System and method for thermal management of high voltage lines in an aircraft

Abstract

The present disclosure relates generally to controlling an aircraft to avoid overheating of components. In one embodiment, a method is disclosed that includes receiving first sensor data indicative of at least one attribute of a high-voltage line of an aircraft, receiving an operating state of the aircraft, determining, based on the first sensor data and the operating state of the aircraft, a proximity of a temperature of the high-voltage line to a temperature limit, and controlling, based on the determined proximity, at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit.

Description

Technical Field

[0001] This disclosure generally relates to the field of powered aircraft. More specifically, but not limited thereto, this disclosure relates to innovations in aircraft using electric propulsion systems. Certain aspects of this disclosure generally relate to monitoring the temperature of high-voltage power lines and controlling the aircraft to prevent the high-voltage lines from overheating. Background Technology

[0002] This disclosure generally relates to preventing overheating of high-voltage lines in aircraft. The temperature of high-voltage lines increases with increasing current, and if high current is maintained, it may exceed safe operating temperatures, posing a risk of wire melting and / or fire. Such events can cause interference, if not harmful, to various systems, including the aircraft's electrical system. Summary of the Invention

[0003] This disclosure generally relates to monitoring the temperature of high-voltage lines and controlling aircraft to prevent overheating of the high-voltage lines. The inventors have recognized several problems that may be associated with supplying power to aircraft, including those using electric or hybrid electric propulsion units (hereinafter referred to as "electric propulsion units" or "EPUs"). For example, the current in the high-voltage lines can vary significantly depending on the aircraft's power requirements, which can vary based on the aircraft's operating state (e.g., hovering, takeoff, winged flight, emergency backup power). The temperature of the high-voltage lines increases with increasing current, and if high currents are maintained, the safe operating temperature may be exceeded, posing a risk of wire melting and / or fire. Furthermore, increasing the diameter of the high-voltage lines to avoid overheating under any flight conditions would increase the weight of the aircraft, thereby reducing range and cargo capacity. Therefore, there is a need for a technique to prevent overheating of high-voltage lines without significantly altering their dimensions.

[0004] One aspect of this disclosure relates to a method comprising: receiving first sensor data indicating at least one attribute of a high-voltage line of an aircraft; receiving an operating state of the aircraft; determining, based on the first sensor data and the operating state of the aircraft, the proximity of the temperature of the high-voltage line to a temperature limit; and controlling at least one electric propulsion unit of the aircraft based on the determined proximity to avoid exceeding the temperature limit.

[0005] Another aspect of this disclosure relates to a system comprising at least one processor configured to execute instructions to cause the system to: receive the operating state of an aircraft; determine, based on first sensor data and the operating state of the aircraft, the proximity of the temperature of a high-voltage line to a temperature limit; and control at least one electric propulsion unit of the aircraft based on the determined proximity to avoid exceeding the temperature limit.

[0006] Another aspect of this disclosure relates to a computer-readable storage medium storing instructions that, when executed by at least one processor, cause at least one processor to: receive the operating state of an aircraft; determine, based on first sensor data and the operating state of the aircraft, the proximity of the temperature of a high-voltage line to a temperature limit; and control at least one electric propulsion unit of the aircraft based on the determined proximity to avoid exceeding the temperature limit.

[0007] Another aspect of this disclosure relates to an aircraft comprising: a first sensor; at least one electric propulsion unit; a high-voltage line; and at least one processor configured to execute instructions to: receive sensor data from the first sensor, the sensor data indicating at least one attribute of the high-voltage line; receive the operating state of the aircraft; determine, based on the first sensor data and the operating state of the aircraft, the proximity of the temperature of the high-voltage line to a temperature limit; and control at least one electric propulsion unit of the aircraft based on the determined proximity to avoid exceeding the temperature limit. Attached Figure Description

[0008] Figure 1 An example electric aircraft consistent with the disclosed embodiments is shown.

[0009] Figure 2A , 2B Images 2C, 2D, 2E, and 2F show exemplary top views of an aircraft consistent with the disclosed embodiments.

[0010] Figure 3 An aircraft control diagram consistent with the disclosed embodiments is shown.

[0011] Figure 4A , 4B 4C, 4D, and 4E demonstrate different models consistent with the disclosed embodiments for controlling the aircraft to avoid overheating of high-voltage lines.

[0012] Figure 5A , 5B 5C and 5D demonstrate how the limits for avoiding overheating, consistent with the disclosed embodiments, vary based on the flight phase of the aircraft.

[0013] Figure 6 A display screen indicating the status of aircraft components, consistent with the disclosed embodiments, is shown. Detailed Implementation

[0014] This disclosure proposes a system to prevent overheating of high-voltage lines in electric or hybrid electric aircraft. The aircraft can be a manned aircraft, an unmanned aircraft (e.g., a UAV), a drone, a helicopter, and / or an airplane. The aircraft includes a fuselage and one or more components configured to allow flight (e.g., wings, tail, propeller). In some embodiments, the aircraft is driven by one or more electric propulsion units (hereinafter referred to as "electric propulsion units" or "EPUs"), which may include at least one engine, at least one rotor, at least one propeller, or any combination thereof. The EPU may be powered by electrical energy (e.g., a battery pack). High-voltage lines can extend from the aircraft's power source (e.g., the battery pack) to the electric propulsion unit. High-voltage lines are crucial for ensuring that the thrust provided by the aircraft's electric propulsion unit maintains the aircraft's stability and controllability.

[0015] To ensure the integrity of the aircraft's high-voltage wiring, one solution is to increase the size of the wiring to ensure it can maintain any expected current rate indefinitely without overheating. However, excessively large wiring increases the aircraft's weight, impacting performance and increasing energy consumption. Another solution is to monitor the wiring's temperature and directly disconnect overheating circuits. However, this solution may not be ideal if the circuitry supplies power to critical aircraft components such as electric propulsion units, other battery packs, etc.

[0016] The disclosed embodiments address these and other problems by monitoring high-voltage line temperature, adjusting the state of one or more battery packs based on the high-voltage line temperature associated with (e.g., powered by) the corresponding battery pack, and / or controlling the aircraft's electric propulsion unit based on the adjusted battery pack state. For example, the state of power of the battery pack can be drated based on an increase in the high-voltage line temperature associated with the battery pack, and the aircraft can be controlled to reduce the use of the battery-powered electric propulsion unit. Therefore, the current through the high-voltage line will be reduced to prevent the line from overheating.

[0017] Furthermore, the disclosed embodiments can predict the time remaining before the high-voltage line reaches its temperature limit and adjust the battery state based on the predicted time. The predicted time before reaching the temperature limit can take into account one or more current or upcoming flight phases of the aircraft. Therefore, even when current consumption varies based on flight phases, the electric propulsion unit can be controlled to prevent the high-voltage line from overheating.

[0018] Figure 1 An example electric aircraft consistent with the disclosed embodiments is shown. Figure 1As shown, in some embodiments, the distributed electric propulsion system of the eVTOL aircraft 100 may include 12 electric propulsion units (e.g., electric motors, fans, turbines, etc.) that may be mounted on booms aft and forward of the main wing of the aircraft 100. The forward electric propulsion unit may tilt (e.g., during flight) between a horizontally oriented position (e.g., generating forward thrust) and a vertically oriented position (e.g., generating vertical thrust). The forward electric propulsion unit may be clockwise or counterclockwise in terms of propeller rotation direction. The rear electric propulsion unit may be fixed in a vertically oriented position (e.g., generating vertical thrust) and may also be clockwise or counterclockwise in terms of propeller rotation direction. As used herein, the term "eVTOL" or "electric aircraft" may refer to an aircraft using at least one electric energy source and may include all-electric aircraft as well as hybrid aircraft using electricity and other fuel sources.

[0019] The aircraft 100 can have various combinations of front and rear electric propulsion units. For example, in some embodiments, the aircraft 100 can have six front electric propulsion units 110 and six rear electric propulsion units 110. In some other embodiments, the aircraft 100 can include four front electric propulsion units and four rear electric propulsion units, or any other combination of front and rear electric propulsion units 110. In some other embodiments, the number of front and rear electric propulsion units is not the same. Figure 2 details different example configurations of the electric propulsion units 110.

[0020] In some embodiments, for vertical takeoff and landing (VTOL) missions, the front electric propulsion unit and the rear electric propulsion unit can provide vertical thrust during takeoff and landing. During the forward flight phase of the aircraft 100, the front electric propulsion unit can provide forward thrust (e.g., in the horizontal direction), while the propeller of the rear electric propulsion unit can be retracted at a fixed position (e.g., its propeller blades are oriented parallel or nearly parallel to the front and rear axes of the aircraft) to minimize drag. The rear electric propulsion unit can be actively retracted via position monitoring.

[0021] In some embodiments, during conventional takeoff and landing (CTOL) missions, the forward electric propulsion unit can provide forward thrust for wing takeoff, cruise, and landing. In some embodiments, the aft electric propulsion unit may not be used to generate thrust during CTOL missions, and the aft propeller may be retracted into place.

