Control method of electric aircraft, electric aircraft and computer program product

Abstract

The invention discloses a control method of an electric aircraft, the electric aircraft and a computer program product, and relates to the technical field of aircraft control. The method comprises the following steps: acquiring an acceleration signal and rotor operation data of the electric aircraft; carrying out falling collision detection analysis on the acceleration signal and the rotor wing operation data, and determining a detection result; in response to the detection result that the electric aircraft is crashed abnormally, the cut-off switch is controlled to be switched to an off state within a first duration range, the closed switch is controlled to be switched to a closed state within a second duration range, the cut-off switch is arranged on a high-voltage main loop in a power supply system of the electric aircraft, and the closed switch is arranged on the high-voltage main loop. And the closed switch is arranged on a high-voltage main loop in a power distribution system of the electric aircraft and is connected in series with the discharge resistor. The technical problems that in the flying process of the electric aircraft, falling collision recognition difficulty is large, and safety is low are solved.

Description

Technical Field

[0001] The present application relates to the field of aircraft control technology, and in particular to a control method for an electric aircraft, an electric aircraft, and a computer program product. Background Art

[0002] Electric aircraft may fall or collide (collectively referred to as crash) during flight. If appropriate measures are not taken in a timely manner after a crash, secondary damage is likely to occur. For example, the electric aircraft may leak electricity, causing the pilot and passengers to face the risk of electric shock, or causing secondary injuries to rescue personnel.

[0003] In this regard, although some aircraft anti-crash control methods have been proposed in the relevant technologies, most of them are aimed at traditional fuel aircraft and lack solutions for accurate monitoring and identification of crash events. However, the internal high-voltage electrical system and control system of electric aircraft are not completely different from those of fuel aircraft. Therefore, the solutions provided by the relevant technologies are difficult to apply to the flight control scenarios of electric aircraft.

[0004] As can be seen from the above, how to accurately identify the electric aircraft crash and how to control the electric aircraft to improve safety in the electric aircraft crash scenario is an important technical problem in the relevant technical field. Currently, no effective solution has been proposed for the above problems. Summary of the invention

[0005] The embodiments of the present application provide a control method for an electric aircraft, an electric aircraft, and a computer program product to at least solve the technical problems of difficulty in crash identification and low safety during the flight of an electric aircraft.

[0006] According to one aspect of an embodiment of the present application, a control method for an electric aircraft is provided, including: obtaining an acceleration signal and rotor operation data of the electric aircraft, wherein the acceleration signal is collected by a flight acceleration sensor provided on the electric aircraft, and the rotor operation data is obtained from a power domain controller of the electric aircraft through a controller local area network bus; performing crash detection analysis on the acceleration signal and the rotor operation data to determine a detection result; in response to the detection result that the electric aircraft has a crash anomaly, controlling a disconnect switch to switch to an open state within a first time range, and controlling a closing switch to switch to a closed state within a second time range, wherein the disconnect switch is arranged on a high-voltage main circuit in a power supply system of the electric aircraft, and the closing switch is arranged on a high-voltage main circuit in a power distribution system of the electric aircraft, and the closing switch is connected in series with a discharge resistor, and the discharge resistor is used to assist in reducing the voltage of the high-voltage main circuit to a safe voltage range.

[0007] Optionally, a crash detection analysis is performed on the acceleration signal and the rotor operation data to determine the detection result, including: determining multiple acceleration components of the electric aircraft in multiple directions based on the acceleration signal; in response to the multiple acceleration components and the rotor operation data satisfying abnormal conditions, determining that the detection result is that the electric aircraft has a crash anomaly, wherein the abnormal condition is determined by multiple crash thresholds corresponding to the crash anomaly.

[0008] Optionally, the rotor operating data includes rotor speed and rotor torque; multiple acceleration components and rotor operating data satisfy abnormal conditions including at least one of the following: the absolute value of at least one acceleration component among the multiple acceleration components is greater than the acceleration crash threshold; the change in the rotor speed within a preset detection period is greater than the speed crash threshold; the change in the rotor torque within a preset detection period is greater than the torque crash threshold.

[0009] Optionally, the first duration range is smaller than the second duration range.

[0010] Optionally, the control method of the electric aircraft also includes: obtaining the real-time flight altitude of the electric aircraft; in response to the detection result that the electric aircraft has a crash abnormality and the real-time flight altitude is within a preset altitude range, controlling the disconnect switch to switch to an open state, and controlling the closing switch to switch to a closed state.

[0011] Optionally, the control method of the electric aircraft further includes: in response to the detection result that the electric aircraft has a crash anomaly, triggering a motor controller of the electric aircraft to perform an active discharge action.

[0012] Optionally, the control method of the electric aircraft further includes: in response to the detection result that the electric aircraft has a crash abnormality, triggering a battery management system of the electric aircraft to disconnect a main positive contactor and a main auxiliary contactor corresponding to the battery module.

[0013] Optionally, the electric aircraft is an electric flying car, an electric airplane or an electric drone.

[0014] According to another aspect of an embodiment of the present application, an electric aircraft is also provided, including: a power supply system, a power distribution system, an electric drive module, a power domain controller and a crash detection control module, wherein the power supply system and the power distribution system are electrically connected, the power supply system and the power domain controller are electrically connected, the power distribution system and the electric drive module are electrically connected, the power domain controller and the electric drive module are electrically connected, the crash control module and the power domain controller are electrically connected, the crash control module and the power domain controller are electrically connected, the crash control module and the power supply system are electrically connected, and the crash control module and the power distribution system are electrically connected; the power supply system includes a battery module and a battery management unit; the distribution system includes a DC converter, a distribution box and an on-board charging unit; the electric drive module includes a motor controller, an electric drive and a rotor; the crash detection control module includes an acceleration sensor, a crash processor, a disconnect switch, a closing switch and a discharge resistor; the disconnect switch is arranged on the high-voltage main circuit in the power supply system, the closing switch is arranged on the high-voltage main circuit in the distribution system, the closing switch is connected in series with the discharge resistor, and the discharge resistor is used to assist in reducing the voltage of the high-voltage main circuit to a safe voltage range; the crash processor is used to execute a computer program to implement any one of the above-mentioned control methods for an electric aircraft.

[0015] According to another aspect of an embodiment of the present application, a computer program product is also provided, including a computer program, and when the computer program is executed by a processor, the control method of the electric aircraft according to any one of the above items is implemented.

