Spacecraft power-up device and spacecraft

Abstract

The utility model discloses a spacecraft power-up device and a spacecraft. The spacecraft power-up device comprises a switch module, a driving switch module and a driving module, the switch module comprises at least one NMOS tube, the NMOS tube is connected between the positive electrode of the battery pack and a load, the driving switch module is connected between the positive electrode of the battery pack and the driving module, the driving module is connected with the switch module, and the driving module is further connected with the negative electrode of the battery pack; the driving switch module is used for controlling the driving module to be powered on when the spacecraft is separated from the carrier rocket; the driving module is used for controlling the switch-on or switch-off of the switch module; the switch module is used for controlling charging and discharging of the battery pack. According to the embodiment of the utility model, the switch module is driven to control the power-on or power-off of the driving module, and the NMOS tube of the switch module is used to control the discharge of the battery pack, so that the heat consumption of the power-on device can be reduced.

Description

Technical Field

[0001] This utility model relates to the field of aerospace technology, and in particular to a spacecraft power supply device and a spacecraft. Background Technology

[0002] Satellite-rocket separation technology is one of the key technologies in the aerospace field. Whether the satellite can separate correctly directly affects the success of the rocket launch and the satellite's successful entry into orbit, significantly impacting the overall performance of the satellite and the completion of the mission. The design of the satellite-rocket separation system must ensure reliable connection and separation between the satellite and the rocket, while also meeting requirements regarding separation speed, separation attitude, and impact response.

[0003] After separation from the launch vehicle, the satellite needs to obtain energy from the solar panels and batteries to achieve normal on-orbit operation and mission execution. However, the solar panels are in a compressed state at this time, and the deployment process takes a certain amount of time, which cannot meet the power requirements of the satellite in time. At this time, the energy supply is entirely provided by the batteries. Therefore, in order to save power, the batteries in the system are not connected to the bus before separation from the launch vehicle. Therefore, it is necessary to design a corresponding power supply device so that the batteries can automatically connect to the bus after separation from the launch vehicle, and complete the power supply of the entire system.

[0004] However, in the existing technology, the internal resistance of the power supply device is relatively large during operation, which also leads to high heat consumption of the autonomous power supply device. Utility Model Content

[0005] This invention provides a spacecraft power supply device and a spacecraft to reduce the heat consumption of the power supply device.

[0006] According to one aspect of the present invention, a spacecraft power supply device is provided, the spacecraft power supply device comprising: a switch module, a drive switch module, and a drive module;

[0007] The switching module includes at least one NMOS transistor connected between the positive terminal of the battery pack and the load. The driving switching module is connected between the positive terminal of the battery pack and the driving module. The driving module is connected to the switching module and is also connected to the negative terminal of the battery pack.

[0008] The drive switch module is used to control the power-on of the drive module when the spacecraft separates from the launch vehicle; the drive module is used to control the conduction or deactivation of the switch module; the switch module is used to control the charging or discharging of the battery pack.

[0009] Optionally, the drive switch module includes: at least one limit switch;

[0010] The first end of the limit switch is connected to the positive terminal of the battery pack, and the second end of the limit switch is connected to the drive module;

[0011] The limit switch is used to control the power-on of the drive module when the spacecraft separates from the launch vehicle.

[0012] Optionally, the driving module includes: a signal generating unit and a driving unit;

[0013] The first end of the signal generating unit is connected to the driving switch module, the second end of the signal generating unit is connected to the first end of the driving unit, the second end of the driving unit is connected to the first end of the signal generating unit, and the third end of the driving unit is connected to the switch module.

[0014] The signal generating unit is used to generate a square wave signal; the driving unit is used to control the switching module to turn on or off according to the square wave signal.

