Control system of electric aircraft and electric aircraft

Abstract

The embodiment of the invention provides a control system of an electric aircraft and the electric aircraft. The control system of the electric aircraft comprises a power module, a first bus bar, a second bus bar, a first load and a second load. The first bus bar and the second bus bar are independently arranged. The power module is connected with the input end of the first bus bar and the input end of the second bus bar, and the output end of the first bus bar is connected with the first load and used for distributing electric energy of the power module to the first load. The first load is equipment of which the ripple sensitivity is smaller than the preset sensitivity in the electric aircraft; the output end of the second bus bar is connected with a second load and is used for distributing the electric energy of the power supply module to the second load; the second load is equipment of which the ripple generation amount is greater than the preset ripple generation amount in the electric aircraft. In this way, by arranging the two bus bars, partitioned power supply can be conducted on the load, and the control flexibility of the electric aircraft is improved.

Description

Technical Field

[0001] The present application relates to the technical field of electric aircraft, and in particular to a control system of an electric aircraft and an electric aircraft. Background Art

[0002] Light Sport Aircraft (LSA) is an important part of the general aviation field, and its application scenarios are becoming more and more extensive.

[0003] Currently, in the electrical system of an electric aircraft, a bus bar is provided to distribute the electrical energy of the battery of the electric aircraft to various loads.

[0004] However, different loads have different demands for electrical energy and use the same busbar. Summary of the Invention

[0005] The present application provides a control system for an electric aircraft and an electric aircraft, which provide two independent bus bars so that different bus bars can supply power to different types of loads.

[0006] In a first aspect, the present application provides a control system for an electric aircraft, comprising a power module, a first bus bar, a second bus bar, a first load, and a second load; the first bus bar and the second bus bar are independently provided;

[0007] The power module is connected to the input end of the first bus bar and the input end of the second bus bar respectively, and the output end of the first bus bar is connected to the first load, for distributing the power of the power module to the first load; the first load is a device in the electric aircraft with a ripple sensitivity less than a preset sensitivity;

[0008] The output end of the second bus is connected to the second load for distributing the electric energy of the power module to the second load; the second load is a device in the electric aircraft whose ripple generation amount is greater than a preset ripple generation amount.

[0009] In one possible implementation, the power module includes a power battery and a storage battery;

[0010] The power battery is connected to the input end of the first bus bar and the input end of the second bus bar, respectively, for supplying power to the first bus bar and the second bus bar, so that the first bus bar distributes the power energy of the power battery to the first load, and the second bus bar distributes the power energy of the power battery to the second load;

[0011] The battery is connected to the first busbar through a first switch unit. When a power battery supplying power to the first busbar fails, the first switch unit is turned on, and the battery supplies power to the first busbar, so that the first busbar distributes the battery's electrical energy to the first load.

[0012] In a possible implementation manner, the first load includes a whole machine controller, a display screen, a motor controller, and a power battery;

[0013] The output end of the first bus bar is respectively connected to the input end of the whole machine controller, the input end of the display screen, and the input end of the motor controller, and is connected to the low voltage input end of the power battery through the second switch unit;

[0014] The second load includes a standby display screen, a cooling water pump and a cooling oil pump;

[0015] The second bus bar is respectively connected to the input end of the standby display screen, the input end of the cooling water pump, and the input end of the cooling oil pump.

[0016] In a possible implementation manner, the distance between the first bus bar and the second bus bar is greater than a preset distance.

[0017] In a possible implementation, the control system of the electric aircraft further includes a first electric energy conversion module and a second electric energy conversion module;

[0018] The first power conversion module and the second power conversion module are both used to convert high voltage power into low voltage power;

[0019] The power battery includes a first power battery and a second power battery;

[0020] The output end of the first power battery is connected to the input end of the first power conversion module, and the output end of the first power conversion module is connected to the input end of the first bus bar;

[0021] The output end of the second power battery is connected to the input end of the second power conversion module, and the output end of the second power conversion module is connected to the input end of the second bus bar;

[0022] The second switch unit includes a first switch and a second switch;

[0023] The output end of the first bus bar is connected to the low-voltage input end of the first power battery through the first switch, and the output end of the second bus bar is connected to the low-voltage input end of the second power battery through the second switch.

[0024] In one possible implementation, the control system of the electric aircraft further includes a motor;

[0025] The output end of the first power battery is connected to the first input end of the motor controller, the output end of the second power battery is connected to the second input end of the motor controller, and the output end of the motor controller is connected to the motor.

[0026] In a possible implementation, the motor includes a first winding and a second winding;

[0027] The motor controller uses the electric energy of the first power battery to power the first winding, and uses the electric energy of the second power battery to power the second winding.

[0028] In one possible embodiment, the first power battery, the second power battery, the first power conversion module, the display screen, the second power conversion module, the spare display screen and the motor controller are respectively connected to the whole machine controller through the corresponding CAN interface and output interface in the whole machine controller; wherein, the first power conversion module and the display screen correspond to the same CAN interface, and the second power conversion module and the spare display screen correspond to the same CAN interface.

[0029] In a possible implementation manner, when a CAN communication corresponding to a CAN interface of a target device fails, the whole device controller controls the target device through an output interface corresponding to the target device;

[0030] The target device is at least one of the first power battery, the second power battery, the first power conversion module, the display screen, the second power conversion module, the spare display screen, and the motor controller.

[0031] In one possible implementation, when the first power battery and the second power battery simultaneously power the motor through the motor controller, and a power supply circuit of a target power battery fails, the whole machine controller controls a relay in the target power battery to turn off through an interface corresponding to the target power battery, so that power is supplied to the motor through the other power batteries.

[0032] The target power battery is one of the first power battery and the second power battery, and the other power batteries are power batteries other than the target power battery.