[0016] In an embodiment of the present application, an acceleration signal and rotor operation data of an electric aircraft are obtained, wherein the acceleration signal is collected by a flight acceleration sensor provided on the electric aircraft, and the rotor operation data is obtained from a power domain controller of the electric aircraft through a controller local area network bus; a crash detection analysis is performed on the acceleration signal and the rotor operation data to determine a detection result; in response to the detection result that the electric aircraft has a crash abnormality, the disconnect switch is controlled to switch to an open state within a first time range, and the closing switch is controlled to switch to a closed state within a second time range, wherein the disconnect switch is arranged on a high-voltage main circuit in a power supply system of the electric aircraft, and the closing switch is arranged on a high-voltage main circuit in a power distribution system of the electric aircraft, and the closing switch is connected in series with a discharge resistor, and the discharge resistor is used to assist in reducing the voltage of the high-voltage main circuit to a safe voltage range. Therefore, in the present application, based on the acceleration signal and rotor operation data of the electric aircraft, detection and analysis are performed to determine whether the electric aircraft has crashed. When it is determined that a crash has occurred, the electric aircraft is controlled after the crash through the built-in disconnect switch, closing switch and discharge resistor of the electric aircraft, and the voltage of the high-voltage main circuit of the electric aircraft is quickly reduced to a safe voltage range, thereby achieving the purpose of accurately detecting the crash of the electric aircraft and automatically controlling the voltage reduction of the high-voltage main circuit of the electric aircraft, thereby achieving the technical effect of improving the accuracy of crash identification during the flight of the electric aircraft and enhancing the safety of the application scenarios of the electric aircraft, thereby solving the technical problems of difficulty in crash identification and low safety during the flight of the electric aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] The drawings described herein are used to provide a further understanding of the present application and constitute a part of the present application. The illustrative embodiments of the present application and the related descriptions of the embodiments are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

[0018] Figure 1 is a hardware structure block diagram of an optional computing terminal for implementing a control method for an electric aircraft according to an embodiment of the present application;

[0019] Figure 2 is a flow chart of a control method of an electric aircraft according to an embodiment of the present application;

[0020] Figure 3 is a schematic diagram of a system architecture of an optional electric aircraft according to an embodiment of the present application;

[0021] Figure 4 It is a structural block diagram of an electric aircraft according to an embodiment of the present application. DETAILED DESCRIPTION

[0022] In order to enable those skilled in the art to better understand the solution of the present application, the technical solution in the embodiments of the present application will be clearly and completely described below in conjunction with the drawings in the embodiments of the present application. Obviously, the described embodiments only include a part of the embodiments of the present application, rather than all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without creative work should fall within the scope of protection of this application.

[0023] It should be noted that the terms "first", "second", etc. in the specification and claims of the present application and the above-mentioned drawings are used to distinguish similar objects, and are not necessarily used to describe a specific order or sequence. It should be understood that the data used in this way can be interchangeable where appropriate, so that the embodiments of the present application described herein can be implemented in an order other than those illustrated or described herein. In addition, the terms "including" and "having" and any of their variations are intended to cover non-exclusive inclusions, for example, a process, method, system, product or device comprising a series of steps or units is not necessarily limited to those steps or units clearly listed, but may include other steps or units that are not clearly listed or inherent to these processes, methods, products or devices.

[0024] According to an embodiment of the present invention, a method embodiment of a control method for an electric aircraft is provided. It should be noted that the steps shown in the flowchart of the accompanying drawings can be executed in a computer system such as a set of computer executable instructions, and although a logical order is shown in the flowchart, in some cases, the steps shown or described can be executed in an order different from that shown here.

[0025] Exemplarily, the operating environment of the above method embodiment is described. Figure 1 is a hardware structure block diagram of a computing terminal for implementing a control method for an electric aircraft according to an embodiment of the present application, such as Figure 1 As shown, a computing terminal 10 (for example, a computer terminal, a mobile smart terminal, a vehicle terminal, a drone terminal, a flying car terminal, or a cloud computing virtual terminal, etc.) may include: one or more processors 102, a memory 104 for storing data, and a transmission device 106 for implementing a communication function, wherein each processor 102 may include but is not limited to a processing component such as a microprocessor (Microcontroller Unit, referred to as MCU) or a programmable logic device (Field Programmable Gate Array, referred to as FPGA).

[0026] The computing terminal 10 may further include: a display device 110, an input / output device 108, a Universal Serial Bus (USB) port (the USB port may be used as one of the ports of a computer bus, not shown in the figure), a network interface (not shown in the figure), a power supply (not shown in the figure) and a camera (not shown in the figure). A person skilled in the art may understand that Figure 1 The structure of the computing terminal 10 shown is only for illustration and does not impose a strict limitation on the structure of the computing terminal 10. For example, the computing terminal 10 may also include Figure 1 More or fewer components than those shown in FIG. 1 , or computing terminal 10 may have the same components as those shown in FIG. Figure 1 Different categories of components are shown.

[0027] It should be noted that the one or more processors 102 and / or other data processing circuits in the computing terminal 10 described above may be embodied in whole or in part as software, hardware, firmware, or any other combination thereof. In addition, the data processing circuit may be a single independent processing module, or may be incorporated in whole or in part into any of the other components in the computing terminal 10 (or mobile device).

[0028] The memory 104 can be used to store software programs and modules of application software, such as the program instructions / data storage device corresponding to the control method of the electric aircraft in the embodiment of the present application. The processor 102 executes various functional applications and data processing by running the software programs and modules stored in the memory 104, that is, realizing the control method of the electric aircraft mentioned above. The memory 104 may include a high-speed random access memory, and may also include a non-volatile memory, such as one or more magnetic storage devices, flash memory, or other non-volatile solid-state memory. In some examples, the memory 104 may further include a memory remotely arranged relative to the processor 102, and these remote memories may be connected to the computing terminal 10 via a network. Examples of the above-mentioned network include, but are not limited to, the Internet, an intranet, a local area network, a mobile communication network, and a combination thereof.

[0029] The transmission device 106 is used to receive or send data via a network. The specific example of the above network may include a wireless network provided by a communication provider of the computing terminal 10. In one example, the transmission device 106 includes a network adapter (Network Interface Controller, referred to as NIC), which can be connected to other network devices through a base station so as to communicate with the Internet. In one example, the transmission device 106 can be a radio frequency (Radio Frequency, referred to as RF) module, which is used to communicate with the Internet wirelessly.

[0030] An embodiment of the present application provides a control method for an electric aircraft. The control method for an electric aircraft can be used to provide control functions of an electric aircraft for preset application scenarios. The preset application scenarios may include the following scenarios in the fields of flying cars, drones, and manned aircraft: commuting autonomous driving scenarios, artificial intelligence (AI) driving scenarios, automatic parking assist (APA) scenarios, and intelligent navigation assistance (Navigation Guided Pilot, NGP) scenarios in urban areas or high-speed areas. In addition, the preset application scenarios may also include, but are not limited to: intelligent driving cargo aircraft scenarios in the field of logistics and transportation, autonomous driving agricultural aircraft scenarios in the field of agricultural machinery, and intelligent flying robots (such as cleaning robots, service robots, delivery robots, etc.) scenarios.

