Electric aircraft power load controller system and working method thereof

Abstract

The invention discloses an electric aircraft power load controller system and a working method thereof, and belongs to the technical field of electric aircraft controllers, and the electric aircraft power load controller system comprises a main power module, a pre-charging module, a discharging module, a monitoring control module, a power supply module, a communication module and a terminal. Through the mode, the solid-state power switch and the high-power mechanical switch are connected in parallel, and the control time sequence of the two switches is set, so that the high-voltage high-power electric load circuit can be suitable for the application scene of a high-voltage high-power electric load of an electric aircraft, and has good capacitive load adaptability and electric arc protection performance; and controllable pre-charging of the equivalent capacitance of the high-voltage line is realized by controlling the semiconductor power switch by utilizing the impedance characteristic of the semiconductor power switch in the saturation region through the pre-charging module, so that the complexity and the weight of equipment are reduced.

Description

Technical Field

[0001] The present invention relates to the technical field of electric aircraft controllers, and in particular to an electric aircraft power load controller system and a working method thereof. Background Art

[0002] Green development is an inevitable requirement for achieving sustainable development of the aviation industry and is the core key to enhancing the future competitiveness of the aviation manufacturing industry. Electric aircraft are important carriers and core products for the development of green aviation and low-altitude economy. The electric propulsion system is an innovative system different from traditional aircraft. The electric propulsion system of electric aircraft has the characteristics of high power and high reliability requirements. It usually uses a high-voltage DC grid to realize energy transmission and control from power batteries to propulsion devices. Due to the electrical characteristics of the high-voltage DC system and electric drive, the high-voltage DC bus is capacitive. During the power-on process of the system and equipment, the system should first slowly charge the equivalent capacitor to prevent the impact of instantaneous large current on the system; during the power-off process of the system and equipment, the system should first release the electric energy stored in the equivalent capacitor to ensure the safety of the system and personnel. At the same time, the high-voltage DC grid also needs to pay attention to arc protection.

[0003] Currently, traditional mechanical power load controllers or solid-state power controllers (SSPC) are commonly used in aviation.

[0004] However, mechanical power load controllers have disadvantages such as slow response speed, limited mechanical control accuracy, high maintenance requirements, and large electromagnetic interference. Solid-state power controllers (SSPC) are mainly suitable for small current loads. For the high-voltage DC grid of electric aircraft, existing mechanical power load controllers or solid-state power controllers are difficult to meet the capacitive load and arc protection requirements.

[0005] Based on this, the present invention designs an electric aircraft power load controller system and a working method thereof to solve the above problems. Summary of the invention

[0006] In view of the above-mentioned shortcomings of the prior art, the present invention provides an electric aircraft power load controller system and a working method thereof.

[0007] To achieve the above objectives, the present invention is implemented through the following technical solutions:

[0008] An electric aircraft power load controller system includes the following parts:

[0009] A main power module is used to control the switching on or off of the high-voltage DC grid;

[0010] Pre-charging module, which charges the equivalent capacitor slowly or pre-charges it with a small current to prevent the impact or damage to the system or load caused by instantaneous large current;

[0011] The discharge module gradually reduces the load voltage to a safe level by first releasing the electrical energy stored in the equivalent capacitor, ensuring that the charge on the load is safely released, thereby avoiding high voltage or arcing when disconnecting;

[0012] Monitor and control the module, collect the status parameters of the main power module, pre-charge module, discharge module and power module, control the working status of each module according to the aircraft's command, and store and report the equipment status;

[0013] A power module, used to receive multiple redundant electric energies and supply power to the monitoring and control module;

[0014] The terminal is used to issue aircraft control instructions to the monitoring and control module and receive device status uploaded by the monitoring and control module;

[0015] A communication module, used for monitoring information transmission between the control module and the terminal;

[0016] The monitoring control module is electrically connected to the communication module, and the communication module is wirelessly connected to the terminal;

[0017] The power module is electrically connected to the monitoring and control module, the main power module is electrically connected to the pre-charging module and the discharge module, and the monitoring and control module is electrically connected to the main power module, the pre-charging module, and the discharge module.

[0018] Furthermore, the main power module includes an a terminal, a b terminal, a c terminal and a d terminal, the two ends of the pre-charging module are electrically connected to the a terminal and the c terminal of the main power module respectively, the two ends of the discharge module are electrically connected to the c terminal and the d terminal of the main power module respectively, and the c terminal and the d terminal of the main power module are electrically connected to the two ends of the load respectively.

[0019] Furthermore, the main power module adopts a high-power mechanical switch, and the pre-charging module includes a solid-state power switch and a current limiting resistor. The solid-state power switch is electrically connected to the current limiting resistor, the solid-state power switch is electrically connected to the A end of the main power module, and the current limiting resistor is electrically connected to the C end of the main power module.