[0015] Optionally, the driving unit includes: a first diode, a first resistor, a first switching transistor, a second resistor, a second switching transistor, a second diode, a third diode, a third resistor, a fourth resistor, a third switching transistor, a fifth resistor, a first capacitor, a second capacitor, and a third capacitor;

[0016] The anode of the first diode is connected to the signal generating unit, the cathode of the first diode is connected to the first end of the first resistor, the second end of the first resistor is connected to the control terminal of the first switching transistor, the first end of the first switching transistor is connected to the first end of the second resistor, the second end of the second resistor is connected to the driving switch module, the second end of the first switching transistor is grounded, the control terminal of the second switching transistor is connected to the first end of the first switching transistor, the first end of the second switching transistor is connected to the driving switch module, the second end of the second switching transistor is connected to the anode of the second diode, the cathode of the second diode is connected to the anode of the third diode, and the cathode of the third diode is connected to the switch module. The first terminal of the three resistors is connected to the second terminal of the second switching transistor. The second terminal of the third resistor is connected to the first terminal of the fourth resistor. The second terminal of the fourth resistor is connected to the first terminal of the third switching transistor. The second terminal of the third switching transistor is grounded. The control terminal of the third switching transistor is connected to the first terminal of the fifth resistor. The second terminal of the fifth resistor is connected to the anode of the first diode. The first terminal of the first capacitor is connected to the cathode of the first diode. The second terminal of the first capacitor is grounded. The first terminal of the second capacitor is connected to the second terminal of the third resistor. The second terminal of the second capacitor is connected to the cathode of the second diode. The first terminal of the third capacitor is connected to the cathode of the third diode. The second terminal of the third capacitor is grounded.

[0017] Optionally, the spacecraft power supply device also includes: a solar panel protection module, a charging module, and a power supply switch module;

[0018] The input terminal of the solar panel protection module is connected to the solar panel, the output terminal of the solar panel protection module is connected to the output terminal of the switch module, the input terminal of the switch module is connected to the positive terminal of the battery pack, the negative terminal of the battery pack is grounded, the charging module is connected in parallel with the switch module, the power-on switch module is connected between the positive terminal of the battery pack and the drive switch module, the control module is connected to the battery pack, and the control module is also connected to the power-on switch module;

[0019] The solar panel protection module is used to limit the direction of the output current of the solar panel; the charging module is used to charge the battery pack when the battery pack is over-discharged; the power-on switch module is used to control the drive switch module to power off when the battery pack is over-discharged; the control module is used to control the power-on switch module to turn on or off.

[0020] Optionally, the solar panel protection module includes: a first protection diode and a second protection diode;

[0021] The anode of the first protection diode is connected to the solar panel, and the cathode of the first protection diode is connected to the switching module. The anode of the second protection diode is connected to the solar panel, and the cathode of the second protection diode is connected to the switching module.

[0022] Optionally, the charging module includes: a plurality of charging diodes;

[0023] The charging diodes are connected in parallel, with the anode of each charging diode connected to the solar panel protection module and the cathode of each charging diode connected to the battery pack.

[0024] Optionally, the power-on switch module includes: at least one power-on relay;

[0025] The input terminal of the power-on relay is connected to the positive terminal of the battery pack, the output terminal of the power-on relay is connected to the drive switch module, the control terminal of the power-on relay is connected to the control module, and the negative terminal of the battery pack is grounded.

[0026] Optionally, the spacecraft power supply device also includes: a test switch module;

[0027] The test switch module is connected between the positive terminal of the battery pack and the drive module;

[0028] The test switch module is used to control the power-on or power-off of the drive module during ground testing.

[0029] According to another aspect of the present invention, a spacecraft is also provided, comprising: a solar array, a battery pack, and a spacecraft power supply device as described in any of the above embodiments.

[0030] This embodiment of the invention detects the release status of the solar array via a drive switch module. When the solar array is pressed down, the drive switch module disconnects the connection between the battery pack and the drive module, thereby turning off the switch module. When the spacecraft separates from the launch vehicle, the drive switch module reconnects the connection between the battery pack and the drive module, thereby turning on the switch module and powering on the load. This embodiment of the invention controls the power-on and power-off of the drive module via the drive switch module and controls the discharge of the battery pack through the NMOS transistor of the switch module, which helps to reduce the heat dissipation of the power supply device.