[0031] Under the above operating environment, the embodiment of the present application provides the following Figure 2 The control method of the electric aircraft shown, Figure 2 is a flow chart of a control method of an electric aircraft according to an embodiment of the present application. Figure 2 As shown, the method includes the following implementation steps S201 to S203.

[0032] Step S201, acquiring the acceleration signal and rotor operation data of the electric aircraft, wherein the acceleration signal is collected by a flight acceleration sensor provided on the electric aircraft, and the rotor operation data is obtained from the power domain controller of the electric aircraft through a controller area network bus.

[0033] The electric aircraft may be an aircraft that uses electricity as a main power source, such as an electric vertical take-off and landing vehicle (eVTOL), an electric manned aircraft, etc.

[0034] The above acceleration signal may refer to the acceleration change in different directions during the movement of the aircraft (including acceleration, deceleration and impact). On electric aircraft, acceleration sensors are used for collection. These sensors are usually installed in key parts of the aircraft, such as wings, fuselage, rotor drive system, etc., to accurately capture the dynamic changes of the aircraft. Acceleration signals can be used to assist in determining whether the aircraft has encountered a crash or abnormal situation.

[0035] The above-mentioned flight acceleration sensor can be a sensor used to monitor the acceleration change of the aircraft. In an electric aircraft, it can monitor the acceleration of the aircraft in different axes (including but not limited to the X-axis, Y-axis and Z-axis, i.e. the lateral, longitudinal and vertical directions) in real time. The data of the flight acceleration sensor is one of the data input sources of the crash detection system.

[0036] The above rotor operation data may include parameters such as rotor speed, torque, vibration, etc., and may also include temperature and pressure information. These rotor operation data are collected by sensors of the rotor system (such as rotor speed sensor, torque sensor, etc.) and transmitted through the control system of the aircraft to evaluate the stability and health status of the aircraft.

[0037] The controller area network bus (CAN bus) is a data bus used for communication between electronic devices and is widely used in control systems in the fields of automobiles, aircraft, etc. In electric aircraft, the CAN bus is used to connect different controllers, such as power domain controllers, motor controllers, battery management systems, etc., to achieve real-time data sharing and efficient communication.

[0038] In the application scenario, the power domain controller of the electric aircraft actively pulls or receives the rotor operation data sent by other controllers (such as the rotor controller) on the electric aircraft through the CAN bus. This data transmission method realizes real-time monitoring and centralized processing of sensor data.

[0039] Step S202, performing crash detection analysis on the acceleration signal and the rotor operation data to determine the detection result.

[0040] The above-mentioned crash detection analysis is a process of determining whether the aircraft has encountered a crash event by real-time monitoring of the aircraft's acceleration signal and rotor operation data, combined with preset thresholds and algorithms. Crash detection analysis is usually performed by the central control system or collision processor of the electric aircraft.

[0041] In the application scenario, after completing the crash detection analysis, the electric aircraft will generate a crash event detection report based on the analysis results, that is, the detection result, which is used to characterize whether the electric aircraft has a crash event. The detection result can assist in determining whether to start the high-voltage electrical isolation and active discharge procedures.

[0042] Step S203, in response to the detection result that the electric aircraft has a crash abnormality, the disconnect switch is controlled to switch to an open state within a first time range, and the closing switch is controlled to switch to a closed state within a second time range, wherein the disconnect switch is arranged on a high-voltage main circuit in the power supply system of the electric aircraft, and the closing switch is arranged on a high-voltage main circuit in the power distribution system of the electric aircraft, and the closing switch is connected in series with a discharge resistor, and the discharge resistor is used to assist in reducing the voltage of the high-voltage main circuit to a safe voltage range.

[0043] The disconnect switch can be a fast disconnect switch, which is in a closed state by default. When certain conditions are met, the fast disconnect switch can be quickly switched to an open state within a short time. The disconnect switch is located on the high-voltage main circuit in the power supply system and is used to quickly disconnect the high-voltage power supply in an emergency, thereby disconnecting the high-voltage power system and reducing the risk of electric shock.

[0044] The above-mentioned closing switch can be a fast closing switch. In the default state, the fast disconnect switch is in the open state. When specific conditions are met, the fast closing switch can be quickly switched to the closed state within a short time. The closing switch is located on the high-voltage main circuit in the power distribution system and is connected in series with the discharge resistor. After the crash detection system confirms that a crash has occurred, the closing switch is activated to form a discharge circuit, which quickly dissipates the electrical energy in the high-voltage circuit through the discharge resistor and reduces the voltage to a safe level.

[0045] The above-mentioned discharge resistor is a component used to quickly convert electrical energy in a high-voltage circuit into heat energy. After the high-voltage main circuit is isolated, the discharge resistor forms a closed loop with the high-voltage main circuit by closing the switch, and the residual electrical energy is quickly released to reduce the voltage to a safe range.

[0046] In the application scenario, when the crash detection system confirms that a crash has occurred, the electric aircraft will immediately control the disconnect switch to switch to the disconnected state within the first time range (for example, 4ms) to quickly isolate the high-voltage power supply. At the same time, the closing switch switches to the closed state within the second time range (for example, 6ms), forming an effective discharge circuit with the discharge resistor and starting to quickly discharge the high-voltage main circuit.

[0047] In the application scenario, by real-time monitoring of acceleration signals and rotor operation data, and combining fast-response high-voltage electrical isolation and active discharge mechanisms, the electric aircraft safety architecture proposed in this application can: quickly identify and respond to crash anomalies in a very short time after a crash, and immediately cut off the high-voltage power supply, thereby minimizing the risk of electric shock and fire; achieve rapid active discharge through discharge resistors, quickly reduce the high-voltage circuit to a safe voltage range, and protect the driver, passengers and rescue personnel from secondary electric shock injuries; improve the safety and reliability of the entire system, provide solid technical guarantees for the safety of personnel in electric aircraft in emergency situations, and fill the gaps in the existing technology for crash electrical isolation response speed and efficiency; realize instant safety isolation and active discharge of high-voltage electrical systems, significantly improving the safety protection capabilities of electric aircraft in emergency situations such as crashes.