[0022] The transition from vertical to forward flight can be achieved via a tilting propeller subsystem, and vice versa. The tilting propeller subsystem can redirect the thrust in the primary vertical direction during the vertical flight phase to the primary horizontal direction during the forward flight phase. The pitch mechanism can change the total angle of the propeller hub assembly blades of the front electric motor for operation during the hovering, transition, and cruise phases.

[0023] like Figure 1 As shown, the high-voltage electrical system (HVPS) of the aircraft 100 may include a power source, such as a battery pack 120, which provides high-voltage electricity that, in other possible uses, is converted into mechanical shaft power to rotate the propeller of the electric propulsion unit 110. The thrust generated by each electric propulsion unit 110 can be controlled by torque commands issued to each electric propulsion unit 110 via a flight control system (FCS) through a digital communication interface.

[0024] In some embodiments, the aircraft includes six battery packs 120, which may be mounted in battery bays within the wings of the aircraft 100. In some embodiments, the six battery packs 120 may have the same design to simplify design, manufacturing, and logistics. The battery packs 120 may power one or more electric propulsion units 110. Although six battery packs 120 are shown, the aircraft 100 may have any number of battery packs 120.

[0025] In some embodiments, a single battery pack 120 may be electrically connected to and power multiple electric propulsion units. For example, in some embodiments, the battery pack 120 may power electric propulsion units 110 located on either side of a longitudinal axis running through the center of the aircraft. In some embodiments, the battery pack 120 may power electric propulsion units 110 located on either side of a horizontal axis running through the wing. In some embodiments, such as Figure 1 As shown, battery pack 120 can power two diagonally opposite electric propulsion units. Therefore, when battery pack 120 fails, the impact on roll or pitch moments can be reduced because lift losses are balanced. In some embodiments, battery pack 120 can power electric propulsion units in different arrangements to reduce roll, pitch, or yaw moments that may be caused by a failure of battery pack 120. For example, in some embodiments, battery pack 120 can be connected to the electric propulsion units in any manner to balance lift and / or forward thrust along the longitudinal and lateral axes of the aircraft.

[0026] Furthermore, the exemplary HVPS system includes a crosslinker 130 with at least one fuse, which allows two or more battery packs 120 to be paired. Through this crosslinker, power from the electric propulsion unit can be shared between the paired battery packs 120. While in Figure 1In the example, battery pack 120 is connected in pairs using three separate crosslinkers 130. However, it should be understood that crosslinkers (e.g., in pairs, triplet, quadruple, or combinations thereof) can be used to connect any number of battery packs and / or some or even all of the batteries can be connected without any crosslinkers. Therefore, multiple battery packs 120 can simultaneously power multiple electric propulsion units. This arrangement provides redundancy and avoids single points of failure, as each paired battery can act as a backup for the others. When a battery pack 120 fails, one or more connected battery packs 120 can continue to power the electric propulsion unit to which the failed battery pack is connected.

[0027] The above-described aircraft and battery pack configuration is provided as an example, but the aircraft may include electric propulsion units with different configurations (e.g., as referenced below). Figures 2A-2F (As shown), battery packs, battery pack connections, and battery pack crosslinking combinations. In some embodiments, each battery pack can power a separate electric propulsion unit (e.g., an electric motor). For example, the aircraft can have four, six, eight, ten, twelve, or any number of electric propulsion units, and the number of battery packs can match the number of electric propulsion units. In some embodiments, each battery pack can power only one electric propulsion unit and can be electrically isolated from all other battery packs. In some embodiments, each battery pack can power one or more partial motors, and each electric propulsion unit (e.g., an electric motor) can include two or more partial motors. Partial motors can include motors capable of independently driving components that can also be independently driven by at least one other motor. Partial motors can be driven by separate motor controllers. Thus, each electric propulsion unit can have a backup power source, but the battery packs can still be electrically isolated.

[0028] In some embodiments, each battery pack can power multiple electric propulsion units. As described above, the battery pack can power multiple sets of electric propulsion units symmetrical about one or more axes of symmetry. In some embodiments, the battery pack can power electric propulsion units (e.g., electric motors) symmetrical about the longitudinal axis, transverse axis, or both of the aircraft. For example, as described above, in some embodiments, different battery packs can power diagonally symmetrical electric propulsion units (e.g., electric motors).

[0029] In some embodiments, the battery pack can power two or more electric propulsion units. In some embodiments, the battery pack 120 can power two or more sets of diagonally symmetrical electric propulsion units. In some embodiments, the battery-powered electric propulsion unit sets may include a pair of inner diagonally symmetrical electric propulsion units and a pair of outer diagonally symmetrical electric propulsion units (e.g., electric motors). In some embodiments, the battery pack can power four or more electric propulsion units arranged symmetrically along a diagonal longitudinal axis.

[0030] In some embodiments, some or all of the battery packs are interconnected. As described above, the crosslinking allows each battery pack to act as backup power for another battery pack. For example, in some embodiments, a first battery pack can directly power a first number of electric propulsion units (e.g., electric motors), while a second battery pack can directly power a second number of electric propulsion units (e.g., electric motors). The first and second battery packs can be crosslinked together to form a battery pack unit. Thus, each battery pack in this unit can act as backup power for another battery pack. When a battery pack in the unit fails, the failed battery pack may disconnect, and the electric propulsion units (e.g., electric motors) will be powered by one or more of the unfailed battery packs in the unit. The battery packs in the battery pack unit can be electrically isolated from other battery pack units.

[0031] In some embodiments, a battery pack unit may include three battery packs, each powering multiple electric propulsion units (e.g., electric motors). For example, in some embodiments, each battery pack may power two diagonally symmetrical electric motors. Therefore, each battery pack unit may power a total of six electric motors, with each electric motor having two backup battery packs. In some embodiments, each battery pack unit may power four electric motors, comprising two sets of diagonally symmetrical electric motors. Therefore, each battery pack unit may power a total of twelve electric motors, with each electric motor having two backup battery packs.

[0032] In some embodiments, the battery pack unit may include four battery packs, each of which powers multiple electric propulsion units (e.g., electric motors). For example, in some embodiments, each battery pack may power two diagonally symmetrical electric propulsion units. Thus, each battery pack unit may power a total of eight electric propulsion units, and each electric propulsion unit has three spare battery packs. In other embodiments, each battery pack unit may power four electric propulsion units, comprising two sets of diagonally symmetrical electric propulsion units. Thus, each battery pack unit may power sixteen electric propulsion units, and each electric propulsion unit has three spare battery packs.

[0033] In some embodiments, all battery packs are connected to a common bus. In some embodiments, the common bus may form a ring power supply, thereby providing additional connection redundancy; while in other embodiments, the common bus may not form a ring power supply.

[0034] In some embodiments, the electric propulsion unit includes a single motor powered by one or more battery packs. In some embodiments, each electric propulsion unit may include two or more partial motors, and the battery packs may power the partial motors. In some embodiments, any of the above-described electric propulsion unit power supply configurations may include powering the partial motors of the battery packs.

[0035] Different configurations of battery packs, electric propulsion units (such as electric motors), the connection between battery packs and electric propulsion units, and battery pack crosslinking components can be selected to optimally balance the aircraft's power requirements, system redundancy, and fault tolerance.

[0036] The high-voltage lines associated with each battery pack can be configured based on variations in the battery pack, electric propulsion unit (e.g., electric motor), the connection between the battery pack and the electric propulsion unit, and the combination of battery pack crosslinking elements. In some embodiments, the temperature of the high-voltage lines can be associated with any connected battery pack that feeds the high-voltage lines (e.g., any battery pack supplying current to the HV lines). Therefore, the battery pack states of the battery packs feeding the high-voltage lines can be appropriately adjusted based on the high-voltage line temperature.

[0037] Furthermore, the electric propulsion units controlled according to battery state can be based on configuration changes. In some embodiments, each electric propulsion unit powered by the battery pack (e.g., each EPU that draws power from the battery pack) can be controlled according to the state of the battery pack. For example, if the battery pack state deteriorates, the aircraft can be controlled to reduce the thrust provided by all electric propulsion units powered by that battery pack.

[0038] Figure 2A , 2B Images 2C, 2D, 2E, and 2F show exemplary top views of an aircraft consistent with the disclosed embodiments.

[0039] Figures 2A-2F This is a top view of an exemplary aircraft consistent with the disclosed embodiments. Numerous design considerations (cost, weight, size, performance, etc.) may exist that influence the number and / or combination of tilting and lifting propellers in the aircraft. The number and orientation of the propellers may affect the number of battery packs and the connections between them (e.g., for controllability and / or stability in the event of a power failure).

[0040] Figure 2A An arrangement of the electric propulsion unit consistent with the disclosed embodiment is shown. (Reference) Figure 2AThe aircraft shown in the figure may be a top view of an exemplary aircraft. The aircraft may include twelve electric propulsion units distributed across the aircraft. In some embodiments, the distribution of the electric propulsion units may include six forward electric propulsion units (265, 266, 267, 268, 269, and 270) and six aft electric propulsion units (271, 272, 273, 274, 275, and 276). In some embodiments, the six forward electric propulsion units may be operatively connected to a tilting propeller, and the six aft electric propulsion units may be operatively connected to a lift propeller. In other embodiments, the six forward electric propulsion units and multiple aft electric propulsion units may be operatively connected to a tilting propeller, and the remaining aft electric propulsion units may be operatively connected to a lift propeller. In other embodiments, all forward and aft electric propulsion units may be operatively coupled to a tilting propeller.