[0033] In a possible implementation, the whole machine controller is connected to the throttle of the electric aircraft through the first input interface and the second input interface;

[0034] The motor controller is connected to the throttle of the electric aircraft through the third input interface and the fourth input interface.

[0035] In a possible implementation, when the CAN communication between the whole machine controller and the motor controller is normal, the whole machine controller obtains a first throttle signal of the throttle through the first input interface and the second input interface, generates a first control signal according to the first throttle signal, and sends the first control signal to the motor controller through the CAN interface corresponding to the motor controller;

[0036] When the CAN communication between the whole machine controller and the motor controller is abnormal, the motor controller obtains a second throttle signal of the throttle through the third input interface and the fourth input interface, and generates a second control signal according to the second throttle signal.

[0037] In a possible implementation, when the electric aircraft is started, the first switch unit is turned on, so that the battery supplies power to the whole machine controller, the motor controller, and the display screen through the first bus bar;

[0038] The whole machine controller controls the first switch and the second switch to be turned on, so that the storage battery supplies power to the first power battery and the second power battery through the first bus bar;

[0039] When the whole machine controller detects through the CAN interfaces corresponding to the first power battery, the second power battery, and the motor controller respectively that the first power battery, the second power battery, and the motor controller are all in good condition, and detects through the first input interface and the second input interface that the throttle is in the zero position, the whole machine controller sends an enable instruction to the first power battery and the second power battery through the CAN interfaces corresponding to the first power battery and the second power battery respectively, so that the relays in the first power battery and the second power battery are turned on, and the first power battery and the second power battery supply power to the motor through the motor controller;

[0040] When the whole machine controller detects through the CAN interfaces corresponding to the first power conversion module and the second power conversion module respectively that the first power conversion module and the second power conversion module are fault-free, the whole machine controller sends an enable instruction to the first power conversion module and the second power conversion module respectively through the CAN interfaces corresponding to the first power conversion module and the second power conversion module, so that the first power battery supplies power to the first bus through the first power conversion module, and the second power battery supplies power to the second bus through the second power conversion module;

[0041] The whole machine controller sends instructions to the cooling water pump and the cooling oil pump to operate the cooling water pump and the cooling oil pump;

[0042] The whole machine controller sends system parameter information to the display screen through the interface corresponding to the display screen, and the display screen displays the system parameter information, so that the user can confirm the status of the electric aircraft based on the system parameter information displayed on the display screen.

[0043] In a second aspect, the present application provides an electric aircraft, comprising the control system of the electric aircraft as described in the first aspect and / or various possible embodiments of the first aspect.

[0044] The present application provides a control system for an electric aircraft and an electric aircraft. The control system includes: a power module, a first bus bar, a second bus bar, a first load, and a second load. The first bus bar and the second bus bar are independently provided. The power module is connected to the input end of the first bus bar and the input end of the second bus bar, respectively. The output end of the first bus bar is connected to the first load, for distributing the power from the power module to the first load; the first load is a device in the electric aircraft with a ripple sensitivity less than a preset sensitivity; the output end of the second bus bar is connected to the second load, for distributing the power from the power module to the second load; the second load is a device in the electric aircraft with a ripple generation greater than a preset ripple generation. Thus, by providing two bus bars, load power can be divided into zones, improving the flexibility of the electric aircraft control. Furthermore, since the first load has a lower ripple sensitivity and the second load has a higher ripple generation, the divided power supply can reduce the impact of the second load on the first load. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the present application and, together with the description, serve to explain the principles of the present application.

[0046] Figure 1 A schematic diagram of the direct connection relationship between a power module, a bus bar, and a load provided in an embodiment of the present application;

[0047] Figure 2 A schematic diagram of the direct connection relationship between another power module, bus bar and load provided in an embodiment of the present application;

[0048] Figure 3 A schematic structural diagram of a control system for an electric aircraft provided in an embodiment of the present application;

[0049] Figure 4 A schematic structural diagram of a control system for another electric aircraft provided in an embodiment of the present application;

[0050] Figure 5A schematic diagram of the communication structure between a whole machine controller and other devices provided in an embodiment of the present application;

[0051] Figure 6 A schematic diagram of a whole machine controller controlling a device through an interface provided in an embodiment of the present application;

[0052] Figure 7 A schematic diagram of a process for starting an electric aircraft provided in an embodiment of the present application;

[0053] Figure 8 A schematic diagram of another process flow for starting an electric aircraft provided in an embodiment of the present application.

[0054] The above drawings illustrate specific embodiments of the present application, which will be described in more detail below. These drawings and the textual description are not intended to limit the scope of the present application in any way, but rather to illustrate the concepts of the present application to those skilled in the art by reference to specific embodiments. DETAILED DESCRIPTION

[0055] Exemplary embodiments will be described in detail herein, with examples illustrated in the accompanying drawings. In the following description, when referring to the drawings, identical numerals in different figures represent identical or similar elements, unless otherwise indicated. The embodiments described in the following exemplary embodiments are not intended to represent all embodiments consistent with the present application. Rather, they are merely examples of apparatus and methods consistent with certain aspects of the present application, as detailed in the appended claims.

[0056] Light sport aircraft (LSA) are small aircraft with a maximum takeoff weight of no more than 650 kg. They are an important component of general aviation. They are primarily used for basic flight training, air sports events, private recreational flights, and low-altitude sightseeing. Compared to engine-powered LSAs, electric LSAs offer lower noise, reduced carbon emissions, and lower maintenance costs.

[0057] Currently, in the electrical system of an electric aircraft, a bus bar is provided to distribute the electrical energy of the battery of the electric aircraft to various loads.