[0048] Through the technical solutions provided by the above steps S201 to S203, in an embodiment of the present application, the acceleration signal and rotor operation data of the electric aircraft are obtained, wherein the acceleration signal is collected by a flight acceleration sensor provided on the electric aircraft, and the rotor operation data is obtained from the power domain controller of the electric aircraft through the controller local area network bus; crash detection and analysis are performed on the acceleration signal and the rotor operation data to determine the detection result; in response to the detection result that the electric aircraft has a crash abnormality, the disconnect switch is controlled to switch to an open state within a first time range, and the closing switch is controlled to switch to a closed state within a second time range, wherein the disconnect switch is arranged on a high-voltage main circuit in the power supply system of the electric aircraft, and the closing switch is arranged on a high-voltage main circuit in the power distribution system of the electric aircraft, and the closing switch is connected in series with a discharge resistor, and the discharge resistor is used to assist in reducing the voltage of the high-voltage main circuit to a safe voltage range. Therefore, in the present application, based on the acceleration signal and rotor operation data of the electric aircraft, detection and analysis are performed to determine whether the electric aircraft has crashed. When it is determined that a crash has occurred, the electric aircraft is controlled after the crash through the built-in disconnect switch, closing switch and discharge resistor of the electric aircraft, and the voltage of the high-voltage main circuit of the electric aircraft is quickly reduced to a safe voltage range, thereby achieving the purpose of accurately detecting the crash of the electric aircraft and automatically controlling the voltage reduction of the high-voltage main circuit of the electric aircraft, thereby achieving the technical effect of improving the accuracy of crash identification during the flight of the electric aircraft and enhancing the safety of the application scenarios of the electric aircraft, thereby solving the technical problems of difficulty in crash identification and low safety during the flight of the electric aircraft.

[0049] In an exemplary application scenario, a method is provided as follows: Figure 3 The system architecture of the electric aircraft is shown in FIG. Figure 3As shown, acceleration sensor 1 and acceleration sensor 2 are used to collect acceleration signals of the electric aircraft. The power domain controller may include rotor operation data of the electric aircraft. The control method of the electric aircraft provided in the embodiment of the present application may be run on a collision processor (including Figure 3 The above-mentioned acceleration sensor 1 and acceleration sensor 2 can be used to collect acceleration signals of the electric aircraft in different directions, and can also be used to collect acceleration signals corresponding to different components of the electric aircraft.

[0050] like Figure 3 As shown, the acceleration signal is transmitted between the acceleration sensor and the collision processor through pulse width modulation (PWM), and the collision processor can obtain the rotor operation data from the power domain controller through the CAN bus.

[0051] Still like Figure 3 As shown, the collision processor (including collision processor 1 and collision processor 2) is connected to the fast closing switch (such as Figure 3 The switch control 1 corresponding to K1 in the figure is connected to the fast cut-off switch (such as Figure 3 When it is determined that the electric aircraft has crashed, the collision processor can control the fast closing switch to switch to the closed state through switch control 1, and can also control the fast disconnecting switch to switch to the disconnected state through switch control 2.

[0052] like Figure 3 As shown, the power distribution unit (i.e., the power distribution system) of the electric aircraft may also include: a fuse R1 corresponding to the air conditioning heater, a fuse R2 corresponding to the air conditioning compressor, a fuse R3 corresponding to the DC converter (DCDC), and a fuse R4 corresponding to the on-board charger (OBC). In addition, in the power distribution unit, the fast discharge resistor connected in series with the fast closing switch is denoted as R5, and the main fuse R6 and main fuse R7 are also provided on the high-voltage main circuit. In addition, an electric drive fuse R8 is also provided between the power distribution unit and the motor controller.

[0053] like Figure 3 As shown, the motor controller of the electric aircraft also includes a microprocessor and a drive motor. The motor controller is connected to the power domain controller via a CAN bus for signal transmission. The power domain controller is connected to the battery management system of the power supply system via a CAN bus.

[0054] As an optional implementation, in the above step S202, performing crash detection analysis on the acceleration signal and the rotor operation data to determine the detection result may also include the following execution steps:

[0055] Step S221, determining multiple acceleration components of the electric aircraft in multiple directions according to the acceleration signal;

[0056] Step S222, in response to multiple acceleration components and rotor operation data satisfying abnormal conditions, determining that the detection result is that the electric aircraft has a crash anomaly, wherein the abnormal condition is determined by multiple crash thresholds corresponding to the crash anomaly.

[0057] The above-mentioned multiple directions include the X direction, Y direction and Z direction in three-dimensional space. Specifically, when the electric aircraft moves in three-dimensional space, the acceleration can be decomposed into three acceleration components along the X axis, Y axis and Z axis. The X axis usually represents the lateral movement of the electric aircraft, the Y axis represents the longitudinal movement of the electric aircraft, and the Z axis represents the vertical movement of the electric aircraft. Electric aircraft crashes are usually accompanied by abnormal acceleration changes in multiple directions, such as rapid descent acceleration in the vertical direction and possible rotational acceleration in the lateral and longitudinal directions.

[0058] In the application scenario, the flight acceleration sensor continuously monitors the acceleration signal of the aircraft, converts the acceleration signal into a digital signal, and transmits it to the central control system or collision processor in real time through the CAN bus. The control system or processor decomposes the received acceleration signal into acceleration components in the X-axis, Y-axis and Z-axis directions for further analysis.

[0059] The abnormal conditions may include a set of criteria for identifying whether the acceleration signal and rotor operation data contain characteristics of a crash event. These abnormal conditions are set based on an analysis of normal flight data and preset thresholds for acceleration and rotor state changes that may be caused by a crash event.

[0060] The above-mentioned crash thresholds may be indicators for determining whether the electric aircraft has encountered a crash. These crash thresholds may include, but are not limited to, the maximum value, minimum value, or rate of change of the acceleration component, and may also include abnormal speed, torque, or vibration level in the rotor operation data. When any one or more of the acceleration signal and the rotor operation data exceeds these preset crash thresholds, it can be considered that the electric aircraft may encounter a crash event.

[0061] The central control system or collision processor of the electric aircraft continuously analyzes the received acceleration components and rotor operation data. When any one or more of these data exceeds the preset crash threshold, it is considered that the abnormal condition is met, and the central control system or collision processor will immediately determine that the electric aircraft may be experiencing or has experienced a crash event.

[0062] After confirming that the acceleration component and rotor operation data meet the abnormal conditions, the collision processor will determine that the detection result is an abnormal crash of the electric aircraft. Subsequently, the high-voltage electrical isolation and active discharge process will be initiated to quickly reduce the high-voltage circuit in the aircraft to a safe state to protect personnel safety.

[0063] Through the technical solutions provided by the above steps S221 to S222, the embodiment of the present application significantly improves the response speed and protection effect to crash events through real-time monitoring and analysis of comprehensive acceleration signals and rotor operation data, and provides strong data analysis technical support for the safe operation of electric aircraft.

[0064] As an optional implementation, in the above step S222, the rotor operation data includes the rotor speed and the rotor torque; and the multiple acceleration components and the rotor operation data satisfying the abnormal condition include at least one of the following:

[0065] The absolute value of at least one acceleration component among the multiple acceleration components is greater than the acceleration crash threshold;

[0066] The change in rotor speed within a preset detection period is greater than the speed crash threshold;

[0067] The change in rotor torque within a preset detection period is greater than the torque crash threshold.

[0068] The above rotor operation data is a set of information reflecting the current status of the rotor provided by the rotor system of the electric aircraft, mainly including rotor speed and rotor torque. These rotor operation data can not only reflect the flight performance of the electric aircraft, but also provide important clues in abnormal situations to help the system quickly diagnose problems.