[0031] It should be understood that the description in this section is not intended to identify key or essential features of the embodiments of this utility model, nor is it intended to limit the scope of this utility model. Other features of this utility model will become readily apparent from the following description. Attached Figure Description

[0032] To more clearly illustrate the technical solutions in the embodiments of this utility model, the drawings used in the description of the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this utility model. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0033] Figure 1 This is a schematic diagram of a spacecraft power supply device provided in an embodiment of this utility model;

[0034] Figure 2 This is a schematic diagram of another spacecraft power supply device provided in an embodiment of the present invention;

[0035] Figure 3 This is a schematic diagram of another spacecraft power supply device provided in this embodiment of the utility model;

[0036] Figure 4 This is a schematic diagram of another spacecraft power supply device provided in this embodiment of the utility model;

[0037] Figure 5 This is a schematic diagram of another spacecraft power supply device provided in this embodiment of the utility model;

[0038] Figure 6 This is a schematic diagram of another spacecraft power supply device provided in this embodiment of the utility model;

[0039] Figure 7 This is a schematic diagram of another spacecraft power supply device provided in this embodiment of the utility model;

[0040] Figure 8 This is a schematic diagram of another spacecraft power supply device provided in this embodiment of the utility model;

[0041] Figure 9 This is a schematic diagram of a spacecraft provided in an embodiment of the present utility model. Detailed Implementation

[0042] To enable those skilled in the art to better understand the present invention, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings of the embodiments. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort should fall within the protection scope of the present invention.

[0043] It should be noted that the terms "first," "second," etc., in the specification, claims, and accompanying drawings of this utility model are used to distinguish similar objects and are not necessarily used to describe a specific order or sequence. It should be understood that such data can be interchanged where appropriate so that the embodiments of the utility model described herein can be implemented in orders other than those illustrated or described herein. Furthermore, the terms "comprising" and "having," and any variations thereof, are intended to cover non-exclusive inclusion; for example, a process, method, system, product, or apparatus that comprises a series of steps or units is not necessarily limited to those steps or units explicitly listed, but may include other steps or units not explicitly listed or inherent to such processes, methods, products, or apparatus.

[0044] This utility model provides a spacecraft power supply device. This device is applied to a spacecraft. In this embodiment, the drive switch module controls the power-on of the drive module when the spacecraft's solar panels deploy, thereby controlling the conduction of the NMOS transistor in the switch module, which helps reduce the heat dissipation of the power supply device. Figure 1 This is a schematic diagram of a spacecraft power supply device provided in an embodiment of this utility model. (Refer to...) Figure 1 The spacecraft power supply device includes: a switch module 110, a drive switch module 120, and a drive module 130.

[0045] The switching module 110 includes at least one NMOS transistor M1, which is connected between the positive terminal of the battery pack 200 and the load 300. The drive switching module 120 is connected between the positive terminal of the battery pack 200 and the drive module 130. The drive module 130 is connected to the switching module 110 and also to the negative terminal of the battery pack 200. The drive switching module 120 is used to control the drive module 130 to power on when the spacecraft separates from the launch vehicle. The drive module 130 is used to control the switching module 110 to turn on or off. The switching module 110 is used to control the discharge of the battery pack 200.

[0046] Specifically, the on / off state of the drive switch module 120 is related to whether the solar array is released. For example, the drive switch module 120 can detect whether the spacecraft has been released by detecting the operational status of the spacecraft unlocking mechanism. When the spacecraft is not separated from the launch vehicle, the drive switch module 120 disconnects the connection between the battery pack 200 and the drive module 130; at this time, the drive module 130 is powered down, and the switch module 110 is turned off. When the spacecraft separates from the launch vehicle, the drive switch module 120 connects the connection between the battery pack 200 and the drive module 130; at this time, the drive module 130 is powered on and drives the switch module 110, turning it on. When the switch module 110 is on, the connection between the battery pack 200 and the load 300 is established, and the battery pack 200 discharges.

[0047] It should be noted that the NMOS transistors M1 in the switching module 110 are driven by the driver module 130. When the switching module 110 has multiple NMOS transistors M1, the NMOS transistors M1 are connected in parallel, and the NMOS transistors M1 in the switching module 110 operate synchronously. That is, all NMOS transistors M1 in the switching module 110 are simultaneously turned on or simultaneously turned off. In practical applications, optionally, at least two NMOS transistors M1 can be set in the switching module 110 for redundancy backup to ensure the normal operation of the switching module 110.