[0020] Furthermore, high power mechanical switches use contactors or relays.

[0021] Furthermore, the solid-state power switch adopts a semiconductor power triode switch.

[0022] In order to better achieve the purpose of the present invention, the present invention also provides a working method of an electric aircraft power load controller system, comprising the following steps:

[0023] Step 1: The high-voltage electrical equipment is powered through the busbar;

[0024] Step 2: The monitoring control module receives instructions through external communication to determine whether the high-voltage electrical equipment needs to work;

[0025] Step 3: If the high-voltage electrical equipment needs to work, the solid-state power switch in the pre-charging module is closed, and the bus bar supplies power to the capacitor in the high-voltage electrical equipment through the current-limiting resistor in the pre-charging module;

[0026] If the high-voltage electrical equipment does not need to work, the thread is terminated directly;

[0027] Step 4: After the bus bar supplies power to the capacitor in the high-voltage electrical equipment through the current-limiting resistor in the pre-charging module, the monitoring and control module determines whether the pre-charging capacitor voltage in the pre-charging module reaches the power supply threshold;

[0028] Step 5: If the pre-charge capacitor voltage does not reach the power supply threshold, return to step 3;

[0029] If the pre-charge capacitor voltage reaches the power supply threshold, the solid-state power switch in the pre-charge module is disconnected, the high-power mechanical switch in the main power module is closed, and the busbar directly charges the high-voltage electrical equipment;

[0030] Step 6: After the bus bar directly charges the high-voltage electrical equipment, the monitoring control module receives instructions through external communication to determine whether the high-voltage electrical equipment stops working;

[0031] Step 7: If the high-voltage electrical equipment continues to work, return to step 5;

[0032] Step 8: If the high-voltage electrical equipment stops working, the capacitor or inductor in the high-voltage electrical equipment is discharged through the discharge module, and then the thread is terminated.

[0033] Compared with the prior art, the present invention has the following beneficial effects: 1. The load control device based on the hybrid power controller of the present invention can be applied to the application scenario of high-voltage and high-power electrical load of electric aircraft by connecting the solid-state power switch and the high-power mechanical switch in parallel and setting the control timing of the two switches, and has good capacitive load adaptability and arc protection, realizes dynamic management of loading and unloading of high-voltage and high-power electrical equipment, and also avoids high-voltage arc;

[0034] 2. The pre-charging module of the present invention utilizes the impedance characteristics of the semiconductor power switch in the saturation region to simulate the pre-charging resistance, and realizes controllable pre-charging of the equivalent capacitance of the high-voltage line by controlling the semiconductor power switch, thereby reducing the complexity and weight of the equipment. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the prior art descriptions are briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention, and for ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0036] Figure 1 A connection block diagram of an electric aircraft power load controller system of the present invention;

[0037] Figure 2 It is a connection block diagram of the main power module, the pre-charge module and the discharge module of the present invention;

[0038] Figure 3 A working diagram of a power load controller system for an electric aircraft according to the present invention;

[0039] Figure 4 This is a diagram of the energy transmission path from the power battery to the motor of the present invention.

[0040] The numbers in the figure represent:

[0041] 1. Main power module; 2. Pre-charging module; 3. Discharge module; 4. Monitoring and control module; 5. Power module; 6. Communication module; 7. Terminal. DETAILED DESCRIPTION

[0042] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0043] Embodiment 1: In some embodiments, please refer to the drawings of the specification Figure 1-Figure 3 , including the following parts:

[0044] A main power module 1 is used to control the switching on or off of the high-voltage DC grid;

[0045] Pre-charging module 2, which slowly charges the equivalent capacitor or pre-charges it with a relatively small current to prevent the impact or damage to the system or load caused by the instantaneous large current;

[0046] The discharge module 3 gradually reduces the load voltage to a safe level by first releasing the electric energy stored in the equivalent capacitor, ensuring that the charge on the load is safely released, thereby avoiding high voltage or arc when disconnecting;

[0047] The monitoring and control module 4 collects the status parameters of the main power module 1, the pre-charging module 2, the discharge module 3 and the power module 5, controls the working status of each module according to the command of the aircraft, and stores and reports the equipment status;

[0048] The power supply module 5 is used to receive multiple redundant electric energies and supply power to the monitoring and control module 4;

[0049] Terminal 7, used to issue aircraft control instructions to monitoring and control module 4 and receive device status uploaded by monitoring and control module 4;

[0050] Communication module 6, used for monitoring information transmission between control module 4 and terminal 7;

[0051] The monitoring control module 4 is electrically connected to the communication module 6, and the communication module 6 is wirelessly connected to the terminal 7;

[0052] The power module 5 is electrically connected to the monitoring and control module 4 , the main power module 1 is electrically connected to the pre-charging module 2 and the discharge module 3 , and the monitoring and control module 4 is electrically connected to the main power module 1 , the pre-charging module 2 , and the discharge module 3 .