[0069] The above rotor speed refers to the number of revolutions per minute (rpm) of the rotor of the electric aircraft. The normal operation of the rotor needs to be maintained within a stable speed range to ensure that the electric aircraft maintains appropriate lift and flight stability.

[0070] Rotor torque refers to the torque that the rotor shaft of an electric aircraft bears when driving the rotor to rotate, reflecting the force output by the motor to the rotor shaft. The magnitude of rotor torque affects the speed of the rotor and the maneuverability of the electric aircraft.

[0071] The above-mentioned multiple acceleration components include acceleration components of the electric aircraft along the lateral direction (X axis), longitudinal direction (Y axis) and vertical direction (Z axis).

[0072] In the operational application scenario of an electric aircraft, if at least one of the following situations a, b and c is detected, it will be determined that the crash abnormality condition is met.

[0073] Case a: The absolute value of at least one of the multiple acceleration components is greater than the acceleration crash threshold (e.g., 5m / s^2). This means that the electric aircraft has experienced an abnormal acceleration change in a certain direction, such as a sudden descent or violent lateral movement. This abnormality is usually closely related to a crash event.

[0074] Case b: The change in rotor speed within the preset detection period is greater than the speed crash threshold. A sharp drop or irregular change in rotor speed may indicate that the rotor system has suffered physical damage, which is a significant sign of a crash event.

[0075] Case c: The change in rotor torque within the preset detection period is greater than the torque crash threshold. The abnormal fluctuation of rotor torque also suggests that the electric aircraft may have encountered an external impact beyond the normal operating range, which is particularly obvious in a crash scenario.

[0076] Through the above-mentioned technical solution for determining whether abnormal conditions are met, the embodiments of the present application can significantly improve the electrical system safety and personnel protection level of electric aircraft in crash events by real-time monitoring of acceleration components and rotor operation data and rapid response based on preset abnormal conditions, making important contributions to the scene application of electric aircraft and the improvement of safety flight standards.

[0077] As an optional implementation, in the above-mentioned control method of the electric aircraft, the first time range is smaller than the second time range.

[0078] The first duration range is set to T+4ms, and the second duration range is set to T+6ms.

[0079] The first time range mentioned above can be the time interval from the detection of abnormal signals to the system response and execution of high-voltage electrical isolation after the electric aircraft crashes abnormally. Within the first time range, the system needs to quickly determine whether it is a crash event and immediately activate the protection mechanism, such as disconnecting the main contactor to cut off the high-voltage power supply to avoid the risk of electric shock. For example, this time range can be set to T+4ms, that is, after the crash signal is detected (recorded as time T), high-voltage electrical isolation is performed within 4 milliseconds.

[0080] The above-mentioned second time range refers to the time interval from the detection of the abnormal crash signal to the completion of active discharge of the electric aircraft, reducing the voltage of the high-voltage circuit to a safe level. The control operation performed within the second time range usually involves disconnecting the connection between the battery and the motor to form a discharge loop, converting the electrical energy in the circuit into heat energy through the discharge resistor and releasing it to ensure that the voltage drops to a safe range that will not cause harm to personnel. For example, the second time range can be set to complete active discharge within 6 milliseconds to several seconds after the crash signal is detected (recorded as time T), and reduce the voltage to below 60V (DC).

[0081] Through the above-mentioned technical scheme, the setting of the first time range and the second time range in the embodiment of the present application ensures that when the electric aircraft encounters a crash anomaly, it can quickly start high-voltage electrical isolation and complete active discharge within a safe time, which has important technical effects on improving personnel safety, optimizing emergency response processes, improving rescue efficiency and reducing system damage.

[0082] As an optional implementation, the control method of the electric aircraft may further include the following execution steps:

[0083] Step S2041, obtaining the real-time flight altitude of the electric aircraft;

[0084] Step S2042, in response to the detection result that the electric aircraft has a crash anomaly and the real-time flight altitude is within a preset altitude range, the disconnect switch is controlled to switch to an open state, and the close switch is controlled to switch to a closed state.

[0085] The above-mentioned real-time flight altitude can be the current flight altitude data provided by the altitude sensor or positioning equipment when the electric aircraft is flying in the air. The real-time flight altitude can be used to evaluate the flight status and environmental conditions of the electric aircraft during the crash detection and response process.

[0086] The above-mentioned preset altitude range can be a height interval pre-set by the system according to the safety design and flight characteristics of the electric aircraft. When the real-time flight altitude of the electric aircraft falls into this range, the system will consider that the electric aircraft is in a specific flight stage or environment and needs to take specific safety measures.

[0087] In the application scenario, the system checks the real-time flight altitude of the electric aircraft while confirming the crash event. If the real-time flight altitude is within the preset crash response altitude range, the system will immediately initiate the safety response procedure, including controlling the disconnect switch to open quickly, cutting off the high-voltage power supply, and controlling the close switch to close, forming a discharge circuit with the discharge resistor, and quickly reducing the voltage in the high-voltage circuit to a safe range.

[0088] Through the technical solutions provided by the above steps S2041 to S2042, the embodiments of the present application can intelligently determine whether it is necessary to immediately start the high-voltage electrical isolation and active discharge procedures by combining the crash detection results and the flight altitude information, thereby avoiding the execution of safety responses at inappropriate flight stages or altitudes, thereby improving the accuracy and effectiveness of the response. When a crash event is confirmed and the electric aircraft is detected to be within a preset altitude range, the system can respond quickly, cut off the high-voltage power supply and perform active discharge, effectively preventing electric shock accidents and fire risks, and protecting the lives of drivers, passengers and rescuers. In addition, the introduction of real-time flight altitude data enables electric aircraft to flexibly adjust safety response strategies according to environmental changes, which is of great significance for improving the safety, reliability and overall flight performance of electric aircraft.

[0089] As an optional implementation, the control method of the electric aircraft may further include the following execution steps:

[0090] Step S205 , in response to the detection result that the electric aircraft has a crash anomaly, triggering a motor controller of the electric aircraft to perform an active discharge action.

[0091] In the application scenario, when the crash detection system determines that at least one of the acceleration component, rotor speed and torque change, and real-time flight altitude of the electric aircraft exceeds a preset abnormality threshold, the system will determine that the electric aircraft has encountered a crash abnormality.

[0092] The above-mentioned motor controller unit (MCU) is a component used to control the operating state of the motor in the electric drive system of an electric aircraft. For example, the MCU can be used to control the start, stop, speed adjustment of the motor, and perform active discharge in an emergency.