[0048] This embodiment of the invention uses a drive switch module 120 to detect the release status of the solar array. When the solar array is pressed down, the drive switch module 120 disconnects the connection between the battery pack 200 and the drive module 130, thereby turning off the switch module 110. When the spacecraft separates from the launch vehicle, the drive switch module 120 connects the battery pack 200 and the drive module 130, thereby turning on the switch module 110 and powering on the load 300. This embodiment of the invention controls the power-on or power-off of the drive module 130 through the drive switch module 120, and controls the discharge of the battery pack 200 through the NMOS transistor M1 of the switch module 110, which helps to reduce the heat dissipation of the power supply device.

[0049] Figure 2 This is a schematic diagram of another spacecraft power supply device provided in an embodiment of the present invention. Optionally, based on the above embodiments, refer to... Figure 2 The drive switch module 120 includes at least one limit switch 121.

[0050] The first end of the limit switch 121 is connected to the positive terminal of the battery pack 200, and the second end of the limit switch 121 is connected to the drive module 130. The limit switch 121 is used to control the drive module 130 to be powered on when the spacecraft separates from the launch vehicle.

[0051] Specifically, the limit switch 121 is linked to the unlocking mechanism of the spacecraft in the launch vehicle. When the spacecraft's unlocking mechanism is pressed, the limit switch 121 is also pressed, at which point the limit switch 121 is turned off, the connection between the battery pack 200 and the drive module 130 is disconnected, and the drive module 130 is powered down; when the spacecraft's unlocking mechanism is released, the limit switch 121 is also released, at which point the limit switch 121 is turned on, the connection between the battery pack 200 and the drive module 130 is established, and the drive module 130 is powered on.

[0052] Figure 3 This is a schematic diagram of another spacecraft power supply device provided in an embodiment of the present invention. Optionally, based on the above embodiments, refer to... Figure 3The drive module 130 includes a signal generation unit 131 and a drive unit 132.

[0053] The first end of the signal generating unit 131 is connected to the drive switch module 120, the second end of the signal generating unit 131 is connected to the first end of the drive unit 132, the second end of the drive unit 132 is connected to the first end of the signal generating unit 131, and the third end of the drive unit 132 is connected to the switch module 110. The signal generating unit 131 is used to generate a square wave signal. The drive unit 132 is used to control the switch module 110 to turn on or off according to the square wave signal.

[0054] Specifically, when the drive switch module 120 is turned on, the signal generation unit 131 and the drive unit 132 are powered on. At this time, the signal generation unit 131 generates a square wave signal. The drive unit 132 acquires the square wave signal generated by the signal generation unit 131 and drives the NMOS transistor in the switch module 110 to turn on according to the square wave signal, thereby powering on the load 300.

[0055] Figure 4 This is a schematic diagram of another spacecraft power supply device provided in an embodiment of the present invention. Optionally, based on the above embodiments, refer to... Figure 4 The driving unit 132 includes: a first diode D1, a first resistor R1, a first switch Q1, a second resistor R2, a second switch Q2, a second diode D2, a third diode D3, a third resistor R3, a fourth resistor R4, a third switch Q3, a fifth resistor R5, a first capacitor C1, a second capacitor C2, and a third capacitor C3.

[0056] The anode of the first diode D1 is connected to the signal generating unit 131. The cathode of the first diode D1 is connected to the first end of the first resistor R1. The second end of the first resistor R1 is connected to the control terminal of the first switch Q1. The first end of the first switch Q1 is connected to the first end of the second resistor R2. The second end of the second resistor R2 is connected to the drive switch module 120. The second end of the first switch Q1 is grounded. The control terminal of the second switch Q2 is connected to the first end of the first switch Q1. The first end of the second switch Q2 is connected to the drive switch module 120. The second end of the second switch Q2 is connected to the anode of the second diode D2. The cathode of the second diode D2 is connected to the anode of the third diode D3. The cathode of the third diode D3 is connected to the switch module 110. The first terminal of resistor R3 is connected to the second terminal of the second switch Q2. The second terminal of resistor R3 is connected to the first terminal of resistor R4. The second terminal of resistor R4 is connected to the first terminal of the third switch Q3. The second terminal of the third switch Q3 is grounded. The control terminal of the third switch Q3 is connected to the first terminal of resistor R5. The second terminal of resistor R5 is connected to the anode of diode D1. The first terminal of capacitor C1 is connected to the cathode of diode D1. The second terminal of capacitor C1 is grounded. The first terminal of capacitor C2 is connected to the second terminal of resistor R3. The second terminal of capacitor C2 is connected to the cathode of diode D2. The first terminal of capacitor C3 is connected to the cathode of diode D3. The second terminal of capacitor C3 is grounded.