[0058] However, distributing electrical energy to various loads through a bus bar is a relatively simple power supply method, which makes the control flexibility of the electrical system of the electric aircraft poor.

[0059] Based on this, the present application provides a control system for an electric aircraft. Two independent busbars are provided in the control system of the electric aircraft. One busbar distributes the power of the power module to loads with low ripple sensitivity, and the other busbar distributes the power of the power module to loads with high ripple generation. In this way, powering different loads through two independent busbars reduces the mutual influence between the two loads and improves the control flexibility of the control system.

[0060] The following specific embodiments describe in detail the technical solution of the present application and how the technical solution of the present application solves the above-mentioned technical problems. The following specific embodiments can be combined with each other, and the same or similar concepts or processes may not be repeated in some embodiments. The embodiments of the present application will be described below in conjunction with the accompanying drawings.

[0061] Figure 1 A schematic diagram of the direct connection relationship between a power module, a bus bar and a load provided in an embodiment of the present application.

[0062] like Figure 1 As shown, the control system of the electric aircraft includes: a power module, a first bus bar, a second bus bar, a first load, and a second load; the first bus bar and the second bus bar are independently arranged.

[0063] The power supply module is connected to the input end of the first bus bar and the input end of the second bus bar respectively, the output end of the first bus bar is connected to the first load, and the output end of the second bus bar is connected to the second load.

[0064] The first bus bar is used to distribute the electric energy of the power module to the first load.

[0065] The second bus bar is used to distribute the electric energy of the power module to the second load.

[0066] The first load is a device in the electric aircraft whose ripple sensitivity is less than a preset sensitivity, and the second load is a device in the electric aircraft whose ripple generation amount is greater than a preset ripple generation amount.

[0067] The preset sensitivity can be set according to the actual situation of each load. For example, the preset sensitivity can be 50mV, which is not limited in the embodiment of the present application.

[0068] The preset ripple generation amount can be determined based on the importance of each load in the electric aircraft to the safe operation of the electric aircraft and the ripple that each load can withstand. For example, the preset ripple generation amount can be 100mV, which is not limited in this embodiment of the present application.

[0069] During the operation of the electric aircraft, the first bus bar can distribute the electric energy of the power module to the first load, and the second bus bar can distribute the electric energy of the power module to the second load.

[0070] Thus, by providing two independent busbars in the electric aircraft's control system, one busbar distributes the power from the power module to loads with low ripple sensitivity, while the other busbar distributes the power from the power module to loads with high ripple generation. This allows power to be supplied to different loads through two independent busbars, and the loads are differentiated based on their electromagnetic interference characteristics, minimizing the mutual influence between the two loads and improving the control flexibility of the control system.

[0071] In the present application, the distance between the first bus bar and the second bus bar is greater than a preset distance.

[0072] Exemplarily, the preset distance may be determined based on the ripple sensitivity and ripple generation capacity of each load. For example, the preset distance may be 50 mm, which is not limited in the embodiment of the present application.

[0073] For example, inductive loads such as the cooling water pump and cooling oil pump in an electric aircraft generate high-frequency switching ripple when in operation. This high-frequency switching ripple can interfere with certain loads in the control system. For example, it can interfere with the motor controller in the electric aircraft, affecting the accuracy of the motor controller's data acquisition.

[0074] In this way, by physically isolating the first bus bar from the second bus bar, the spatial coupling interference between the loads can be reduced, the accuracy of the load operation process can be improved, and the safety of the electric aircraft can be improved.

[0075] In this application, the power module includes a power battery and a storage battery.

[0076] The power battery is connected to the input end of the first bus bar and the input end of the second bus bar respectively, and the storage battery is connected to the first bus bar through the first switch unit.

[0077] The power battery is used to supply power to the first bus bar and the second bus bar, so that the first bus bar distributes the power battery's power to the first load, and the second bus bar distributes the power battery's power to the second load.

[0078] Illustratively, during aircraft operation, when a power battery supplying power to the first bus fails, the first switch unit is turned on, and the battery supplies power to the first bus, so that the first bus distributes the battery's electrical energy to the first load.

[0079] In this way, the first busbar is equipped with both a power battery and a storage battery for power supply. Under normal circumstances, the power battery powers the first load. If the first battery fails, the storage battery provides power to the first load, thus ensuring uninterrupted power supply to the first load. Since the first load is a critical device with low ripple sensitivity, providing uninterrupted power to this critical device allows the electric aircraft to continue operating for a period of time even if a power battery fails, thereby improving the safety of the electric aircraft. The second, less critical load does not require uninterrupted operation and does not require a storage battery, thus saving energy and reducing the weight of the electric aircraft.

[0080] In this application, the control system of the electric aircraft further includes a first power conversion module and a second power conversion module. The first load may include an aircraft controller, a display screen, a motor controller, and a power battery. The second load may include a backup display screen, a cooling water pump, and a cooling oil pump. The power battery includes a first power battery and a second power battery.

[0081] Figure 2 A schematic diagram of the direct connection relationship between another power module, bus bar and load provided in an embodiment of the present application.

[0082] Figure 2 The direction of the arrow indicates the direction of flow of electrical energy.

[0083] like Figure 2 As shown, the output end of the first bus bar is respectively connected to the input end of the whole machine controller, the input end of the display screen, and the input end of the motor controller, and is connected to the low voltage input end of the power battery through the second switch unit.

[0084] The second bus bar is respectively connected to the input end of the standby display screen, the input end of the cooling water pump, and the input end of the cooling oil pump.

[0085] The output end of the first power battery is connected to the input end of the first power conversion module, and the output end of the first power conversion module is connected to the input end of the first bus bar;

[0086] The output end of the second power battery is connected to the input end of the second power conversion module, and the output end of the second power conversion module is connected to the input end of the second bus bar.