[0093] The above-mentioned active discharge action may refer to the process in which, after detecting a crash abnormality signal, the motor controller rapidly releases the electrical energy in the high-voltage electrical system to below a safe level through an internal discharge circuit or discharge device (such as a discharge resistor in the embodiment of the present application). The execution of the above-mentioned active discharge action usually involves disconnecting the connection between the high-voltage battery and the motor, and dissipating the stored electrical energy in the circuit in the form of heat energy through a discharge resistor.

[0094] For example, after confirming that an electric aircraft crashes abnormally, the system will trigger the motor controller to perform active discharge. Specifically, the system (such as a collision processor) sends a command to the motor controller, instructing it to prepare to execute the active discharge procedure; after receiving the command, the motor controller disconnects the contactor (Relay) between the high-voltage battery and cuts off the power supply; the motor controller activates the internal discharge circuit, closes the discharge switch, and forms a path between the high-voltage circuit and the discharge resistor; the electrical energy is quickly converted into heat energy through the discharge resistor, and the voltage in the high-voltage circuit drops rapidly to a safe voltage level within a few milliseconds to tens of milliseconds; after the active discharge is completed, the motor controller will report the status to the system to confirm that the high-voltage circuit has been safely discharged.

[0095] Through the technical solution provided in the above step S25, the embodiment of the present application can quickly respond and eliminate the safety threat of the high-voltage electrical system when a crash anomaly is detected by integrating and optimizing the active discharge function of the motor controller. This is of great significance for improving the safety performance of electric aircraft, protecting the lives of passengers and rescue personnel, and optimizing aircraft design and improving rescue efficiency.

[0096] Specifically, by quickly performing active discharge after a crash, the residual voltage in the high-voltage electrical system can be significantly reduced, avoiding the risk of electric shock and electric shock, and protecting drivers, passengers and potential rescuers from harm. Active discharge can quickly dissipate the energy stored in the high-voltage circuit, prevent fires caused by circuit short circuits, arcs, etc., ensure the safety of aircraft wreckage, and reduce the risk of fire and explosion after the accident. After receiving the abnormal crash signal, the motor controller can perform active discharge in a very short time, which improves the system's response speed to emergencies and ensures the immediacy of safety measures. By integrating the active discharge function, the motor controller can not only control the normal operation of the motor, but also participate in the safety protection of the high-voltage electrical system, which promotes the overall optimization of the electrical architecture of electric aircraft. Active discharge reduces the risk of high-voltage electrical systems, creates a safer rescue environment for rescuers, and improves the safety and efficiency of rescue operations.

[0097] As an optional implementation, the control method of the electric aircraft may further include the following execution steps:

[0098] Step S206, in response to the detection result that the electric aircraft has a crash anomaly, triggering the battery management system of the electric aircraft to disconnect the main positive contactor and the main auxiliary contactor corresponding to the battery module.

[0099] like Figure 3 As shown, the power supply system of the electric aircraft includes a battery module, a battery management system, a pre-charge resistor R9, and an X capacitor (such as Figure 3 C1) and multiple contactors. Figure 3A1 in the figure represents the main positive contactor, A2 represents the pre-charge contactor, and A3 represents the main negative contactor. The battery module is also connected to two CY capacitors (such as Figure 3 C2 and C3) are connected as shown.

[0100] The Battery Management System (BMS) is a component of the power supply system of an electric aircraft that monitors and manages key parameters of the high-voltage battery module, such as the health status, power level, temperature, and voltage. BMS can ensure the safe and efficient operation of the battery module under different flight conditions through complex algorithms and control logic.

[0101] The battery module is the main component of the high-voltage power supply of electric aircraft. The battery module is usually composed of multiple battery cells connected in series and / or in parallel. The battery module provides high voltage and high current for the drive motor and other high-voltage accessories.

[0102] The main positive contactor is located at the positive output end of the battery module and is responsible for connecting and disconnecting the circuit between the high-voltage battery and the electric drive system, power distribution system and other high-voltage accessories. During normal flight, the main positive contactor is in a closed state to ensure power supply; in an emergency, such as a crash, the BMS will control the main positive contactor to disconnect to cut off the high-voltage power supply.

[0103] The main and auxiliary contactors are located in the high-voltage circuit and are used to smooth the voltage when starting the high-voltage system to prevent the system from being damaged by the instantaneous high-voltage shock. In emergency situations such as crashes, the BMS will also control the main and auxiliary contactors to disconnect to further ensure the isolation of the high-voltage system.

[0104] In the application scenario, when the crash detection system determines that the electric aircraft has crashed abnormally, the battery management system will implement the following control behaviors. The BMS controls the main positive contactor of the battery module to disconnect, thereby achieving a physical disconnection between the high-voltage battery and the electric drive system, distribution system and other high-voltage accessories, quickly cutting off the power supply, preventing current from flowing in the circuit after the crash, and avoiding the risk of electric shock. The BMS controls the main and auxiliary contactors of the battery module to disconnect, further ensuring the complete isolation of the high-voltage electrical system, avoiding any potential circuit short circuit or arc discharge, and reducing the risk of fire.

[0105] Through the technical solution provided in the above step S206, the embodiment of the present application realizes the rapid isolation of the high-voltage electrical system after a crash event by integrating the main positive contactor and the main and auxiliary contactors of the BMS disconnect battery module, which is of great significance for protecting personnel safety, reducing fire risks, optimizing battery management functions, and improving the overall safety standards of electric aircraft.

[0106] Specifically, by quickly disconnecting the main positive contactor and the main auxiliary contactor, the BMS can effectively cut off the high-voltage power supply, preventing the driver, passengers and rescue personnel from electric shock injuries after the crash, and significantly improving the personnel safety of electric aircraft in emergency situations. Disconnecting the high-voltage circuit not only prevents electric shock accidents, but also reduces the possibility of fire caused by circuit short circuits or arc discharges, protects the safety of aircraft wreckage, and reduces secondary disasters after the accident. The rapid response mechanism of the BMS enables the isolation of the high-voltage electrical system to be completed in the instant of the crash, shortening the time from the crash to the complete safety of the high-voltage system, creating favorable conditions for rescue operations. The technical solution of the present application expands the function of the BMS, allowing it to participate in the emergency safety response of the aircraft while monitoring and managing the health status of the battery, enhancing the core value of the BMS in the design of electric aircraft. By integrating an efficient high-voltage electrical isolation mechanism, the safety performance of electric aircraft has been significantly improved, setting a new benchmark for safety standards in the field of electric aviation.

[0107] As an optional implementation, in the above-mentioned control method of the electric aircraft, the electric aircraft is an electric flying car, an electric airplane or an electric drone.

[0108] The above-mentioned electric flying car is a new type of transportation that combines land driving and air flying capabilities. It is usually designed as a vertical take-off and landing vehicle (eVTOL) and can provide fast and efficient travel solutions in urban environments.

[0109] The above-mentioned electric aircraft refers to an aircraft that is mainly driven by an electric motor. Electric aircraft have significant advantages in environmental protection and noise control. Electric aircraft can be designed as fixed-wing aircraft, rotorcraft or tilt-rotor aircraft, so as to be suitable for different flight missions and environments.