[0053] When the aircraft commands the high-voltage and high-power load to power on, the monitoring and control module 4 first sends a pre-charging instruction to the pre-charging module 2, and then closes the main power module 1 when the output voltage meets the requirements, and the high-voltage and high-power load enters the working state; when the aircraft commands the high-voltage and high-power load to power off, the monitoring and control module 4 first sends an instruction to discharge excess electric energy to the discharge module 3, and then disconnects the main power module 1 when the output voltage meets the requirements, and the high-voltage and high-power load shuts down.

[0054] Embodiment 2: In some embodiments, Figure 1 and Figure 2 As shown, as a preferred embodiment of the present invention, the main power module 1 includes an a terminal, a b terminal, a c terminal and a d terminal, the two ends of the pre-charging module 2 are electrically connected to the a terminal and the c terminal of the main power module 1 respectively, the two ends of the discharge module 3 are electrically connected to the c terminal and the d terminal of the main power module 1 respectively, and the c terminal and the d terminal of the main power module 1 are electrically connected to the two ends of the load respectively.

[0055] The main power module 1 adopts a high-power mechanical switch, and the pre-charging module 2 includes a solid-state power switch and a current limiting resistor. The solid-state power switch is electrically connected to the current limiting resistor, the solid-state power switch is electrically connected to the a end of the main power module 1, and the current limiting resistor is electrically connected to the c end of the main power module 1.

[0056] The high-power mechanical switch and the solid-state power switch are connected in parallel.

[0057] The solid-state power switch adopts a semiconductor power triode switch. The impedance characteristics of the semiconductor power triode switch in the saturation region (when the base current of the triode increases to a certain extent, the collector current no longer increases with the increase of the base current, but remains unchanged near a certain value. At this time, the triode loses the current amplification function and is in a saturated state. When saturated, the emitter junction and the collector junction of the triode are both in a forward voltage) are used to simulate the pre-charging resistance, and the controllable pre-charging of the equivalent capacitance of the high-voltage line is achieved by controlling the semiconductor power triode switch.

[0058] The circuit operation of the load, main power module 1, pre-charging module 2 and discharge module 3 is as follows:

[0059] 1. When the load needs to work, the high-power mechanical switch of the main power module 1 is disconnected, the solid-state power switch in the pre-charging module 2 is closed, and the current is charged through the pre-charging module 2;

[0060] 2. When the load stops working, the high-power mechanical switch of the main power module 1 is disconnected, and the load charge is discharged through the discharge module 3;

[0061] 3. When the load is working, the high-power mechanical switch of the main power module 1 is closed, and the current only passes through the main power module 1, and the pre-charge module 2 and the discharge module 3 are both disconnected.

[0062] Embodiment 3: In some embodiments, Figure 3 As shown, as a preferred embodiment of the present invention, a working method of an electric aircraft power load controller system includes the following steps:

[0063] Step 1: The high-voltage electrical equipment is powered through the busbar;

[0064] Step 2: The monitoring control module 4 receives instructions through external communication to determine whether the high-voltage electrical equipment needs to work;

[0065] Step 3: If the high-voltage electrical equipment needs to work, the solid-state power switch in the pre-charging module 2 is closed, and the bus bar supplies power to the capacitor in the high-voltage electrical equipment through the current-limiting resistor in the pre-charging module 2;

[0066] If the high-voltage electrical equipment does not need to work, the thread is terminated directly;

[0067] Step 4: After the bus bar supplies power to the capacitor in the high-voltage electrical equipment through the current-limiting resistor in the pre-charging module 2, the monitoring and control module 4 determines whether the pre-charging capacitor voltage in the pre-charging module 2 reaches the power supply threshold;

[0068] Step 5: If the pre-charge capacitor voltage does not reach the power supply threshold, return to step 3;

[0069] If the pre-charge capacitor voltage reaches the power supply threshold, the solid-state power switch in the pre-charge module 2 is disconnected, the high-power mechanical switch in the main power module 1 is closed, and the busbar directly charges the high-voltage electrical equipment;

[0070] Step 6: After the bus bar directly charges the high-voltage electrical equipment, the monitoring control module 4 receives instructions through external communication to determine whether the high-voltage electrical equipment stops working;

[0071] Step 7: If the high-voltage electrical equipment continues to work, return to step 5;

[0072] Step 8: If the high-voltage electrical equipment stops working, the capacitor or inductor in the high-voltage electrical equipment is discharged through the discharge module 3, and then the thread ends.