[0057] Figure 5 This is a schematic diagram of another spacecraft power supply device provided in an embodiment of the present invention. Optionally, based on the above embodiments, refer to... Figure 5 The spacecraft's power supply device also includes: a solar panel protection module 140, a charging module 150, and a power-on switch module 160.

[0058] The input terminal of the solar panel protection module 140 is connected to the solar panel 400, and the output terminal of the solar panel protection module 140 is connected to the output terminal of the switch module 110. The input terminal of the switch module 110 is connected to the positive terminal of the battery pack 200, and the negative terminal of the battery pack 200 is grounded. The charging module 150 is connected in parallel with the switch module 110. The power-on switch module 160 is connected between the positive terminal of the battery pack 200 and the drive switch module 120. The control module 170 is connected to the battery pack 200 and is also connected to the power-on switch module 160. The solar panel protection module 140 is used to limit the direction of the output current of the solar panel 400. The charging module 150 is used to charge the battery pack 200 when it is over-discharged. The power-on switch module 160 is used to control the drive switch module 120 to power off when the battery pack 200 is over-discharged. The control module 170 is used to control the power-on or power-off of the power-on switch module 160.

[0059] Specifically, the electrical energy stored in the battery pack 200 is limited by its own capacity, and the amount of electrical energy stored in the battery pack 200 is related to its output voltage; the greater the electrical energy stored in the battery pack 200, the greater its output voltage. After the spacecraft launches, it generates electricity through its solar array 400, which is then transmitted to the load 300 via the solar array protection module 140. In other words, after the solar array 400 deploys, the load 300 is powered by both the battery pack 200 and the solar array 400. When the output voltage of the battery pack 200 is lower than the voltage of the electrical energy generated by the solar array 400, the battery pack 200 is charged via the switching module 110. That is, at this time, the solar array 400 simultaneously supplies power to the load 300 and charges the battery pack 200.

[0060] When the spacecraft's solar array 400 is not fully deployed, the load 300 is powered entirely by the battery pack 200. Therefore, when the solar array 400 is fully deployed, the battery pack 200 may be over-discharged. The control module 170 detects the state of the battery pack 200 and obtains the deployment state of the solar array 400. When the solar array 400 is fully deployed and the battery pack 200 is over-discharged, the control module 170 controls the power-on switch module 160 to disconnect the connection between the battery pack 200 and the drive switch module 120, at which point the drive switch module 120 is powered down. When the drive switch module 120 is powered down, the drive module 130 is also powered down simultaneously, and the NMOS transistor M1 in the switch module 110 is turned off. In this situation, the electrical energy generated by the solar array 400 is delivered to the battery pack 200 through the solar array protection module 140 and the charging module 150 to charge the battery pack 200. When the battery pack 200 is charged to a non-over-discharged state, the control module 170 controls the power-on switch module 160 to turn on again, thereby turning on the NMOS transistor in the switch module 110. After the switch module 110 is turned on, the battery pack 200 is charged through the solar panel protection module 140 and the switch module 110.

[0061] Figure 6 This is a schematic diagram of another spacecraft power supply device provided in an embodiment of the present invention. Optionally, based on the above embodiments, refer to... Figure 6 The solar panel protection module 140 includes: a first protection diode D4 and a second protection diode D5.

[0062] The anode of the first protection diode D4 is connected to the solar panel 400, and the cathode of the first protection diode D4 is connected to the switch module 110. The anode of the second protection diode D5 is connected to the solar panel 400, and the cathode of the second protection diode D5 is connected to the switch module 110.