[0087] The second switch unit includes a first switch and a second switch.

[0088] like Figure 2 As shown, the output end of the first bus bar is connected to the low voltage input end of the first power battery through the first switch, and the output end of the second bus bar is connected to the low voltage input end of the second power battery through the second switch.

[0089] The first power conversion module and the second power conversion module are both used to convert high-voltage power into low-voltage power.

[0090] Exemplarily, the first power conversion module and the second power conversion module may be DC / DC, or other devices capable of converting high voltage into low voltage, which is not limited in the embodiments of the present application.

[0091] The first power battery and the second power battery are both high-voltage power sources, and the storage battery is a low-voltage power source. For example, the storage battery may be a low-voltage power source with an output voltage of 14V.

[0092] Illustratively, during the operation of the aircraft, the first power battery transmits high-voltage power to the first power conversion module, the first power conversion module converts the high-voltage power into low-voltage power and transmits it to the first bus, and the first bus distributes power to the whole machine controller, display screen, motor controller, first power battery and second power battery according to their power requirements, so that the whole machine controller, display screen and motor controller can operate, and the first power battery and the second power battery can be enabled and controlled.

[0093] For the second power battery, during the operation of the aircraft, the second power battery transmits high-voltage power to the second power conversion module, and the second power conversion module converts the high-voltage power into low-voltage power and transmits it to the second bus. The second bus distributes power to the backup display screen, cooling water pump and cooling oil pump according to their power requirements, so that the backup display screen, cooling water pump and cooling oil pump can operate.

[0094] In this way, the first bus distributes power to critical loads such as the aircraft controller, display, motor controller, and power battery, ensuring uninterrupted power supply to these loads. This allows the aircraft controller, display, and motor controller to continue operating even if the first power battery fails, and the aircraft controller can still control the second power battery. Furthermore, the second bus supplies power to loads with less impact on the electric aircraft's operation, such as the backup display, cooling water pump, and cooling oil pump, thus enabling the division of power supply circuits according to load conditions.

[0095] In the present application, the control system of the electric aircraft further includes a motor, and the motor includes a first winding and a second winding.

[0096] In combination with the above embodiments, Figure 3 A schematic structural diagram of a control system for an electric aircraft provided in an embodiment of the present application.

[0097] like Figure 3As shown, the output of the first power battery is connected to the first input of the motor controller, the output of the second power battery is connected to the second input of the motor controller, and the output of the motor controller is connected to the motor. The motor controller uses the power of the first power battery to power the first winding and the power of the second power battery to power the second winding.

[0098] For example, based on Figure 3 As can be seen from the control system shown, the total output power of the motor is the sum of the power input from the first power battery and the power input from the second power battery.

[0099] Combine Figure 3 As shown, during the operation of the electric aircraft, the first power battery and the second power battery both power the motor through the electrode controller, and the first power battery powers the first winding of the motor, and the second power battery powers the second winding of the motor.

[0100] For example, when a fault is detected in the first power battery, the motor controller bus channel, or the first winding, the relay in the first power battery can be controlled to shut down, and electrical isolation can be performed through PWM wave sealing action to reduce the impact of the fault in the power supply circuit of the first power battery on other devices.

[0101] In this way, the motor is powered by both the first and second power batteries, allowing the high-voltage electrical system in the control system to include redundant first and second power batteries. These two battery packs independently power the motor, ensuring that if one power battery fails, the other can still power the motor, ensuring continued operation and improving aircraft safety. Furthermore, the motor's dual windings effectively increase its power output.

[0102] In combination with the above embodiment, taking the first power conversion module as DC / DC1, the second power conversion module as DC / DC2, and the first switch unit including a relay and a third switch as an example, the control system of the electric aircraft is described in detail. Figure 4 A schematic structural diagram of a control system for another electric aircraft provided in an embodiment of the present application.

[0103] like Figure 4 As shown, the positive output terminal (HV+) of the first power battery is connected to the first positive input terminal (HV1+) of the motor controller and the positive input terminal (HV+) of DC / DC1 respectively, and the negative output terminal (HV-) of the first power battery is connected to the first negative input terminal (HV1-) of the motor controller and the negative input terminal (HV-) of DC / DC1 respectively.

[0104] The positive output terminal (HV+) of the second power battery is respectively connected to the second positive input terminal (HV2+) of the motor controller and the positive input terminal (HV+) of DC / DC2, and the negative output terminal (HV-) of the second power battery is respectively connected to the second negative input terminal (HV2-) of the motor controller and the negative input terminal (HV-) of DC / DC2.

[0105] The positive output terminal (Out+) of DC / DC1 is connected to the first positive input terminal of the first busbar through a fuse, and the negative output terminal (Out-) of DC / DC1 is connected to the first negative input terminal of the first busbar.

[0106] The first switch unit includes a relay and a third switch S3. The battery's positive output terminal (Out+) is connected to the second positive input terminal of the first busbar through a fuse and a relay, and is also connected to the third negative input terminal of the first busbar through the relay and the third switch S3. The battery's negative output terminal (Out-) is connected to the second negative input terminal of the first busbar.

[0107] The first positive output terminal of the first bus is connected to the positive terminal (12V+) of the whole machine controller through a fuse, and the first negative output terminal of the first bus is connected to the ground (GND) of the whole machine controller. The second positive output terminal of the first bus is connected to the positive terminal (12V+) of the motor controller through a fuse, and the second negative output terminal of the first bus is connected to the ground (GND) of the motor controller. The third positive output terminal of the first bus is connected to the positive terminal (12V+) of the first display screen through a fuse, and the third negative output terminal of the first bus is connected to the ground (GND) of the first display screen. The fourth positive output terminal of the first bus is connected to the positive input terminal (12V+) of the first power battery through a fuse and the first switch S1, and the fourth negative output terminal of the first bus is connected to the ground (GND) input terminal of the first power battery. The fifth positive output terminal of the first busbar is connected to the positive input terminal (12V+) of the second power battery through the fuse and the second switch S2, and the fifth negative output terminal of the first busbar is connected to the ground (GND) input terminal of the second power battery.