[0041] Figure 2B An alternative arrangement of the electric propulsion unit, consistent with the disclosed embodiment, is shown. (Reference) Figure 2B The aircraft shown in the figure may be a top view of an exemplary aircraft. The aircraft may include eight electric propulsion units distributed across the aircraft. In some embodiments, the distribution of the electric propulsion units may include four front electric propulsion units (277, 278, 279, and 280) and four rear electric propulsion units (281, 282, 283, and 284). In some embodiments, the four front electric propulsion units may be operatively connected to a tilting propeller, and the four rear electric propulsion units may be operatively connected to a lift propeller. In other embodiments, the four front electric propulsion units and multiple rear electric propulsion units may be operatively connected to a tilting propeller, and the remaining rear electric propulsion units may be operatively connected to a lift propeller. In other embodiments, all front and rear electric propulsion units may be operatively coupled to a tilting propeller.

[0042] Figure 2C An alternative arrangement of the electric propulsion unit consistent with embodiments of this disclosure is shown. (Reference) Figure 2C The aircraft may be a top view of an exemplary aircraft. In some embodiments, the aircraft may include ducted fans operatively connected to the electric propulsion unit. In some embodiments, the aircraft may include a set of ducted fans (261, 262, 263, 264) on each wing of the aircraft, and the set of ducted fans may be connected to tilt together (e.g., between lift and forward thrust configurations). In some embodiments, the aircraft includes left and right canards and left and right afts. In some embodiments, each wing of the aircraft includes a set of connected ducted fans. In some embodiments, each set of connected ducted fans may tilt (e.g., between lift and forward thrust configurations), while in other embodiments, only the fan sets on the canards may tilt.

[0043] Figure 2D An alternative arrangement of the electric propulsion unit, consistent with the disclosed embodiment, is shown. (Reference) Figure 2D The aircraft shown in the figure may be a top view of an exemplary aircraft. The aircraft may include six electric propulsion units distributed across the aircraft. In some embodiments, the distribution of the electric propulsion units may include a first group of four electric propulsion units 285, 286, 287, and 288, which are coplanar in a first plane; and a second group of two electric propulsion units 289 and 290, which are coplanar in a second plane. In some embodiments, the first group of electric propulsion units 285, 286, 287, and 288 may be operatively connected to a tilting propeller, and the second group of electric propulsion units 289 and 290 may be operatively connected to a lift propeller. In other embodiments, the first group of electric propulsion units 285, 286, 287, and 288, as well as the rearward second group of electric propulsion units 289 and 290, may all be operatively connected to the tilting propeller.

[0044] Figure 2E An alternative arrangement of the electric propulsion unit, consistent with the disclosed embodiment, is shown. (Reference) Figure 2E The aircraft shown in the figure may be a top view of an exemplary aircraft. The aircraft may include four electric propulsion units distributed across the aircraft. In some embodiments, the distribution of the electric propulsion units may include four coplanar electric propulsion units 291, 292, 293, and 294. In some embodiments, all electric propulsion units may be operatively connected to a tilting propeller.

[0045] Figure 2F An alternative arrangement of the electric propulsion unit, consistent with the disclosed embodiment, is shown. (Reference) Figure 2F The aircraft shown in the figure may be a top view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may include six electric propulsion units distributed across the aircraft. For example, in some embodiments, the aircraft may include four forward electric propulsion units 295, 296, 297, and 298 operatively connected to a tilting propeller, and two aft electric propulsion units 299 and 300 operatively connected to a lift propeller. In some embodiments, the aircraft may include ten electric propulsion units distributed across the aircraft. For example, in some embodiments, the aircraft may include six forward electric propulsion units operatively connected to a tilting propeller, and four aft electric propulsion units operatively connected to a lift propeller. In some embodiments, some or all of the aft electric propulsion units may be operatively connected to the tilting propeller.

[0046] like Figure 2FAs shown, in some embodiments, the aircraft may have a flying wing configuration, such as a tailless fixed-wing aircraft without a fixed fuselage. In some embodiments, the aircraft may adopt a flying wing configuration with the fuselage integrated into the wing. In some embodiments, when the tilt propeller operates in a lift configuration, the tilt propeller can rotate in a plane above the aircraft fuselage.

[0047] Figure 3 An aircraft control diagram consistent with the disclosed embodiments is shown. The Flight Control System (FCS) 320 and Battery Management Unit (BMU) 310 may be implemented by at least one processor (e.g., at least one microprocessor-based controller) configured to execute software code stored in a storage medium (e.g., a computer-readable medium, a non-transitory computer-readable medium) to implement any combination of the functions described herein. The FCS 320 and BMU 310 may also be implemented in hardware or a combination of hardware and software, and may be configured to repeatedly execute individual steps or sequences until a desired or commanded result is achieved.

[0048] Pilot input 303 represents input from a pilot input device indicating a commanded aircraft state. The pilot input device is configured to receive pilot input and generate or influence signals. The pilot input device may include buttons, switches, knobs, joysticks, sliders, controllers, any combination thereof, or any other device configured to generate or influence signals based on the pilot's physical actions. For example, the pilot input device may include one or more of a right controller (e.g., moving the right controller left / right and / or moving the right controller forward / backward), a left controller (e.g., moving the left controller left / right and / or moving the left controller forward / backward), and / or controller switches. In some embodiments, the pilot input device may include an interface to an autopilot system (e.g., a display screen, switch, button, joystick, and / or other interface). Pilot input 303 may also represent input directly from the autopilot system, such as roll commands, climb commands, and / or other commands for controlling the aircraft.

[0049] Control command 304 represents one or more executed processes for providing commands to control the aircraft (e.g., a torque command issued to control assignment 305). Such processes may include modeling the commanded aircraft response (e.g., determining the shape of the ideal aircraft response), feedback and feedforward processes (e.g., determining the appropriate forces to achieve the desired changes in the aircraft), outer-loop processes (e.g., determining the appropriate yaw, roll, command, pitch, and / or thrust), and / or one or more inner-loop control law processes (e.g., determining the torque command). Each of these processes may involve one or more (e.g., stored in memory) control laws, rules, functions, models, and / or algorithms, and may dynamically adjust its output based on inputs (e.g., from pilot input 303, aircraft sensors 307, BMU 310, and / or thermal limit 302).

[0050] Control assignment 305 can control flight elements of the aircraft based on inputs from control command 304, thermal limit 302, BMU 310, and / or aircraft sensors 307. For example, control assignment 305 can control (e.g., send one or more commands to) one or more electric propulsion units (e.g., electric motors, propellers, actuators, rotors, etc.) of the aircraft. Control assignment 305 can further control one or more control surfaces of the aircraft, including flaps, ailerons, rudders, spoilers, and / or elevators.

[0051] Control assignment 305 may include one or more models or functions (e.g., optimization functions) for controlling commands to the electric propulsion unit and / or control surfaces to satisfy a commanded aircraft state while satisfying one or more hard and / or soft constraints. For example, control assignment 305 may weight different priorities (e.g., for aircraft stability or controllability), such as maintaining lift and / or forward thrust, satisfying battery requirements (e.g., dated battery state), satisfying engine thermal requirements (e.g., established by thermal limit 302), and / or preventing propeller speeds corresponding to vibrations from exceeding thresholds. In some embodiments, control assignment 305 may determine multiple control solutions to satisfy the commanded aircraft state and (e.g., based on the relative weights between priorities) select the solution that best satisfies the priority. Aircraft dynamics 306 represents the controlled flight elements (e.g., the electric propulsion system and / or control surfaces) and aircraft dynamics (e.g., how the aircraft responds to flight element controls (aircraft orientation, motion, etc.)).

[0052] Aircraft sensor 307 can detect aircraft dynamics, aircraft state, and / or the flight phase of the aircraft. Aircraft sensor 307 may include one or more sensors configured to detect aircraft dynamics and / or aircraft state, such as acceleration and / or pitch orientation sensors (e.g., accelerometers, triaxial accelerometers, gyroscopes, and / or triaxial gyroscopes), airspeed sensors (e.g., pitot tube sensors), and / or ground speed sensors. Aircraft sensor 307 may further include one or more inertial measurement units (IMUs) for determining the aircraft state. The aircraft state may include at least one of the following: aircraft position (e.g., yaw, roll, pitch, and / or any other orientation along one or two axes), aircraft velocity, aircraft angular rate (e.g., roll, pitch, and / or yaw), and / or aircraft acceleration (e.g., longitudinal, lateral, and / or vertical acceleration), or any physical characteristic of the aircraft or its components.