[0110] The above-mentioned electric drones refer to unmanned aerial vehicles driven by electric motors, which are widely used in aerial photography, logistics distribution, agricultural spraying and emergency rescue. The design of electric drones usually includes two types: multi-rotor and fixed-wing, which can perform various complex flight missions.

[0111] Applying the technical solution of the present application to different types of electric aircraft (electric flying cars, electric airplanes or electric drones) can provide a consistent high-voltage electrical system safety protection mechanism for electric aircraft of different types and uses, ensuring that in the event of a crash or abnormal situation, the high-voltage power supply of the electric aircraft can be quickly cut off and actively discharged to reduce the voltage to a safe range, thereby avoiding secondary damage.

[0112] By integrating crash detection, high-voltage electrical isolation and active discharge functions, the technical solution of the present application can improve the response speed of different types of electric aircraft when encountering emergencies, ensuring the safety of personnel and aircraft wreckage.

[0113] Through the technical solution provided in the embodiments of the present application, electric aircraft, electric airplanes and electric drones can share the same crash protection and electric shock protection architecture, which will help promote the establishment of technical standards for the electric aircraft industry and promote technical integration and sharing among different aircraft.

[0114] Through the above-mentioned technical scheme, the embodiment of the present application achieves wide applicability, unified safety protection standards and enhanced emergency response capabilities by providing a crash and electric shock protection architecture suitable for a variety of electric aircraft, thereby providing a solid technical guarantee for the safe flight of electric aircraft and the safety of passengers and rescue personnel in emergency situations.

[0115] In summary, the embodiments of the present application propose a crash protection and electric shock protection scheme for electric aircraft. The core of the scheme is to respond to crash events instantly and quickly start efficient and safe electrical isolation and discharge mechanisms. Specifically, the technical solution provided by the present application integrates high-precision crash detection sensors and intelligent control units to cut off the high-voltage power supply within a short time (such as 4ms) after a crash to avoid current flow. At the same time, by optimizing the control logic between the battery management system (BMS) and the electric drive system, the motor controller (MCU) is triggered to perform active discharge, and the voltage of the high-voltage circuit is quickly reduced to a safe level (below 60V) within 11ms, significantly shortening the time required for traditional active discharge (usually 1 to 3 seconds).

[0116] It should be noted that the user information (including but not limited to user device information, user personal information, etc.) and data (including but not limited to data used for analysis, stored data, displayed data, etc.) involved in this application are all information and data authorized by the user or fully authorized by all parties, and the collection, use and processing of relevant data must comply with the relevant laws, regulations and standards of the relevant countries and regions, and provide corresponding operation entrances for users to choose to authorize or refuse.

[0117] It should be noted that, for the above method embodiments, for the sake of simplicity, the technical solutions in the method embodiments are described as a series of action combinations, but those skilled in the art should be aware that the present application is not limited by the order of actions in the described action combinations, because according to the present application, some of the above steps can be performed in other orders or simultaneously. In addition, those skilled in the art should also be aware that the embodiments described in the specification of this application belong to preferred embodiments, and the actions and modules involved are not necessarily necessary to implement the technical solutions of the present application.

[0118] According to an embodiment of the present invention, an embodiment of an electric aircraft is also provided. Figure 4 is a structural block diagram of an electric aircraft according to an embodiment of the present application, such as Figure 4 As shown, the electric aircraft includes a power supply system, a power distribution system, an electric drive module, a power domain controller and a crash control module. The power supply system includes a battery module and a battery management unit; the power distribution system includes a DC converter, a distribution box and an on-board charging unit; the electric drive module includes a motor controller, an electric drive and a rotor; the crash control module includes an acceleration sensor, a crash processor, a disconnect switch, a closing switch and a discharge resistor; the disconnect switch is set on the high-voltage main circuit in the power supply system, the closing switch is set on the high-voltage main circuit in the distribution system, the closing switch is connected in series with the discharge resistor, and the discharge resistor is used to assist in reducing the voltage of the high-voltage main circuit to a safe voltage range; the crash processor is used to execute a computer program to implement any of the above control methods for electric aircraft.

[0119] Still like Figure 4 As shown, the power supply system is electrically connected to the power distribution system, the power supply system is electrically connected to the power domain controller, the power distribution system is electrically connected to the electric drive module, the power domain controller is electrically connected to the electric drive module, the crash control module is electrically connected to the power domain controller, the crash control module is electrically connected to the power supply system, and the crash control module is electrically connected to the power distribution system.

[0120] According to an embodiment of the present invention, an embodiment of an electronic device is also provided. The electronic device includes a memory and a processor. The memory stores a computer program. The processor is configured to run the computer program to implement the above-mentioned control method of the electric aircraft.

[0121] According to an embodiment of the present invention, a computer-readable storage medium is also provided, the computer-readable storage medium including a stored program, wherein when the program is executed, the device where the computer-readable storage medium is located is controlled to implement the above-mentioned control method of the electric aircraft.

[0122] Optionally, the above-mentioned computer storage media may include but are not limited to: hard disk drive (HDD), solid state drive (SSD), USB flash drive, optical disk, memory card, cloud storage media and network storage device (NAS).

[0123] Optionally, the computer-readable storage medium may be configured to store a computer program for executing the following steps: obtaining an acceleration signal and rotor operation data of the electric aircraft, wherein the acceleration signal is collected by a flight acceleration sensor provided on the electric aircraft, and the rotor operation data is obtained from a power domain controller of the electric aircraft through a controller local area network bus; performing crash detection analysis on the acceleration signal and the rotor operation data to determine a detection result; in response to the detection result that the electric aircraft has a crash abnormality, controlling the disconnect switch to switch to an open state within a first time range, and controlling the closing switch to switch to a closed state within a second time range, wherein the disconnect switch is arranged on a high-voltage main circuit in a power supply system of the electric aircraft, and the closing switch is arranged on a high-voltage main circuit in a power distribution system of the electric aircraft, and the closing switch is connected in series with a discharge resistor, and the discharge resistor is used to assist in reducing the voltage of the high-voltage main circuit to a safe voltage range.

[0124] According to an embodiment of the present invention, a computer program product is also provided. The computer program product includes a computer program, and when the computer program is executed by a processor, the control method of the electric aircraft can be implemented.

[0125] Optionally, the computer program product may provide control services for an electric aircraft based on the control method for an electric aircraft.

[0126] Optionally, in this embodiment, the computer program product may be a set of instructions and codes pre-written according to the control method of the electric aircraft. The computer program product may run on various computer platforms, including personal computers, servers, mobile devices, etc.