[0108] The positive output terminal (Out+) of DC / DC2 is connected to the positive input terminal of the second busbar through a fuse, and the negative output terminal (Out-) of DC / DC2 is connected to the negative input terminal of the second busbar.

[0109] The first positive output terminal of the second bus bar is connected to the positive terminal (12V+) of the second display screen through a fuse, and the first negative output terminal of the second bus bar is connected to the ground (GND) of the second display screen. The second positive output terminal of the second bus bar is connected to the positive terminal (12V+) of the cooling water pump through a fuse, and the second negative output terminal of the second bus bar is connected to the ground (GND) of the cooling water pump. The third positive output terminal of the second bus bar is connected to the positive terminal (12V+) of the cooling oil pump through a fuse, and the third negative output terminal of the second bus bar is connected to the ground (GND) of the cooling oil pump.

[0110] based on Figure 4 As shown, the first bus bar and the battery form a dual-channel redundant topology.

[0111] Under normal operating conditions, DCDC1 converts the high-voltage DC (for example, 400V) of the first power battery into 14V DC, which is used to power the first bus and charge the battery simultaneously.

[0112] Each load branch corresponding to the first bus bar is equipped with a fuse to achieve overcurrent protection.

[0113] The fault tolerance mechanism includes: when DCDC1 or the first power battery failure is detected, the relay in the first power battery is controlled to cut off the fault link, and the battery continues to power the first bus, ensuring that the critical load can continue to operate for a period of time, for example, at least 30 minutes, to meet the requirements of emergency landing conditions.

[0114] The above embodiment describes the circuit structure of the control system of the electric aircraft. Next, the communication structure in the control system will be described.

[0115] Figure 5 A schematic diagram of the communication structure between a whole machine controller and other devices provided in an embodiment of the present application.

[0116] like Figure 5 As shown, the first power battery, the second power battery, the first power conversion module, the display screen, the second power conversion module, the spare display screen and the motor controller are connected to the whole machine controller through the corresponding CAN interface and output interface in the whole machine controller respectively.

[0117] The first power conversion module and the display screen correspond to the same CAN interface, and the second power conversion module and the spare display screen correspond to the same CAN interface.

[0118] Specifically, the whole machine controller is connected to the first power battery through the first CAN interface (CAN1H and CAN1L) and the first output interface (Out1). The whole machine controller is connected to the second power battery through the second CAN interface (CAN2H and CAN2L) and the second output interface (Out2).

[0119] The whole machine controller is connected to the first power conversion module and the display screen respectively through the third CAN interface (CAN3H and CAN3L), and is connected to the first power conversion module through the third output interface (Out3).

[0120] The whole machine controller is connected to the second power conversion module and the standby display screen respectively through the fourth CAN interface (CAN4H and CAN4L), and is connected to the second power conversion module through the fourth output interface (Out4).

[0121] The whole machine controller is connected to the motor controller through the fifth CAN interface (CAN5H and CAN5L) and the fifth output interface (Out5).

[0122] The whole machine controller is connected to the cooling oil pump through the sixth CAN interface (CAN6H and CAN6L).

[0123] The whole machine controller is connected to the cooling water pump through a pulse width modulation (PWM) interface.

[0124] The whole machine controller is connected to the throttle of the electric aircraft through the first input interface (In1) and the second input interface (In2);

[0125] The motor controller is connected to the throttle of the electric aircraft through the third input interface (In3) and the fourth input interface (In4).

[0126] It should be noted that the above-mentioned output interfaces may be DO interfaces, and the input interfaces may be AI interfaces.

[0127] based on Figure 5 As shown in Table 1 below: Table 1 Correspondence between the interfaces, devices and control / signal types of the whole machine controller

[0128]

[0129] This application proposes a hierarchical redundant architecture that employs dual-channel control for key equipment. This means that the entire machine controller can connect to key equipment such as the motor controller and power battery via two interfaces. Furthermore, the power control link is configured with dual backup for signal acquisition and command transmission. This allows the output interface to be used to control the equipment in the event of a CAN communication failure.

[0130] Combined with the above Figure 5 As shown in the figure, the control strategy of the whole machine controller is explained when a power supply circuit of the power battery fails.

[0131] For example, when the first power battery and the second power battery simultaneously power the motor through the motor controller, and the power supply circuit of the target power battery fails, the whole machine controller controls the relay in the target power battery to shut down through the interface corresponding to the target power battery, so that the motor can be powered by other power batteries.

[0132] The target power battery is one of the first power battery and the second power battery, and the other power batteries are power batteries other than the target power battery.

[0133] That is, when a fault is detected in the power supply circuit of the first power battery, the whole machine controller sends a control signal to the first power battery through the first CAN interface to control the relay in the first power battery to be turned off.

[0134] The situation where the power supply circuit of the first power battery fails may include: the first power battery, the bus channel corresponding to the first power battery in the motor controller, or the first winding fails.

[0135] In this way, if the power supply circuit of one power battery fails, the other power battery can still supply power to the motor, allowing the motor to operate and improving the safety of the electric aircraft.

[0136] In the present application, when a CAN communication corresponding to the CAN interface of the target device fails, the whole machine controller controls the target device through the output interface corresponding to the target device.

[0137] The target device is at least one of the first power battery, the second power battery, the first power conversion module, the display screen, the second power conversion module, the spare display screen and the motor controller.