[0053] Aircraft sensor 307 may include one or more sensors for detecting the flight phase of the aircraft. For example, aircraft sensor 307 may include an atmospheric data system and / or an airspeed sensor (e.g., a pitot tube), a ground speed sensor, a GPS sensor, a propeller speed sensor (e.g., a Hall effect sensor and / or an optical sensor), a propeller tilt sensor (e.g., a magnetic sensor, a position displacement sensor, a linear displacement sensor, etc., for measuring nacelle tilt angle), an accelerometer, and / or a gyroscope to determine the flight phase. In some embodiments, the flight phase may be determined based on comparing one or more of these sensor measurements with pre-stored thresholds. For example, aircraft sensor 307 may determine the flight phase based on airspeed measurements (and / or ground speed measurements) and / or propeller tilt angle measurements. For example, a higher airspeed and tilt angle in a forward thrust configuration may correspond to a forward flight phase, and a lower airspeed and tilt angle in a lift configuration may correspond to a hovering and / or takeoff flight phase. A "phase of flight" or "flight phase" (e.g., hovering, cruise, forward flight, takeoff, landing, transition) can be defined by a combination of flight conditions (e.g., a specific range of combinations of flight conditions), which may include one or more of the following: airspeed, altitude, pitch angle (e.g., the pitch angle of the aircraft), roll angle (e.g., the roll angle of one or more propellers), roll angle, rotational speed (e.g., the rotational speed of the propellers), torque value, pilot command, or any other value indicating at least part of the current or requested (e.g., commanded) state of the aircraft. A flight phase may include one or more of the following: vertical takeoff, short takeoff, hovering, forward flight, wing-borne flight, vertical descent, conventional landing, and / or any other flight motion that affects the current consumption of the electric propulsion unit.

[0054] High-voltage sensing, specifically HV sensor 301, can detect the state of one or more components of the aircraft's high-voltage electrical system. For example, HV sensor 301 may include one or more sensors associated with battery pack 120 for collecting sensor data indicating at least one attribute of the aircraft's high-voltage lines (e.g., HV line temperature, wing ambient temperature, current, voltage, power, and / or frequency). For example, HV sensor 301 may include one or more current sensors, voltage sensors, and / or temperature sensors, which may be associated with one or more battery packs. In some embodiments, the temperature sensor may be located on portions of the positive and / or negative high-voltage channels that feed (are configured to power) each electric propulsion unit 110, and / or on portions of the high-voltage lines that feed (are configured to power) another battery pack (e.g., crosslink 130). In some embodiments, the temperature sensor may be mounted within the aircraft wing to monitor the ambient air temperature (wing ambient temperature) near one or more high-voltage channels. In some embodiments, current sensors may be placed on portions of the positive and / or negative high-voltage channels that feed (are configured to power) at least one (e.g., each) electric propulsion unit, and / or on portions of the high-voltage lines that feed (are configured to power) another battery pack (e.g., crosslink 130). In some embodiments, one or more sensors may detect when the battery pack is disconnected. For example, one or more sensors may detect whether the battery pack is disconnected based on a fuse blowing, current measurement (e.g., a current measurement of zero or falling below a threshold), and / or voltage measurement (e.g., a voltage measurement of zero or falling below a threshold).

[0055] Combined with the following text Figures 4A-4C Further described, thermal limit 302 includes one or more processes for determining proximity to a thermal limit (e.g., a temperature limit) imposed by the high-voltage line temperature. As described below, in some embodiments, the processes performed by thermal limit 302 may be performed by FCS 320; while in other embodiments, the processes performed by thermal limit 302 may be performed by BMU 310. In some embodiments, one or more steps may be performed by BMU 310, while other steps may be performed by FCS 320. In some embodiments, the processes performed by thermal limit 302 may be performed by one or more processors independent of BMU 310 and FCS 320.

[0056] As detailed below, the limits established by high-voltage line temperature can be considered alone or in conjunction with other battery limits. For example, the limits imposed by high-voltage line temperature can be applied by derating the battery pack state considered by control command 304 and / or control assignment 305. Dereasing the battery pack state can include adjusting descriptive values ​​or weights associated with the battery pack, such as by adjusting the state of power, state of charge, state of health, state of energy, available current, and / or any other capacity, state, or processing metrics of the battery pack (e.g., weights within an algorithm). As mentioned above, the battery pack state (discussed further below) can include the battery pack's state of power, state of health, state of energy, and / or any other battery operating information. The battery pack state can be determined based on measurements from one or more sensors (e.g., from aircraft sensor 307 and / or HV sensor 301). For example, the battery pack state can be determined based on monitoring current, voltage, temperature, and / or tracking the battery pack's lifetime. In some embodiments, the state of charge can indicate the battery pack's ability to store (or provide) charge at a given moment. State of charge (SOC) can be expressed as an absolute value (e.g., coulombs or ampere-hours, Ah) or as a ratio or percentage relative to the maximum capacity of the battery pack to store (or provide) charge. In some embodiments, SOC can refer to the available battery pack capacity relative to the battery pack's rated capacity. For example, SOC can be based on the battery pack's open-circuit voltage (OCV), where OCV is the resting voltage of the battery pack (e.g., based on a battery pack with no current flowing for a set period of time). When the battery pack is not resting for a set period of time, the SOC of the battery pack can be estimated (e.g., determined in the resting state) using coulomb counts and initial SOC. For example, SOC = (initial SOC + integral (current)) dt)) / battery capacitance. In other embodiments, other methods may be used to determine the SOC of the battery pack. For example, Kalman filtering may be used to determine the SOC, which uses a model that correlates the SOC with battery dynamics, current and / or voltage measurements. In some embodiments, the state of energy may refer to the amount of energy available in the battery pack. For example, the state of energy may be determined by taking into account the health condition of the battery pack (e.g., lifetime, degradation, etc.) to determine the amount of energy available in the battery pack. In some embodiments, the state of power may indicate the available power that can be provided by the battery pack within a certain time range (e.g., not exceeding at least one system constraint (such as battery pack voltage constraint, battery cell temperature constraint, HV line current constraint, etc.)). The state of power may be expressed as an absolute value (e.g., kW, W) or as a ratio or percentage relative to the maximum rated power (e.g., maximum rated system power). The following will combine Figures 4A-4EThe method for determining the state of power is further described in detail. In some embodiments, the state of health may refer to the current condition or overall health of the battery relative to its original, acceptable, preferred, or optimal state. For example, the state of health may reflect the magnitude of impedance that adversely affects battery pack performance and / or the charge that can be drawn from the battery. In some embodiments, the state of health may indicate the battery pack's ability to store (or provide) charge at a given time relative to its maximum original charge storage (or provision) capacity. The state of health may be expressed as a ratio or percentage relative to the battery pack's maximum charge storage (or provision) capacity.

[0057] In other embodiments, the limit established by the wire temperature can be applied separately from other battery limits. For example, proximity to a temperature limit can directly result in adjustments to the power allocated to one or more electric propulsion units. In some embodiments, based on proximity to a temperature limit, in addition to controlling the power consumption of the electric propulsion units, the BMU 310 and / or FCS 320 can also (e.g., via a pilot interface associated with the FCS 320) provide warnings to the pilot and / or command certain flight maneuvers (e.g., emergency landing). Proximity to a temperature limit can be, can include, or can be based on the time it takes to reach the temperature limit.

[0058] Figure 4A A model 400 for controlling an aircraft to avoid overheating of high-voltage lines, consistent with the disclosed embodiments, is shown. In some embodiments, model 400 may be implemented as a process, a computer-implemented method, a computerized model, and / or a computerized algorithm or program. For ease of description, the various parts of model 400 will be described in the form of steps, but it should be understood that they may also be implemented as modules, computerized functions, programs, calculations, or any combination thereof. It is worth noting that effective implementation by a human user is required. Figure 4A , 4E The steps of the exemplary model depicted in 4D are impossible (or at least impractical), especially considering that these functions are frequently (e.g., continuously, constantly) implemented during aircraft flight (including takeoff or landing), and / or dynamically implemented based on (e.g., in response to) received signals (e.g., aircraft sensors that can continuously measure various conditions, and / or pilot input devices). Generally, it can be understood that... Figure 4A , 4E Any / all steps of the exemplary 4D model may be executed by at least one processor (e.g., a processor associated with the FCS and / or BMU), such as by executing one or more instructions stored on a computer-readable medium (e.g., a non-transitory computer-readable medium).

[0059] In step 401, the ambient air temperature inside the aircraft wing can be determined (e.g., received, requested, measured, and / or calculated). For example, the ambient air temperature can be received from one or more nearby temperature sensors (e.g., within one foot, several feet, several meters, or within the same aircraft wing) of the high-voltage conductor being evaluated. In some embodiments, as further described below in step 402, other measurements can be collected to determine the initial conductor temperature. Current, battery temperature, electric propulsion unit temperature, coolant temperature, and / or coolant flow rate can be measured. For example, a current sensor can collect the initial current through a high-voltage line configured to power the electric propulsion unit. Temperature sensors coupled to the battery pack, electric propulsion unit, and / or coolant piping can determine the temperature of these respective components. Flow rate or pressure sensors coupled to the coolant piping can determine the coolant flow rate.