[0127] Optionally, in this embodiment, the instructions and codes corresponding to the computer program product are used to implement the following method steps: obtaining the acceleration signal and rotor operation data of the electric aircraft, wherein the acceleration signal is collected by a flight acceleration sensor provided on the electric aircraft, and the rotor operation data is obtained from the power domain controller of the electric aircraft through the controller local area network bus; performing crash detection analysis on the acceleration signal and the rotor operation data to determine the detection result; in response to the detection result that the electric aircraft has a crash abnormality, controlling the disconnect switch to switch to an open state within a first time range, and controlling the closing switch to switch to a closed state within a second time range, wherein the disconnect switch is arranged on a high-voltage main circuit in the power supply system of the electric aircraft, and the closing switch is arranged on a high-voltage main circuit in the power distribution system of the electric aircraft, and the closing switch is connected in series with a discharge resistor, and the discharge resistor is used to assist in reducing the voltage of the high-voltage main circuit to a safe voltage range.

[0128] In the above-mentioned multiple embodiments of the present application, the description of each embodiment has different emphases. For parts that are not described in detail in a certain embodiment, reference can be made to the relevant descriptions of other embodiments.

[0129] In the several embodiments provided in this application, it should be understood that the disclosed technical content can be implemented in other ways. For example, the device embodiments described above are only schematic. For example, the division of multiple modules can be a logical function division, and there can be any other possible division methods when actually implemented in the application scenario; for another example, multiple modules (or units or components in the modules) can be combined with each other and can be integrated into another system. For example, some features of the method embodiments described above can be ignored or skipped.

[0130] It should be noted that in the above-mentioned embodiments, the modules, components or units described as separate parts can be physically separated or physically integrated. The parts displayed as modules or units can be physical modules or physical units, or virtual modules or virtual units, that is, multiple modules or multiple units can be in the same position, or distributed in multiple positions or multiple spaces. In the application scenario, according to the actual needs of the scenario, some or all of the multiple modules or multiple units can be selected to implement the technical solutions of the embodiments of the present application, thereby achieving the corresponding technical objectives.

[0131] In particular, for an integrated functional module or functional unit, if it is implemented in the form of a software functional unit and sold or used as an independent product, the module or functional unit can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present application can essentially or in other words, the part that contributes to the prior art or all or part of the technical solution can be embodied in the form of a software product, which is stored in a storage medium and includes several instructions for enabling a computer device (which can be a personal computer, a server or a network device, etc.) to execute all or part of the steps of the method described in each embodiment of the present application.

[0132] The above is only a preferred implementation of the present application. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principles of the present application. These improvements and modifications should also be regarded as the scope of protection of the present application.

Claims

1. A control method for an electric aircraft, characterized in that: include: Acquiring an acceleration signal and rotor operation data of the electric aircraft, wherein the acceleration signal is collected by a flight acceleration sensor provided on the electric aircraft, and the rotor operation data is obtained from a power domain controller of the electric aircraft via a controller area network bus; Performing crash detection analysis on the acceleration signal and the rotor operation data to determine a detection result; In response to the detection result that the electric aircraft has a crash abnormality, the disconnect switch is controlled to switch to an open state within a first time range, and the closing switch is controlled to switch to a closed state within a second time range, wherein the disconnect switch is arranged on a high-voltage main circuit in the power supply system of the electric aircraft, and the closing switch is arranged on a high-voltage main circuit in the power distribution system of the electric aircraft, and the closing switch is connected in series with a discharge resistor, and the discharge resistor is used to assist in reducing the voltage of the high-voltage main circuit to a safe voltage range.

2. The control method of an electric aircraft according to claim 1, characterized in that: Performing crash detection analysis on the acceleration signal and the rotor operation data to determine the detection result includes: determining a plurality of acceleration components of the electric aircraft in a plurality of directions according to the acceleration signal; In response to the multiple acceleration components and the rotor operation data satisfying abnormal conditions, the detection result is determined as the electric aircraft having the crash anomaly, wherein the abnormal condition is determined by multiple crash thresholds corresponding to the crash anomaly.

3. The control method of an electric aircraft according to claim 2, characterized in that: The rotor operation data includes a rotor speed and a rotor torque; the multiple acceleration components and the rotor operation data satisfying the abnormal condition include at least one of the following: The absolute value of at least one acceleration component among the multiple acceleration components is greater than the acceleration crash threshold; The change in the rotor speed within a preset detection period is greater than the speed crash threshold; The change in the rotor torque within a preset detection period is greater than a torque crash threshold.

4. The control method of an electric aircraft according to claim 1, characterized in that: The first duration range is smaller than the second duration range.

5. The control method of an electric aircraft according to claim 1, characterized in that: The control method of the electric aircraft also includes: Obtaining the real-time flight altitude of the electric aircraft; In response to the detection result that the electric aircraft has the crash anomaly and the real-time flight altitude is within a preset altitude range, the cut-off switch is controlled to switch to an open state, and the closing switch is controlled to switch to a closed state.

6. The control method of an electric aircraft according to claim 1, characterized in that: The control method of the electric aircraft also includes: In response to the detection result that the electric aircraft has experienced the crash anomaly, a motor controller of the electric aircraft is triggered to perform an active discharge action.

7. The control method of an electric aircraft according to claim 1, characterized in that: The control method of the electric aircraft also includes: In response to the detection result that the electric aircraft has experienced the crash anomaly, the battery management system of the electric aircraft is triggered to disconnect the main positive contactor and the main auxiliary contactor corresponding to the battery module.

8. The control method of an electric aircraft according to claim 1, characterized in that: The electric aircraft is an electric flying car, an electric airplane or an electric drone.

9. An electric aircraft, characterized in that: include: A power supply system, a power distribution system, an electric drive module, a power domain controller and a crash control module, wherein the power supply system is electrically connected to the power distribution system, the power supply system is electrically connected to the power domain controller, the power distribution system is electrically connected to the electric drive module, the power domain controller is electrically connected to the electric drive module, the crash control module is electrically connected to the power domain controller, the crash control module is electrically connected to the power supply system, and the crash control module is electrically connected to the power distribution system; The power supply system includes a battery module and a battery management unit; The power distribution system includes a DC converter, a power distribution box and an on-board charging unit; The electric drive module includes a motor controller, an electric drive and a rotor; The crash control module includes an acceleration sensor, a crash processor, a cut-off switch, a closing switch and a discharge resistor; The disconnect switch is arranged on the high-voltage main circuit in the power supply system, the closing switch is arranged on the high-voltage main circuit in the power distribution system, the closing switch is connected in series with the discharge resistor, and the discharge resistor is used to assist in reducing the voltage of the high-voltage main circuit to a safe voltage range; The crash processor is used to execute a computer program to implement the control method of the electric aircraft as described in any one of claims 1 to 7.

10. A computer program product, characterized in that The invention comprises a computer program, which, when executed by a processor, implements the control method of an electric aircraft according to any one of claims 1 to 8.

Images