[0073] Embodiment 4: In some embodiments, Figure 4 As shown, attached Figure 4 The figure describes a DC high-voltage power grid for electric aircraft, including the energy transmission path from the power battery to the motor. The figure only shows a typical end-to-end power supply link and does not involve the overall power grid architecture. It is only for illustrating typical application scenarios and is not limited to this type of power grid.

[0074] The above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit the same. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that the technical solutions described in the aforementioned embodiments may still be modified, or some of the technical features may be replaced by equivalents. Such modifications or replacements will not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. An electric aircraft power load controller system, characterized in that: Includes the following sections: A main power module (1), used for controlling the switching on or off of the high voltage DC grid; A pre-charging module (2) is used to slowly charge the equivalent capacitor or pre-charge it with a relatively small current to prevent the system or load from being impacted or damaged by a large instantaneous current; The discharge module (3) gradually reduces the load voltage to a safe level by first releasing the electric energy stored in the equivalent capacitor, thereby ensuring that the charge on the load is safely released, thereby avoiding the generation of high voltage or arc when disconnecting; A monitoring control module (4) collects status parameters of the main power module (1), the pre-charging module (2), the discharge module (3) and the power module (5), controls the working status of each module according to the command of the aircraft, and stores and reports the equipment status; A power supply module (5), used for receiving multiple redundant electric energies and supplying power to the monitoring and control module (4); The terminal (7) is used to issue control instructions of the aircraft to the monitoring and control module (4) and receive the device status uploaded by the monitoring and control module (4); A communication module (6) for monitoring information transmission between the control module (4) and the terminal (7); The monitoring and control module (4) is electrically connected to the communication module (6), and the communication module (6) is wirelessly connected to the terminal (7); The power module (5) is electrically connected to the monitoring and control module (4), the main power module (1) is electrically connected to the pre-charging module (2) and the discharge module (3), and the monitoring and control module (4) is electrically connected to the main power module (1), the pre-charging module (2) and the discharge module (3).

2. The electric aircraft power load controller system according to claim 1, characterized in that: The main power module (1) comprises an a terminal, a b terminal, a c terminal and a d terminal, the two ends of the pre-charging module (2) are respectively electrically connected to the a terminal and the c terminal of the main power module (1), the two ends of the discharge module (3) are respectively electrically connected to the c terminal and the d terminal of the main power module (1), and the c terminal and the d terminal of the main power module (1) are respectively electrically connected to the two ends of the load.

3. The electric aircraft power load controller system according to claim 2, characterized in that: The main power module (1) adopts a high-power mechanical switch.

4. The electric aircraft power load controller system according to claim 3, characterized in that: The pre-charging module (2) comprises a solid-state power switch and a current-limiting resistor, the solid-state power switch is electrically connected to the current-limiting resistor, the solid-state power switch is electrically connected to the a terminal of the main power module (1), and the current-limiting resistor is electrically connected to the c terminal of the main power module (1).

5. The electric aircraft power load controller system according to claim 4, characterized in that: The high-power mechanical switch and the solid-state power switch are connected in parallel.

6. The electric aircraft power load controller system according to claim 5, characterized in that: The high-power mechanical switch adopts a contactor or a relay.

7. The electric aircraft power load controller system according to claim 6, characterized in that: The solid-state power switch adopts a semiconductor power triode switch.

8. A working method of an electric aircraft power load controller, used in the electric aircraft power load controller system of claim 7, characterized in that: The following steps are involved: Step 1: The high-voltage electrical equipment is powered through the busbar; Step 2: The monitoring control module (4) receives instructions through external communication to determine whether the high-voltage electrical equipment needs to work; Step 3: If the high-voltage electrical equipment needs to work, the solid-state power switch in the pre-charging module (2) is closed, and the bus bar supplies power to the capacitor in the high-voltage electrical equipment through the current-limiting resistor in the pre-charging module (2); If the high-voltage electrical equipment does not need to work, the thread is terminated directly; Step 4: After the bus bar supplies power to the capacitor in the high-voltage electrical equipment through the current-limiting resistor in the pre-charging module (2), the monitoring control module (4) determines whether the pre-charging capacitor voltage in the pre-charging module (2) reaches the power supply threshold; Step 5: If the pre-charge capacitor voltage does not reach the power supply threshold, return to step 3; If the pre-charge capacitor voltage reaches the power supply threshold, the solid-state power switch in the pre-charge module (2) is disconnected, the high-power mechanical switch in the main power module (1) is closed, and the busbar directly charges the high-voltage electrical equipment; Step 6: After the bus bar directly charges the high-voltage electrical equipment, the monitoring control module (4) receives instructions through external communication to determine whether the high-voltage electrical equipment stops working; Step 7: If the high-voltage electrical equipment continues to work, return to step 5; Step 8: If the high-voltage electrical equipment stops working, the capacitor or inductor in the high-voltage electrical equipment is discharged through the discharge module (3), and then the thread ends.

Images