[0060] In step 402, the conductor core temperature of the high-voltage line (e.g., the HV line to the electric propulsion unit) can be estimated. In some embodiments, a thermal resistance model can be used to determine the initial conductor temperature based on the conductor's resistance and size, the proximity between conductors, the conductor sheath, variations in ambient air temperature, and / or the current flowing through the conductor. For example, a thermal resistance model similar to resistance model 402a can be used, and the initial conductor core temperature can be calculated.

[0061]

[0062] T i n+1 = Conductor temperature at the current time step (e.g., initial core temperature), (T i n+1 – T j n+1 = Temperature variation across nodes (e.g., core, sheath, environment), T i n = Conductor temperature at the previous time step, Δt = time change, C i = thermal mass, R i-j = Thermal resistance of conduction, convection and / or radiation, P i = Ohmic heat generated by the conductor (as a function of current and resistance). In some embodiments, the previous time step T can be approximated (e.g., based on historical information, ambient temperature, etc.). i n The temperature.

[0063] In some embodiments, the initial conductor temperature can be approximated based on the average current consumption and / or average conductor temperature during a typical flight of the aircraft. In some embodiments, experimental data or simulations can be used to develop algorithms, models, or lookup tables to determine the initial conductor temperature based on one or more influencing parameters, such as measured battery temperature, ambient air temperature, engine temperature, coolant temperature, coolant flow rate, time since the last flight, and / or details of the last flight (e.g., power consumption of the electric propulsion unit during flight, flight phases experienced during flight, flight duration, and / or flight distance).

[0064] After solving for the initial conductor temperature, the conductor temperature at subsequent time points can be determined (e.g., estimated). For example, the conductor temperature can be determined using a time-transient model that considers (e.g., based on and / or configured to receive inputs such as constraints) the materials of the conductor and sheath, the spacing between the conductors, the dimensions of the conductor and sheath, and the spacing between components. For example, a time-transient model can be used to estimate the conductor temperature based on (e.g., using determined or estimated values) at least one of the heat generated by the current, the heat dissipated by conduction, the heat dissipated by convection, or the heat dissipated to the surrounding environment by radiation. In some embodiments, the resistance model 402a described above can be used to determine the subsequent temperature of the conductor; while in other embodiments, the following model can be used to determine the subsequent temperature of the conductor:

[0065]

[0066] T n+1 = Conductor temperature at the current time step; T n = Initial conductor temperature and temperature at the previous time step; dt = time step; D c = Conductor diameter; c p,c = Thermal conductivity; I n = Current; r c = conductor radius; ρ c,r = Conductor density; α = Conductor thermal diffusivity; r s = Surface radius; T s = Conductor surface temperature (e.g., outer sheath surface); e = thermal emissivity; T e = Ambient temperature (e.g., ambient wing temperature).

[0067] The conductor temperature can be calculated iteratively at different times (e.g., using discrete time intervals), where T n+1 The calculated value becomes T in the next time step. nIn some embodiments, the computation rate can be set based on experimental data and / or modeling. In some embodiments, the computation rate can be automatically adjusted based on the increase in conductor core temperature. In some embodiments, the conductor temperature can be calculated by integrating the corresponding equation for solving dT / dt at different time steps. In some embodiments, the ambient thermal temperature T can be measured using one or more sensors installed within the aircraft wing. e (For example, changes in the temperature around the wing) can update the above model accordingly.

[0068] In step 403, the conductor core temperature is compared to one or more temperature limits and / or ranges. In some embodiments, the normal load temperature limit may indicate a safe temperature range. For example, a safe temperature range may indicate a temperature range where the risk of damage to the high-voltage line is limited or nonexistent. Based on the comparison results (i.e., one or more comparisons), model 400 may take additional actions. For example, if it is determined that the conductor temperature is within the normal load temperature limit range, no further steps are taken, and the conductor temperature continues to be monitored. Overload temperature may indicate an undesirable temperature range due to its proximity to the conductor's thermal design limits. For example, in some embodiments, the overload temperature limit may indicate a temperature range where adjustments to aircraft controls and / or warnings are provided to avoid reaching design limits. Design limits may indicate temperatures at which the conductor risks damaging itself and / or surrounding components (e.g., melting, fire, etc.). In some embodiments, design limits may be based on limits established by the conductor manufacturer. In some embodiments, when a design limit is reached, a warning may be issued to the flight control interface, and / or the battery state may be derated to zero (or a minimum set value). Additionally or alternatively, temperature limits may include the total current flow over a period of time (e.g., the current integral). Additionally or alternatively, temperature limits may include functions, models, ranges (e.g., the ranges discussed above), and / or curves (such as...) Figure 4C (Exemplary curves depicted), or represented by them.

[0069] In step 404a, the remaining time for the distance to reach the limit can be determined. For example, the time for the distance to reach the design limit and / or overload limit can be determined based on the conductor temperature determined in step 402. In some embodiments, the time for the distance to reach the limit can be determined based on one or more models, functions, and / or lookup tables that take into account (e.g., based on and / or configured to receive inputs such as constraints) the physical properties of the conductor and the conductor temperature. Figure 4BAs shown, the time to reach the distance limit can be determined based on temperature profiles (e.g., reflecting the physical properties of the conductor). For example, the time to reach the distance limit can be determined based on an I²t curve (and / or a curve proportional to the I²t curve) that considers the conductor structure, material, and cross-sectional area. Therefore, different temperature limits and their proximity to the corresponding temperature limits can be determined for different battery packs. Furthermore, as... Figure 4C As shown, in some embodiments, the curve can be based on changes in the ambient temperature of the wing. For example, as illustrated, for lower wing ambient temperatures (e.g., 50 degrees Celsius), the time to reach the limit can be longer, and for higher wing ambient temperatures (e.g., 90 degrees Celsius), the time to reach the limit can be shorter. Curves reflecting conductor properties can be used... Figure 4A and 4D -4E outlines any control process to determine when the distance reaches the limit. For example... Figure 4B As shown, the curves represent limits (e.g., overload limits and / or design limits) and show that when the current drawn through the high-voltage line is less than the first current value (line 1) of the second current value (line 2), the conductor will require a longer time to reach the limit.

[0070] In some embodiments, one or more models, functions, and / or lookup tables can be manipulated (e.g., generating one or more predictions) using upcoming changes in current consumption (e.g., predictive changes, anticipated changes based on flight phases and / or flight plans). Figure 4B As shown, current consumption can vary based on the aircraft's operating state (e.g., flight phase or emergency situation). In the first flight phase (line 3), current consumption may increase, resulting in a shorter time to reach the distance limit; while in the second flight phase (line 4), current consumption may decrease, resulting in a longer time to reach the distance limit. As described above, in some embodiments, the flight phase can be determined based on information from aircraft sensor 307. In some embodiments, the flight phase can be determined by comparing the aircraft's position (e.g., measured by GPS in aircraft sensor 307) with pre-stored navigation information indicating flight phase information for different segments of the journey. This will be discussed in conjunction with the following... Figures 5A-5D As shown, current consumption can vary significantly depending on the flight phase. This variation in current consumption can be based on (e.g., determined by) simulation data and / or experimental data (e.g., using current sensors during flight). In some embodiments, model 400 can determine limits (e.g., in step 403) and / or proximity to limits (e.g., in step 404a) based on one or more upcoming flight phases (e.g., based on analysis of at least one upcoming flight phase or otherwise using at least one upcoming flight phase, such as input variables or constraints).

[0071] In some embodiments, one or more models, functions, and / or lookup tables may provide (e.g., store, determine, predict, and / or calculate) the time to reach a limit based on the conductor temperature determined in step 402, one or more values ​​used to determine the conductor temperature, conductor current consumption, one or more existing flight phases, and / or one or more upcoming flight phases. In some embodiments, analyzing one or more upcoming flight phases may include aggregating the effect of the expected flight duration in each flight phase on the conductor temperature.

[0072] In step 405, the battery state of the aircraft can be determined. For example, the battery pack state can include the state of power, state of charge, state of health, state of energy, and / or any other battery state for each battery pack. At least one battery state can be determined based on measurements from one or more sensors (e.g., from aircraft sensor 307 and / or HV sensor 301). Additionally or alternatively, at least one battery state can be determined by the BMU. In some embodiments, the state of power of the battery pack can be determined. The state of power can indicate the available power that the battery pack can provide over a certain time period (e.g., not exceeding at least one system constraint (such as battery pack voltage constraint, battery cell temperature constraint, HV line current constraint, etc.)). The state of power can be expressed as an absolute value (e.g., kW, W) or as a ratio or percentage relative to the maximum rated power (e.g., maximum rated system power). In some embodiments, multiple states of power over different time periods can be determined, and each state of power can be derated (as described below in conjunction with 404b). In some embodiments, the state of power can be determined using the state of charge (e.g., based on the initial state of charge and current consumption) and the state of health (e.g., based on the overall condition of the battery pack). In some embodiments, the state of power (SOP) at different time intervals can be determined based on the state of charge (SOC), state of health (SOH), and / or additional electrical conditions (e.g., measured and / or expected current or voltage). For example, a lookup table or model can retrieve and / or compute the SOP for one or more time intervals based on SOC, SOH, and / or additional electrical conditions (e.g., using a function or algorithm). In some embodiments, an electrothermal model can be used to determine the SOP, which details how the battery pack can be determined based on changes in current and voltage (e.g., changes in temperature). In some embodiments, a neural network trained on various inputs (e.g., SOC, SOH, voltage, current, etc.) can be used to determine the SOP.