[0063] Specifically, the first protection diode D4 and the second protection diode D5 are used to limit the direction of the electrical energy generated by the solar panel 400 in order to prevent the electrical energy output by the battery pack 200 from impacting the solar panel 400.

[0064] Based on the above embodiments, optionally, refer to... Figure 6 The charging module 150 includes: multiple charging diodes D6.

[0065] The charging diodes D6 are connected in parallel. The anode of the charging diode D6 is connected to the solar panel protection module 140, and the cathode of the charging diode D6 is connected to the battery pack 200. Specifically, the charging diodes D6 are used to limit the direction of electrical energy during charging of the battery pack 200, thereby preventing the battery pack 200 from discharging.

[0066] Based on the above embodiments, optionally, refer to... Figure 6 The power-on switch module 160 includes at least one power-on relay S1.

[0067] The input terminal of the power-on relay S1 is connected to the positive terminal of the battery pack 200, the output terminal of the power-on relay S1 is connected to the drive switch module 120, the control terminal of the power-on relay S1 is connected to the control module 170, and the negative terminal of the battery pack 200 is grounded. It should be noted that in practical applications, the power-on switch module 160 can be equipped with at least two power-on relays S1 for redundancy, ensuring the normal operation of the power-on switch module 160.

[0068] Figure 7 This is a schematic diagram of another spacecraft power supply device provided in this embodiment of the utility model. Figure 8 This is a schematic diagram of another spacecraft power supply device provided by this utility model embodiment. Based on the above embodiments, optionally, it can be combined with... Figure 7 and Figure 8 The spacecraft power supply device also includes: test switch module 180.

[0069] The test switch module 180 is connected between the positive terminal of the battery pack 200 and the drive module 130; the test switch module 180 is used to control the power-on or power-off of the drive module 130 during ground testing. Optionally, the test switch module 180 may include a test relay S2.

[0070] Specifically, during ground testing of the spacecraft, the power-on or power-off of the drive module 130 is controlled by the test switch module 180 to reduce the wear and tear on the drive switch module 120 during testing and improve the reliability of the drive switch module 120 after launch.

[0071] This utility model embodiment also provides a spacecraft. Figure 9This is a schematic diagram of a spacecraft provided in an embodiment of this utility model. (Refer to...) Figure 9 The spacecraft 10 includes: a solar array 400, a battery pack 200, and a spacecraft power supply device 100 provided in any of the above embodiments.

[0072] The spacecraft 10 provided in this embodiment has the beneficial effects of the spacecraft power supply device 100 provided in any of the above embodiments, which will not be described in detail here.

[0073] It should be understood that the various forms of the process shown above can be used, with steps reordered, added, or deleted. For example, the steps described in this utility model can be executed in parallel, sequentially, or in different orders, as long as the desired result of the technical solution of this utility model can be achieved, and this is not limited herein.

[0074] The specific embodiments described above do not constitute a limitation on the scope of protection of this utility model. Those skilled in the art should understand that various modifications, combinations, sub-combinations, and substitutions can be made according to design requirements and other factors. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of this utility model should be included within the scope of protection of this utility model.

Claims

1. A spacecraft power supply device, characterized in that, include: Switch module, drive switch module, and drive module; The switching module includes at least one NMOS transistor connected between the positive terminal of the battery pack and the load. The driving switching module is connected between the positive terminal of the battery pack and the driving module. The driving module is connected to the switching module and is also connected to the negative terminal of the battery pack. The drive switch module is used to control the power-on of the drive module when the spacecraft separates from the launch vehicle; the drive module is used to control the conduction or deactivation of the switch module; the switch module is used to control the charging or discharging of the battery pack.

2. The spacecraft power supply device according to claim 1, characterized in that, The drive switch module includes: at least one limit switch; The first end of the limit switch is connected to the positive terminal of the battery pack, and the second end of the limit switch is connected to the drive module; The limit switch is used to control the power-on of the drive module when the spacecraft separates from the launch vehicle.