[0138] Exemplarily, the detection of whether a CAN communication failure occurs may include: performing CRC check on the data frame during the data communication process, updating the data frame technology through a rolling counter after the CRC check, counting the sending status of the data frame of each device through a sending error counter, or counting the receiving status of the data frame through a receiving error counter.

[0139] The send error counter is incremented when a send error occurs and decremented when a send is successful. The receive error counter is incremented when a receive error occurs and decremented when a receive is successful.

[0140] When the whole machine controller detects that the count TEC of the sending error counter is greater than the first value or the count TEC of the receiving error counter REC is greater than the second value, it determines that the CAN communication of the corresponding device has failed, and can control the device through the corresponding digital output interface Out.

[0141] For example, after a CAN communication failure is detected, a timer may be used. If the failure disappears after a preset time, the CAN interface may continue to be used for communication.

[0142] In this application, the process of the whole machine controller controlling the device through the interface can be found in Figure 6 As shown, Figure 6 A schematic diagram of a whole machine controller controlling a device through an interface provided in an embodiment of the present application.

[0143] like Figure 6 As shown, the whole machine controller determines whether CAN communication is valid through the enable signal. If valid, CAN communication is used to execute CAN instructions. If invalid, DO communication is used to execute DO instructions.

[0144] It should be understood that when the device detects a valid CAN enable frame, it ignores the digital input state; only when the CAN bus enters the Busoff state does it switch to DO signal control.

[0145] In this way, the whole machine controller implements dual-mode enabling control of the power battery, power conversion module and motor controller. When the CAN bus is valid, the bus instructions are executed first, and when a bus fault is detected, it automatically switches to the digital control channel.

[0146] Combined with the above Figure 5 As shown, the process of generating a control signal according to the throttle signal of the electric aircraft may include:

[0147] When the CAN communication between the whole machine controller and the motor controller is normal, the whole machine controller obtains the first throttle signal of the throttle through the first input interface and the second input interface, generates a first control signal according to the first throttle signal, and sends the first control signal to the motor controller through the CAN interface corresponding to the motor controller.

[0148] That is, under normal operating conditions, the whole machine controller collects the throttle signal through AI1-AI2, converts it into the motor speed command through the calibration curve, and transmits it to the motor controller through the fifth CAN interface to control the motor.

[0149] When the CAN communication between the whole machine controller and the motor controller is abnormal, the motor controller obtains the second throttle signal of the throttle through the third input interface and the fourth input interface, and generates a second control signal according to the second throttle signal.

[0150] For example, when the Busoff of the fifth CAN interface lasts for more than 200ms, the motor controller enables the local spare AD channel to directly read the throttle signal and independently completes the control conversion to control the motor.

[0151] For example, the throttle signal may be a signal indicating the degree of opening and closing of the accelerator pedal. The control signal may be a speed command of the motor.

[0152] In this way, when the communication interface between the whole machine controller and the throttle fails, the motor controller can obtain the throttle signal through the backup interface and generate a control signal, which can reduce the situation where the control signal error of the electric aircraft is caused by the failure of the interface corresponding to the throttle, thereby improving the safety of the electric aircraft.

[0153] In this application, the process of starting the electric aircraft can be found in Figure 7 As shown, Figure 7 A schematic diagram of a process flow for starting an electric aircraft provided in an embodiment of the present application.

[0154] like Figure 7 As shown, the startup process of an electric aircraft may include:

[0155] S701: Initial power supply.

[0156] When the electric aircraft is started, the first switch unit is turned on, so that the battery supplies power to the whole machine controller, the motor controller and the display screen through the first bus bar.

[0157] Exemplarily, the first switch unit may be turned on when a start instruction is received.

[0158] The first switch unit is turned on, including the third switch and the relay. The battery supplies power to the whole machine controller, the motor controller, and the display screen, forming a minimum control system.

[0159] S702: Pre-check and low-voltage wake-up.

[0160] The whole machine controller controls the first switch and the second switch to be turned on, so that the storage battery supplies power to the first power battery and the second power battery through the first bus bar.

[0161] The whole machine controller detects whether there are faults in the first power battery, second power battery, and motor controller through their corresponding CAN interfaces. Specifically, the first and second CAN interfaces are used to monitor whether the power battery has no fault codes, and the fifth CAN interface is used to verify whether the motor controller has no fault codes.

[0162] And check whether the throttle analog input is at zero position.

[0163] S703, high voltage power on.

[0164] When the whole machine controller detects through the CAN interfaces corresponding to the first power battery, the second power battery and the motor controller that there are no faults in the first power battery, the second power battery and the motor controller, that is, there is no fault code, it can be determined that the pre-check has passed. If there is a fault, it can be determined that the pre-check has failed, the process is interrupted, and the start of the electric aircraft is interrupted.

[0165] After the pre-check passes, when it is detected through the first input interface and the second input interface that the throttle is in the zero position, the whole machine controller sends an enable instruction to the first power battery and the second power battery through the CAN interfaces corresponding to the first power battery and the second power battery respectively, so that the relays in the first power battery and the second power battery are both turned on, and the first power battery and the second power battery power the motor through the motor controller.

[0166] Among them, after the relay controlling the power battery is turned on, the whole machine controller can monitor the bus voltage climbing rate of the power battery in real time to monitor the discharge condition of the power battery.

[0167] S704, low voltage system switching.

[0168] The whole machine controller detects the states of the first power conversion module and the second power conversion module through the third CAN interface and the fourth CAN interface.