[0073] In step 404b, the battery state can be adjusted (e.g., derating) based on the time it takes for the distance to reach a limit. For example, as the time to reach a limit (e.g., an overload temperature limit or a design limit) decreases, the battery state may degrade. In some embodiments, battery state derating may follow a linear model. In some embodiments, battery state derating may follow a nonlinear model (e.g., a sinusoidal model). In some embodiments, the drated battery state can be sent to the flight control system (e.g., FCS 320). In some embodiments, other information indicating proximity to the limit (e.g., the time it takes for the distance to reach the limit and / or other indications of proximity to the limit) can be sent to the flight control system FCS 320.

[0074] In step 406, the flight control system (e.g., control command 304 and / or control assignment 305) may receive the dated battery status (and / or other indications of proximity to design limits) and control the electric propulsion unit based on the dated battery status. For example, the FCS 320 may store and / or receive indications of which electric propulsion units are powered by which battery packs and control the electric propulsion units based on the status of their associated battery packs (e.g., battery packs supplying power to the EPU and / or providing backup power). For example, the FCS 320 may determine (e.g., relative to the power requirements of components connected to other buses and / or relative to the initial power requirements allowed by the battery packs in a dated state before dating) to reduce the power requirements of components connected to the dated battery packs (e.g., reduce power requirement limits). In some embodiments, the flight control system 320 may determine that the electric propulsion unit is receiving power from a backup battery pack based on information from the battery management system (e.g., an indication from the BMU 321 that a battery pack fuse has blown or is no longer being powered by the battery pack). For example, the flight control system 320 can determine which additional electric propulsion units are being powered by the battery pack in an emergency based on the relationships between the received and / or stored battery pack, electric propulsion units, and interconnects.

[0075] In some embodiments, a derated battery state can be provided as a battery pack state. As detailed above, control allocation 305 can receive commands (e.g., torque commands, thrust commands, etc.) for achieving an aircraft state and determine torque and / or thrust commands for the electric propulsion units to achieve the commanded aircraft state, while also satisfying certain hard constraints (e.g., aircraft stability and / or controllability) and / or soft constraints. In some embodiments, control allocation 305 can determine multiple solutions for achieving the commanded state and can prioritize solutions that better satisfy the constraints. In some embodiments, solutions can be determined based on prioritization schemes for the electric propulsion units (e.g., schemes that prioritize EPUs based on their position relative to the aircraft wings and / or fuselage and / or their relative control rights). For example, dependency weights can be assigned to different electric propulsion units (EPUs), and these weights can be adjusted based on the derated battery state associated with one or more EPUs. For example, the dependency weights of the electric propulsion units can be adjusted based on the battery pack's state of power, and these weights can be reduced based on the reduced state of power due to devaluation (e.g., proportionally). Therefore, as the HV line associated with the first EPU approaches the temperature limit, the reliance on the first electric propulsion unit may decrease compared to the second electric propulsion unit. For example, commands to the first electric propulsion unit (e.g., electric or torque commands) can be adjusted such that the thrust provided by the first electric propulsion unit decreases relative to the second electric propulsion unit as the HV line associated with the first EPU approaches the temperature limit. As detailed above, the flight control system can manage the power consumption of multiple electric propulsion units (e.g., each electric propulsion unit of the aircraft) based on a determined proximity to prevent each electric propulsion unit from exceeding its temperature limit while maintaining the stability of the aircraft.

[0076] In some embodiments, the flight control system may receive an undegraded battery state in addition to a dated battery state. Therefore, in an emergency, if the risk of overheating of the high-voltage lines is less than the risk of reducing the utilization of the associated electric propulsion units, the flight control system 320 may be configured to rely on (e.g., use) an undegraded battery state instead of a dated battery state (override the dated battery state). The flight control system may store one or more flight conditions for which an undegraded battery state (e.g., power state) should be used. For example, in some embodiments, an undegraded battery state (e.g., power state) may be used for EPU control if the power required to maintain aircraft control or stability exceeds the power provided by the dated battery pack, or if the aircraft is performing an emergency landing. In some embodiments, the flight control system 320 will rely on the undegraded battery state when it determines an emergency. For example, the flight control system 320 may determine an emergency based on one or more electric propulsion unit failures, one or more battery pack failures (e.g., short circuits, overcurrent conditions, battery cell failures, battery depletion), control surface failures, tilt propeller system failures, and / or other failures of critical systems.

[0077] In some embodiments, while the aircraft is in flight, based on at least one of changes in wing ambient temperature, conductor core temperature, HV line current, and / or changes in aircraft operating status (e.g., during flight phases or emergency situations), the determination of proximity to a temperature limit (step 404a) and corresponding battery state derating and control adjustments (steps 404b and 406) are continuously performed. In some embodiments, these steps are performed iteratively at preset time intervals.

[0078] Figure 4D A model for controlling an aircraft to avoid overheating of high-voltage lines, consistent with the disclosed embodiments, is shown. In some embodiments, in addition to or instead of using the above-described combination... Figure 4A The models described in steps 401 and 402 can also use empirical data to correlate wing ambient temperature and current with the proximity of temperature limits. For example, simulation and / or experimental data can be used to determine the relationship between wing ambient temperature, current, and the time it takes for the distance to reach the limit. In some embodiments, this relationship can further include the influence of flight phases. The relationship information can be stored as one or more models, functions, and / or lookup tables that retrieve the time it takes for the distance to reach the design limit based on wing ambient temperature and current measurements. The remaining steps are combined with those described above. Figure 4A It has been described.

[0079] Figure 4EAnother model for controlling an aircraft to avoid overheating of high-voltage lines, consistent with the disclosed embodiments, is shown. As described above, in some embodiments, the flight control system (FCS) 320 may perform one or more steps to determine the proximity to a high-voltage line temperature limit. As shown, in some embodiments, the BMU 310 provides the measured wing ambient temperature, battery status, and high-voltage line current consumption to the flight control system. The flight control system may determine the time close to reaching the limit and derating the battery pack status as described above. In some embodiments, as combined with the above... Figure 4A The time it takes for the distance to reach the limit is based on one or more models used to estimate the conductor temperature. In other embodiments, as described above... Figure 4B The time it takes for the distance to reach the limit is based on empirical data that establishes the relationship between wing ambient temperature, current, and the time it takes for the distance to reach the limit.

[0080] In some embodiments, Figure 4A The process shown and Figure 4D-4E The method shown can be based on temperature measurements of the high-voltage line (e.g., temperature sensors connected to the high-voltage line), and, in addition to or replacing the methods described above for determining the high-voltage line temperature (e.g., based on wing ambient temperature modeling, etc.), the closeness to the temperature limit can be based on this measurement. For example, the temperature measurement can come from a temperature sensor connected to a high-voltage line feeding the motor and / or a line feeding the battery pack (e.g., crosslink 130). In some embodiments, the measured temperature of the high-voltage line can be combined with the methods described above to determine the closeness to the temperature limit. For example, the measured temperature can be combined with the above. Figure 4A The average temperature of the model described in step 402 is taken.

[0081] In some embodiments, consistent with the disclosed embodiments, Figure 4A and 4D The processes shown in -4E can be executed continuously and the control can be adjusted based on determined (e.g., measured, estimated, calculated) values. In some embodiments, these processes can be executed once at set time intervals.

[0082] Figures 5A-5D This demonstrates how current consumption and the corresponding time to reach the temperature limit can vary based on changes in the aircraft's flight phase. Consistent with the disclosed embodiments, the effects of existing and / or upcoming flight phases (e.g., on current consumption, temperature limits, and / or the time to reach the temperature limit) can be included in one or more models, functions, and / or lookup tables for determining proximity to the temperature limit (e.g., the time to reach the temperature limit). Although Figures 5A-5DThe charts in the diagram are shown with specific current, time, and temperature scales, but it should be understood that these scales are merely illustrative and that the values ​​of current and / or temperature may vary over time in different scenarios (though the trends depicted may remain constant or similar).

[0083] Figure 5A An example of current consumption during the takeoff phase of an aircraft (e.g., vertical takeoff) is illustrated. As shown, the current consumption through the high-voltage lines is greater than in other phases (e.g., the forward flight shown below). Consequently, the high-voltage line temperature rises more rapidly. Furthermore, during the takeoff phase, both current consumption and the corresponding high-voltage line temperature increase. Therefore, during takeoff, based on the high starting current and the expected increase in current consumption, it is determined (e.g., as combined above)... Figure 4A , 4D The time it takes for the distance to reach the temperature limit (or to be determined by one or more of the four phases of flight) can be shorter than in other flight phases.