3. The spacecraft power supply device according to claim 1, characterized in that, The driving module includes: a signal generating unit and a driving unit; The first end of the signal generating unit is connected to the driving switch module, the second end of the signal generating unit is connected to the first end of the driving unit, the second end of the driving unit is connected to the first end of the signal generating unit, and the third end of the driving unit is connected to the switch module. The signal generating unit is used to generate a square wave signal; the driving unit is used to control the switching module to turn on or off according to the square wave signal.

4. The spacecraft power supply device according to claim 3, characterized in that, The driving unit includes: a first diode, a first resistor, a first switching transistor, a second resistor, a second switching transistor, a second diode, a third diode, a third resistor, a fourth resistor, a third switching transistor, a fifth resistor, a first capacitor, a second capacitor, and a third capacitor; The anode of the first diode is connected to the signal generating unit, the cathode of the first diode is connected to the first end of the first resistor, the second end of the first resistor is connected to the control terminal of the first switching transistor, the first end of the first switching transistor is connected to the first end of the second resistor, the second end of the second resistor is connected to the driving switch module, the second end of the first switching transistor is grounded, the control terminal of the second switching transistor is connected to the first end of the first switching transistor, the first end of the second switching transistor is connected to the driving switch module, the second end of the second switching transistor is connected to the anode of the second diode, the cathode of the second diode is connected to the anode of the third diode, and the cathode of the third diode is connected to the switch module. The first terminal of the three resistors is connected to the second terminal of the second switching transistor. The second terminal of the third resistor is connected to the first terminal of the fourth resistor. The second terminal of the fourth resistor is connected to the first terminal of the third switching transistor. The second terminal of the third switching transistor is grounded. The control terminal of the third switching transistor is connected to the first terminal of the fifth resistor. The second terminal of the fifth resistor is connected to the anode of the first diode. The first terminal of the first capacitor is connected to the cathode of the first diode. The second terminal of the first capacitor is grounded. The first terminal of the second capacitor is connected to the second terminal of the third resistor. The second terminal of the second capacitor is connected to the cathode of the second diode. The first terminal of the third capacitor is connected to the cathode of the third diode. The second terminal of the third capacitor is grounded.

5. The spacecraft power supply device according to claim 1, characterized in that, Also includes: Solar panel protection module, charging module, and power-on switch module; The input terminal of the solar panel protection module is connected to the solar panel, the output terminal of the solar panel protection module is connected to the output terminal of the switch module, the input terminal of the switch module is connected to the positive terminal of the battery pack, the negative terminal of the battery pack is grounded, the charging module is connected in parallel with the switch module, the power-on switch module is connected between the positive terminal of the battery pack and the drive switch module, the control module is connected to the battery pack, and the control module is also connected to the power-on switch module; The solar panel protection module is used to limit the direction of the output current of the solar panel; the charging module is used to charge the battery pack when the battery pack is over-discharged; the power-on switch module is used to control the drive switch module to power off when the battery pack is over-discharged; the control module is used to control the power-on switch module to turn on or off.

6. The spacecraft power supply device according to claim 5, characterized in that, The solar panel protection module includes: a first protection diode and a second protection diode; The anode of the first protection diode is connected to the solar panel, and the cathode of the first protection diode is connected to the switching module. The anode of the second protection diode is connected to the solar panel, and the cathode of the second protection diode is connected to the switching module.

7. The spacecraft power supply device according to claim 5, characterized in that, The charging module includes: multiple charging diodes; The charging diodes are connected in parallel, with the anode of each charging diode connected to the solar panel protection module and the cathode of each charging diode connected to the battery pack.

8. The spacecraft power supply device according to claim 5, characterized in that, The power-on switch module includes: at least one power-on relay; The input terminal of the power-on relay is connected to the positive terminal of the battery pack, the output terminal of the power-on relay is connected to the drive switch module, the control terminal of the power-on relay is connected to the control module, and the negative terminal of the battery pack is grounded.

9. The spacecraft power supply device according to claim 1, characterized in that, Also includes: Test switch module; The test switch module is connected between the positive terminal of the battery pack and the drive module; The test switch module is used to control the power-on or power-off of the drive module during ground testing.

10. A spacecraft, characterized in that, include: Solar array, battery pack, and spacecraft power supply device as described in any one of claims 1-9.

Images