[0169] When the whole machine controller detects that the first power conversion module (i.e., DC / DC1) and the second power conversion module (i.e., DC / DC2) are fault-free through the CAN interfaces corresponding to the first power conversion module (i.e., DC / DC1) and the second power conversion module (i.e., DC / DC2), the whole machine controller sends an enable instruction to the first power conversion module and the second power conversion module through the CAN interfaces corresponding to the first power conversion module and the second power conversion module, so that the first power battery supplies power to the first bus through the first power conversion module, and the second power battery supplies power to the second bus through the second power conversion module.

[0170] S705: The cooling system starts.

[0171] The whole machine controller sends instructions to the cooling water pump and the cooling oil pump to make the cooling water pump and the cooling oil pump run.

[0172] S706: Human-computer interaction confirmation.

[0173] The whole machine controller sends system parameter information to the display screen through the interface corresponding to the display screen, and the display screen displays the system parameter information, so that the user can confirm the status of the electric aircraft based on the system parameter information displayed on the display screen.

[0174] When the user confirms that all status flags are correct, the throttle enable flag is set.

[0175] S707, power output.

[0176] When the throttle opening is greater than a certain value (for example, 5%), the whole machine controller sends a torque command to the motor controller through CAN5, and the motor is driven.

[0177] Combined with the above steps S701-S707, the specific process of starting the electric aircraft can be found in Figure 8 As shown, Figure 8 A schematic diagram of another process flow for starting an electric aircraft provided in an embodiment of the present application.

[0178] In this way, the electric aircraft is started by firstly applying low voltage and then applying high voltage, which makes the starting process of the electric aircraft safer.

[0179] The present application also provides an electric aircraft, which includes any one of the control systems described in the above embodiments.

[0180] The present application also provides a computer program product, including a computer program, which implements the above method when executed by the whole machine controller.

[0181] The present application also provides a computer-readable storage medium, in which computer-executable instructions are stored. When a processor executes the computer-executable instructions, the above method is implemented.

[0182] The readable storage medium may be implemented by any type of volatile or non-volatile memory device, or a combination thereof, such as static random access memory (SRAM), electrically erasable programmable read-only memory (EEPROM), erasable programmable read-only memory (EPROM), programmable read-only memory (PROM), read-only memory (ROM), magnetic storage, flash memory, magnetic disk, or optical disk. The readable storage medium may be any available medium that can be accessed by a general-purpose or special-purpose computer.

[0183] An exemplary readable storage medium is coupled to a processor so that the processor can read information from the readable storage medium and write information to the readable storage medium. Of course, the readable storage medium can also be an integral part of the processor. The processor and the readable storage medium can be located in an application specific integrated circuit (ASIC). Of course, the processor and the readable storage medium can also exist in the device as discrete components.

[0184] The division of units is merely a logical functional division; actual implementations may employ alternative divisions, such as combining or integrating multiple units or components into another system, or omitting or disabling certain features. Furthermore, any coupling or direct coupling or communication connection shown or discussed may be an indirect coupling or communication connection between devices or units, whether electrical, mechanical, or otherwise, through some interface.

[0185] Units described as separate components may or may not be physically separate, and components shown as units may or may not be physical units, that is, they may be located in one place or distributed across multiple network units. Some or all of these units may be selected to achieve the purpose of this embodiment according to actual needs.

[0186] In addition, each functional unit in each embodiment of the present invention may be integrated into one processing unit, or each unit may exist physically separately, or two or more units may be integrated into one unit.

[0187] If a function is implemented as a software functional unit and sold or used as an independent product, it can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present invention, or the portion that contributes to the prior art, or a portion of the technical solution, can be embodied in the form of a software product. This computer software product is stored in a storage medium and includes several instructions for causing a computer device (which can be a personal computer, server, or network device, etc.) to execute all or part of the steps of the various embodiments of the method of the present invention. The aforementioned storage medium includes various media that can store program code, such as USB flash drives, mobile hard drives, read-only memories (ROMs), random access memories (RAMs), magnetic disks, or optical disks.

[0188] Those skilled in the art will appreciate that all or part of the steps in the above-described method embodiments can be implemented using hardware associated with program instructions. The aforementioned program can be stored in a computer-readable storage medium. When executed, the program performs the steps of the above-described method embodiments. The aforementioned storage medium includes various media capable of storing program code, such as ROM, RAM, magnetic disks, or optical disks.

[0189] Finally, it should be noted that those skilled in the art will readily identify other embodiments of the present invention after considering the specification and practicing the invention disclosed herein. The present invention is intended to cover any variations, uses, or adaptations of the present invention that follow the general principles of the present invention and include common knowledge or customary techniques in the art not disclosed herein. The present invention is not limited to the precise structure described above and illustrated in the accompanying drawings, and various modifications and variations may be made without departing from the scope thereof. The scope of the present invention is limited solely by the appended claims.

Claims

1. A control system for an electric aircraft, characterized in that: include: A power module, a first bus bar, a second bus bar, a first load, and a second load; the first bus bar and the second bus bar are independently provided; The power module is connected to the input end of the first bus bar and the input end of the second bus bar respectively, and the output end of the first bus bar is connected to the first load, for distributing the power of the power module to the first load; the first load is a device in the electric aircraft with a ripple sensitivity less than a preset sensitivity; The output end of the second bus is connected to the second load for distributing the electric energy of the power module to the second load; the second load is a device in the electric aircraft whose ripple generation amount is greater than a preset ripple generation amount.

2. The control system of the electric aircraft according to claim 1, characterized in that: The power module includes a power battery and a storage battery; The power battery is connected to the input end of the first bus bar and the input end of the second bus bar, respectively, for supplying power to the first bus bar and the second bus bar, so that the first bus bar distributes the power energy of the power battery to the first load, and the second bus bar distributes the power energy of the power battery to the second load; The battery is connected to the first busbar through a first switch unit. When a power battery supplying power to the first busbar fails, the first switch unit is turned on, and the battery supplies power to the first busbar, so that the first busbar distributes the battery's electrical energy to the first load.