[0084] Figure 5B An example of current consumption during the forward flight phase is shown. As illustrated, current consumption through the high-voltage lines is lower than in other flight phases (e.g., takeoff or hovering). Furthermore, current consumption increases when maintaining a constant power consumption. The corresponding high-voltage temperature of the lines can rise more slowly than in other flight phases (e.g., hovering and / or takeoff). Therefore, when the aircraft transitions to forward flight, the time it takes to reach the temperature limit for a given distance can be longer than in other flight phases (e.g., hovering and / or takeoff).

[0085] Figure 5C An example of current consumption during an aircraft emergency is shown. An emergency could include a situation where the battery pack provides backup power to one or more additional electric propulsion units. For example, Figure 5C The current consumption figure illustrates how the current changes during emergency situations encountered in the hovering phase of flight, such as during a vertical landing. As shown, the current consumption through the high-voltage lines is greater than in other phases (e.g., forward flight and / or takeoff as shown above). Consequently, the high-voltage line temperature rises more rapidly. Furthermore, both the current consumption and the corresponding high-voltage line temperature increase. Therefore, when the aircraft is in an emergency, the time it takes to reach the temperature limit at a given distance can be shorter than in other flight phases (e.g., forward flight and / or takeoff as shown above).

[0086] Figure 5DAn example of aircraft takeoff, followed by an example of current consumption during an emergency, is shown. As described above, in some embodiments, one or more upcoming flight phases and / or aircraft states may be considered when determining the time to reach the temperature limit. For example, determining the time to reach the temperature limit (proximity to the temperature limit) includes considering the impact of at least two upcoming flight phases on the high-voltage line temperature (e.g., predicted temperature increase or predicted temperature decrease). Figure 5D The scenario depicts an aircraft taking off and then encountering an emergency approximately 300 seconds later. For example, an emergency could include the battery pack supplying power to one or more additional electric propulsion units. Once an emergency begins, current consumption and the corresponding high-voltage line temperatures rise. Therefore, if one of the predicted upcoming aircraft states includes an emergency, the time it takes to reach the temperature limit can be determined more quickly.

[0087] Figure 6 A display screen indicating the status of aircraft components, consistent with the disclosed embodiments, is shown. This display screen can be shown on a screen within the aircraft (e.g., in the cockpit). In some embodiments, the FCS 320 can adjust the display screen based on information received from the BMU 310 and / or other components of the aircraft. In some embodiments, the status of one or more aircraft components can be displayed (e.g., on display screen 601 associated with the FCS 320). In some embodiments, the status of an aircraft component can indicate the temperature associated with that aircraft component. For example, battery pack temperature (BT) can indicate the temperature of one or more battery pack components and / or the temperature of associated high-voltage lines. In some embodiments, the highest temperature of one or more battery pack components and high-voltage lines can be displayed (e.g., on the display temperature section 602). Oil temperature (OT) can indicate the temperature of the oil used in the electric propulsion unit of the aircraft. Stator winding temperature (ST) can indicate the temperature of the stator windings of the electric propulsion unit.

[0088] In some embodiments, the proximity of one or more aircraft components to a corresponding thermal limit (e.g., a temperature limit) can be displayed. For example, oil temperature (OT), stator winding temperature (SWT), and / or battery temperature (BT) relative to one or more limits (represented by different shaded lines at the top of the instrument) can be displayed. In some embodiments, the aircraft component closest to its corresponding temperature limit can be selected (e.g., via BMU 310 and / or FCS 320) for display, and its proximity to the temperature limit can be shown. For example, only one of the oil temperature (OT), stator winding temperature (SWT), battery temperature (BT), or other aircraft component temperatures can be selected to be displayed on the display temperature section 602 based on their proximity to the corresponding thermal limit (e.g., a temperature limit).

[0089] In some embodiments, the display screen may show an estimated or predicted future state (e.g., future temperature) of one or more aircraft components. For example, as shown in the extended display screen of the temperature section 602 of the display screen, a trend line 603 may indicate the expected temperature of the component within a set time interval.

[0090] In some embodiments, additional details about the aircraft components may be displayed (e.g., on a separate display panel and / or a separate optional interface). For example, details about which battery pack component and the corresponding battery pack is closest to the thermal limit may be provided. For example, the layout of the aircraft components and sub-components may be displayed visually (e.g., as a schematic diagram), and the component closest to the corresponding thermal limit may be indicated (e.g., through shading or text).

[0091] In some embodiments, the display may indicate one or more control actions being performed automatically (e.g., adjusting the EPU and / or emergency landing) to prevent aircraft components from overheating. In some embodiments, the display may advise the pilot to take one or more control actions to prevent aircraft components from overheating. For example, the display may advise the pilot to land at an alternative destination, perform a specific type of landing (e.g., a conventional landing or hovering), and / or allocate a specific amount of time before the next flight to cool one or more aircraft components. The display may be dynamically and / or continuously updated, such as based on at least one process or model (such as those incorporated herein). Figure 4A , 4D The repeated determinations made (as described in 4E).

[0092] Additional aspects of this disclosure may be further described by the following terms:

[0093] 1. A method for controlling an aircraft, the method comprising:

[0094] Receive first sensor data, which indicates at least one property of the high-voltage lines of the aircraft;

[0095] Receive the operational status of the aircraft;

[0096] The approximation of the high-voltage line's temperature to the temperature limit is determined based on the first sensor data and the aircraft's operating status; and

[0097] The aircraft's at least one electric propulsion unit is controlled based on a determined proximity to avoid exceeding the temperature limit.

[0098] 2. The method according to Clause 1, wherein the at least one attribute includes at least one of the following: the temperature of the high-voltage line, the ambient temperature of the wing, or the current passing through the high-voltage line.

[0099] 3. The method according to Clause 1 or 2, wherein the at least one attribute includes the temperature of the high-voltage line received from a temperature sensor on the high-voltage line, and wherein the high-voltage line is configured to provide power to at least one of: an electric propulsion unit or a battery pack.

[0100] 4. The method according to any one of Clauses 1-3, wherein the at least one property includes the wing ambient temperature and the current through the high-voltage line.

[0101] 5. The method according to any one of Clauses 1-4, wherein the operational state includes at least one of a flight phase or an emergency state.

[0102] 6. The method according to any one of Clauses 1-5, wherein the operating state includes an emergency state indicating whether the battery pack is providing backup power.

[0103] 7. The method according to any one of Clauses 1-6, wherein the operating state includes a flight phase, which includes at least one of the following: hovering, takeoff, or forward flight.

[0104] 8. The method according to any one of Clauses 1-7, wherein the operating state is based on at least one of: airspeed measurement, altitude measurement, ground speed measurement, propeller tilt angle measurement, or GPS measurement.

[0105] 9. The method according to any one of Clauses 1-8,

[0106] This operational status includes at least two upcoming flight phases; and

[0107] Determining the proximity to the temperature limit includes taking into account the impact of the at least two upcoming flight phases on the high-voltage line temperature.

[0108] 10. The method according to any one of clauses 1-9, wherein the proximity to the temperature limit is determined based on at least one of a model or a lookup table.

[0109] 11. The method according to any one of clauses 1-10, wherein the proximity to the temperature limit is based on the proximity to a curve that corresponds to (high-voltage line current)^2. (High-voltage line temperature) is proportional.

[0110] 12. The method according to any one of clauses 1-11, wherein the proximity to the temperature limit is determined based on a model or lookup table, and wherein the inputs to the model or lookup table include: the temperature of the high-voltage line, the current through the high-voltage line, and the flight phase of the aircraft.

[0111] 13. The method according to any one of Clauses 1-12, wherein the proximity to the temperature limit is the time from when the high-voltage line temperature reaches the temperature limit.

[0112] 14. The method according to any one of clauses 1-13, wherein, while the aircraft is in flight, the following operations are performed more than once: determining the proximity to the temperature limit and controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit.

[0113] 15. The method according to any one of Clauses 1-14, wherein controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit includes controlling multiple electric propulsion units according to a priority scheme.

[0114] 16. The method according to any one of clauses 1-15,

[0115] The high-voltage line is configured to supply power to the first electric propulsion unit; and

[0116] Controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit includes reducing the thrust provided by the first electric propulsion unit compared to the second electric propulsion unit as the proximity to the temperature limit increases.

[0117] 17. The method according to any one of Clauses 1-16, wherein controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit comprises: degrading the battery state of the battery pack associated with the high-voltage line, and controlling the at least one electric propulsion unit based on the derated battery state.

[0118] 18. The method according to Clause 17, wherein the dated battery state is the dated power state of the battery pack associated with the high-voltage line.

[0119] 19. The method according to any one of Clauses 1-18,

[0120] Receiving the first sensor data includes receiving first sensor data for each of a plurality of high-voltage line segments associated with a plurality of battery packs.

[0121] Determining the proximity to the temperature limit includes determining the proximity of each high-voltage line segment to the corresponding temperature limit; and

[0122] Controlling at least one electric propulsion unit of the aircraft to prevent exceeding the temperature limit includes:

[0123] The battery state of each battery pack is adjusted based on the proximity of the corresponding high-voltage line segment of each battery pack to the corresponding temperature limit.