3. The control system of the electric aircraft according to claim 2, characterized in that: The first load includes a whole machine controller, a display screen, a motor controller, and a power battery; The output end of the first bus bar is respectively connected to the input end of the whole machine controller, the input end of the display screen, and the input end of the motor controller, and is connected to the low voltage input end of the power battery through the second switch unit; The second load includes a standby display screen, a cooling water pump and a cooling oil pump; The second bus bar is respectively connected to the input end of the standby display screen, the input end of the cooling water pump, and the input end of the cooling oil pump.

4. The control system of the electric aircraft according to claim 3, characterized in that: A distance between the first bus bar and the second bus bar is greater than a preset distance.

5. The control system of the electric aircraft according to claim 3, characterized in that: The control system of the electric aircraft further includes a first electric energy conversion module and a second electric energy conversion module; The first power conversion module and the second power conversion module are both used to convert high voltage power into low voltage power; The power battery includes a first power battery and a second power battery; The output end of the first power battery is connected to the input end of the first power conversion module, and the output end of the first power conversion module is connected to the input end of the first bus bar; The output end of the second power battery is connected to the input end of the second power conversion module, and the output end of the second power conversion module is connected to the input end of the second bus bar; The second switch unit includes a first switch and a second switch; The output end of the first bus bar is connected to the low-voltage input end of the first power battery through the first switch, and the output end of the second bus bar is connected to the low-voltage input end of the second power battery through the second switch.

6. The control system of the electric aircraft according to claim 5, characterized in that: The control system of the electric aircraft further comprises a motor; The output end of the first power battery is connected to the first input end of the motor controller, the output end of the second power battery is connected to the second input end of the motor controller, and the output end of the motor controller is connected to the motor.

7. The control system of the electric aircraft according to claim 6, characterized in that: The motor includes a first winding and a second winding; The motor controller uses the electric energy of the first power battery to power the first winding, and uses the electric energy of the second power battery to power the second winding.

8. The control system of an electric aircraft according to any one of claims 1 to 7, characterized in that: The first power battery, the second power battery, the first power conversion module, the display screen, the second power conversion module, the spare display screen and the motor controller are respectively connected to the whole machine controller through the corresponding CAN interface and output interface in the whole machine controller; wherein, the first power conversion module and the display screen correspond to the same CAN interface, and the second power conversion module and the spare display screen correspond to the same CAN interface.

9. The control system of the electric aircraft according to claim 8, characterized in that: When a CAN communication corresponding to the CAN interface of the target device fails, the whole device controller controls the target device through the output interface corresponding to the target device; The target device is at least one of the first power battery, the second power battery, the first power conversion module, the display screen, the second power conversion module, the spare display screen, and the motor controller.

10. The control system of the electric aircraft according to claim 9, characterized in that: When the first power battery and the second power battery simultaneously power the motor through the motor controller, and a power supply circuit of the target power battery fails, the whole machine controller controls the relay in the target power battery to turn off through the interface corresponding to the target power battery, so that the motor is powered by the other power batteries; The target power battery is one of the first power battery and the second power battery, and the other power batteries are power batteries other than the target power battery.

11. The control system of the electric aircraft according to claim 8, characterized in that: The whole machine controller is connected to the throttle of the electric aircraft through the first input interface and the second input interface; The motor controller is connected to the throttle of the electric aircraft through the third input interface and the fourth input interface.

12. The control system of the electric aircraft according to claim 11, characterized in that: When the CAN communication between the whole machine controller and the motor controller is normal, the whole machine controller obtains a first throttle signal of the throttle through the first input interface and the second input interface, generates a first control signal according to the first throttle signal, and sends the first control signal to the motor controller through the CAN interface corresponding to the motor controller; When the CAN communication between the whole machine controller and the motor controller is abnormal, the motor controller obtains a second throttle signal of the throttle through the third input interface and the fourth input interface, and generates a second control signal according to the second throttle signal.

13. The control system of the electric aircraft according to claim 11, characterized in that: When the electric aircraft is started, the first switch unit is turned on, so that the battery supplies power to the whole machine controller, the motor controller and the display screen through the first bus bar; The whole machine controller controls the first switch and the second switch to be turned on, so that the storage battery supplies power to the first power battery and the second power battery through the first bus bar; When the whole machine controller detects through the CAN interfaces corresponding to the first power battery, the second power battery, and the motor controller respectively that the first power battery, the second power battery, and the motor controller are all in good condition, and detects through the first input interface and the second input interface that the throttle is in the zero position, the whole machine controller sends an enable instruction to the first power battery and the second power battery through the CAN interfaces corresponding to the first power battery and the second power battery respectively, so that the relays in the first power battery and the second power battery are turned on, and the first power battery and the second power battery supply power to the motor through the motor controller; When the whole machine controller detects through the CAN interfaces corresponding to the first power conversion module and the second power conversion module respectively that the first power conversion module and the second power conversion module are fault-free, the whole machine controller sends an enable instruction to the first power conversion module and the second power conversion module respectively through the CAN interfaces corresponding to the first power conversion module and the second power conversion module, so that the first power battery supplies power to the first bus through the first power conversion module, and the second power battery supplies power to the second bus through the second power conversion module; The whole machine controller sends instructions to the cooling water pump and the cooling oil pump to operate the cooling water pump and the cooling oil pump; The whole machine controller sends system parameter information to the display screen through the interface corresponding to the display screen, and the display screen displays the system parameter information, so that the user can confirm the status of the electric aircraft based on the system parameter information displayed on the display screen.

14. An electric aircraft, characterized in that: A control system for an electric aircraft comprising the control system of any one of claims 1-13.

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