[0124] Power is distributed across multiple electric propulsion units based on the adjusted battery pack status.

[0125] 20. The method according to Clause 19, further comprising: wherein each of the plurality of high-voltage line segments includes a high-voltage line feeding the electric propulsion unit.

[0126] 21. The method according to any one of clauses 1-20, wherein controlling the at least one electric propulsion unit of the aircraft comprises: managing the power consumption of the at least one electric propulsion unit based on the determined proximity to prevent the electric propulsion unit from exceeding the temperature limit.

[0127] 22. The method according to any one of clauses 1-21, wherein controlling the at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit comprises: managing the power consumption of each of the plurality of electric propulsion units of the aircraft based on the determined proximity to prevent each electric propulsion unit from exceeding the temperature limit while maintaining the stability of the aircraft.

[0128] 23. A system for controlling an aircraft, the system comprising at least one processor configured to execute instructions to cause the system to perform any one of the provisions of 1-22.

[0129] 24. A computer-readable storage medium storing instructions that, when executed by at least one processor, cause the at least one processor to perform the method according to any one of clauses 1-22.

[0130] 26. An aircraft comprising:

[0131] First sensor;

[0132] At least one electric propulsion unit;

[0133] High-voltage lines;

[0134] At least one processor, which is configured to execute instructions to cause the system to perform any of the provisions of clauses 1-22.

[0135] The foregoing description has been presented for illustrative purposes. It is not exhaustive and does not limit the invention to the precise forms or embodiments disclosed herein. Modifications and adaptations of the invention will be apparent to those skilled in the art upon consideration of the description and practice of the disclosed embodiments of the invention.

[0136] The features and advantages of this disclosure are apparent from the detailed description; therefore, the appended terms are intended to cover all systems and methods falling within the true spirit and scope of this disclosure. As used herein, the indefinite articles “a” and “an” mean “one or more”. Similarly, the use of plural terms does not necessarily indicate a plural number unless this is explicit in the given context. Unless otherwise expressly stated, words such as “and” or “or” mean “and / or”. As used herein, unless otherwise expressly stated, “based on” can include dependent on, interdependent on, derived from (e.g., used), associated with, at least partially defined by, affected by, or in response to. As used herein, “related to” or “involved with” can include including, expressed by, indicated by, or based on. Furthermore, since many modifications and variations are readily apparent from studying this disclosure, it is not intended to limit this disclosure to the exact structures and operations shown and described; therefore, all suitable modifications and equivalents may be employed within the scope of this disclosure.

[0137] Other embodiments will be apparent to those skilled in the art upon consideration of the description and practice of the embodiments disclosed herein. The architectures and circuit arrangements shown in the figures are for illustrative purposes only and are not intended to limit the invention to the specific arrangements described and shown in the figures. The description and examples are also intended to be considered exemplary only, and the true scope and spirit of the invention are indicated by the following terms. The foregoing description has been presented for illustrative purposes. It is not exhaustive and does not limit the invention to the precise forms or embodiments disclosed. Modifications and adaptations of the invention will be apparent to those skilled in the art upon consideration of the description and practice of the disclosed embodiments of the invention disclosed herein. The sequence of steps shown in the figures is also for illustrative purposes only and is not intended to limit the invention to any particular sequence of steps. Therefore, those skilled in the art will understand that these steps may be performed in different orders while implementing the same method.

Claims

1. A method for controlling an aircraft, the method comprising: During flight, first sensor data is received, the first sensor data indicating at least one property of the high-voltage line of the aircraft; Receive the operational status of the aircraft during flight; During flight, the proximity of the temperature of the high-voltage line to the temperature limit is determined based on the data from the first sensor and the operating status of the aircraft. as well as At least one electric propulsion unit of the aircraft is controlled based on a determined proximity to avoid exceeding the temperature limit.

2. The method of claim 1, wherein the at least one property includes at least one of the following: the temperature of the high-voltage line, the ambient temperature of the wing, or the current passing through the high-voltage line.

3. The method of claim 1, wherein the at least one attribute includes the temperature of the high-voltage line received from a temperature sensor on the high-voltage line, and wherein the high-voltage line is configured to provide power to at least one of: an electric propulsion unit or a battery pack.

4. The method of claim 1, wherein the at least one property includes the wing ambient temperature and the current through the high-voltage line.

5. The method of claim 1, wherein the operating state includes at least one of a flight phase or an emergency situation state.

6. The method of claim 5, wherein the operating state includes an emergency state, the emergency state indicating whether the battery pack is providing backup power.

7. The method of claim 1, wherein the operating state includes a flight phase, the flight phase including at least one of: hovering, takeoff, or forward flight.

8. The method of claim 1, wherein the operating state is based on at least one of the following: airspeed measurement, altitude measurement, ground speed measurement, propeller tilt angle measurement, or GPS measurement.

9. The method according to claim 1, The operational state described therein includes at least two upcoming flight phases; and Determining the proximity to the temperature limit includes taking into account the impact of the at least two upcoming flight phases on the temperature of the high-voltage line.

10. The method of claim 1, wherein the proximity to the temperature limit is determined based on at least one of a model or a lookup table.

11. The method of claim 1, wherein the proximity to the temperature limit is based on the proximity to a curve relating to (the current of the high-voltage line)^2. The temperature of the high-voltage line is proportional to the temperature of the line.

12. The method of claim 1, wherein the proximity to the temperature limit is determined based on a model or lookup table, and wherein the input to the model or lookup table includes: The temperature of the high-voltage line, the current flowing through the high-voltage line, and the flight phase of the aircraft.

13. The method according to claim 1, wherein the proximity to the temperature limit is the time remaining until the high-voltage line temperature reaches the temperature limit.

14. The method of claim 1, wherein while the aircraft is in flight, the following operations are performed more than once: determining the proximity to the temperature limit and controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit.

15. The method of claim 1, wherein controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit comprises controlling multiple electric propulsion units according to a priority scheme.

16. The method according to claim 1, The high-voltage line is configured to supply power to the first electric propulsion unit; and Controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit includes: As the determined proximity to the temperature limit increases, the thrust provided by the first electric propulsion unit decreases compared to the second electric propulsion unit.

17. The method of claim 1, wherein controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit comprises: The battery state of the battery pack associated with the high-voltage line is derated, and the at least one electric propulsion unit is controlled based on the derated battery state.

18. The method of claim 17, wherein the dated battery state is the dated power state of the battery pack associated with the high-voltage line.

19. The method according to claim 1, Receiving the first sensor data includes receiving first sensor data for each of a plurality of high-voltage line segments associated with a plurality of battery packs. Determining the proximity to the temperature limit includes determining the proximity of each high-voltage line segment to the corresponding temperature limit; and Controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit includes: The battery pack status of each battery pack is adjusted based on the proximity of the corresponding high-voltage line segment of each battery pack to the corresponding temperature limit. Power is distributed across multiple electric propulsion units based on the adjusted battery pack status.

20. The method of claim 19, further comprising: Each of the plurality of high-voltage line segments includes a high-voltage line that supplies power to the electric propulsion unit.

21. The method of claim 1, wherein controlling the at least one electric propulsion unit of the aircraft comprises: The power consumption of the at least one electric propulsion unit is managed based on the determined proximity to prevent the electric propulsion unit from exceeding the temperature limit.

22. The method of claim 1, wherein controlling the at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit comprises: The power consumption of each of the multiple electric propulsion units of the aircraft is managed based on the determined proximity to prevent each electric propulsion unit from exceeding the temperature limit, while maintaining the stability of the aircraft.

23. A system for controlling an aircraft, the system comprising at least one processor configured to execute instructions to cause the system to perform the following steps: During flight, first sensor data is received, the first sensor data indicating at least one property of the high-voltage line of the aircraft; Receive the operational status of the aircraft during flight; During flight, the proximity of the high-voltage line's temperature to a temperature limit is determined based on the data from the first sensor and the operating status of the aircraft; and At least one electric propulsion unit of the aircraft is controlled based on a determined proximity to avoid exceeding the temperature limit.

24. A computer-readable storage medium storing instructions, which, when executed by at least one processor, cause the at least one processor to perform the following steps: During flight, first sensor data is received, which indicates at least one property of the aircraft's high-voltage lines; Receive the operational status of the aircraft during flight; During flight, the proximity of the high-voltage line's temperature to a temperature limit is determined based on the data from the first sensor and the operating status of the aircraft; and At least one electric propulsion unit of the aircraft is controlled based on a determined proximity to avoid exceeding the temperature limit.

25. An aircraft comprising: First sensor; At least one electric propulsion unit; High-voltage lines; At least one processor, the at least one processor being configured to execute instructions to perform the following steps: During flight, sensor data from the first sensor is received, the sensor data indicating at least one attribute of the high-voltage line; Receive the operational status of the aircraft during flight; During flight, the proximity of the temperature of the high-voltage line to the temperature limit is determined based on sensor data from the first sensor and the operating status of the aircraft. as well as At least one electric propulsion unit of the aircraft is controlled based on a determined proximity to avoid exceeding the temperature limit.

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