Satellite electric propulsion system

Abstract

This utility model relates to a satellite electric propulsion system, addressing the problems of high cost, poor control precision, and low efficiency and reliability of existing chemical thrusters for micro and nanosatellites, as well as the issues of low efficiency and low reliability of resistance-heated thrusters. The system includes a propellant tank, a power supply and control module, a dual-path pressure regulation module, and a dual-thrust module. The propellant tank uses a full TC4 titanium alloy spherical structure to store liquid ammonia. The power supply and control module controls the high-voltage arc heating of the working fluid via a CAN bus. The dual-path pressure regulation module achieves redundant pressure reduction through solenoid valves, preheaters, and pressure reducing valves. The dual thrusters are arranged vertically at their centers of mass, providing mutual cold backup. This design reduces size and weight through modular integration, improves electrothermal conversion efficiency, orbit climbing mission success rate, and working fluid utilization efficiency, effectively solving the technical bottlenecks of existing propulsion systems and is suitable for high-precision orbit control of micro and nanosatellites.

Description

Technical Field

[0001] This utility model relates to the field of space thruster technology, and in particular to a satellite electric propulsion system. Background Technology

[0002] With the rapid development of the commercial space industry, microsatellites and nanosatellites, due to their significant advantages of high flexibility and low cost, have become the core carriers for building satellite internet constellations, and their application demand in fields such as remote sensing monitoring, communication, and navigation continues to grow. As a key subsystem of microsatellites and nanosatellites, the performance of the propulsion system directly affects the satellite's orbital control accuracy, mission lifespan, and overall cost. Currently, microsatellite propulsion systems mainly adopt two technical solutions: chemical thrusters and resistance-heated thrusters, but both have insurmountable technical bottlenecks.

[0003] Traditional chemical thrusters rely on the combustion of chemical fuels to generate thrust. While they can provide high instantaneous thrust, their inherent limitations restrict the large-scale application of microsatellites and nanosatellites. On the one hand, propellant storage requires high-pressure containers, and the consumption per mission is large, resulting in high system manufacturing and maintenance costs. On the other hand, chemical thrusters have poor control precision and slow pulse adjustment response speed, making it difficult to meet the high-precision orbit maintenance and attitude control requirements of microsatellites and nanosatellites for complex missions. In addition, their propellant storage and supply components are large in size and weight, occupying too much of the limited payload space of microsatellites and further compressing the design space of the payload.

[0004] In existing technologies, the high cost and low precision of chemical thrusters, and the low efficiency and low reliability of resistance-heated thrusters, have become major technical pain points for micro- and nano-satellite propulsion systems. How to improve thrust performance and propellant utilization efficiency while reducing the weight and manufacturing cost of the propulsion system, and enhancing the reliability of satellite operation in orbit, is a key problem that urgently needs to be solved in the field of micro- and nano-satellite propulsion technology.

[0005] In summary, there is currently a lack of a satellite electric propulsion system to improve the propulsion performance of micro and nano satellites. Utility Model Content

[0006] The purpose of this invention is to overcome the defects of the existing technology and provide a satellite electric propulsion system to solve or partially solve the problems of lack of redundancy backup and poor redundancy backup.

[0007] The objective of this utility model can be achieved through the following technical solutions:

[0008] This utility model provides a satellite electric propulsion system, including:

[0009] The storage tank contains liquid ammonia gas.

[0010] The power supply and control module is used to generate a high-voltage, high-temperature electric arc and control the operation of the thruster.

[0011] A dual-path pressure regulation module, connected to the storage tank, is used to convert the high-pressure gas in the storage tank into working pressure gas and output it to the thrust module in two paths;

[0012] The dual-thrust module, connected to the dual-path pressure regulation module, is used to generate thrust by heating ammonia propellant through electric arc discharge.

[0013] As a preferred technical solution, the dual-channel pressure regulation module includes:

[0014] The FDV1 valve is connected at one end to the storage tank and is used to control the injection and release of ammonia.

[0015] A dual-path pressure reducing sensor unit is connected to the other end of the FDV1 valve and is used to convert the high-pressure gas in the tank into gas at the working pressure.

[0016] The first pressure sensor P1 is connected to the storage tank.

[0017] As a preferred technical solution, in the dual-path pressure reduction sensing unit, one branch includes:

[0018] The first solenoid valve SV1 is connected to the feed / discharge valve FDV1;

[0019] The first preheater PH1 is connected to the first solenoid valve SV1 and is used to preheat the gas to the working temperature.

[0020] The first pressure reducing valve RV1 is connected to the first preheater PH1 and is used to output the pressure-reduced gas.

[0021] The second pressure sensor P2 is located at the first pressure reducing valve RV1 and is electrically connected to the power supply and control module, and is used to detect the pressure of the output gas.

[0022] As a preferred technical solution, in the dual-path pressure reducing sensing unit, the two branches have the same structure, and the other branch includes a second solenoid valve SV2, a second preheater PH2 and a second pressure reducing valve RV2 connected in sequence, and a third pressure sensor P3 located in the second pressure reducing valve RV2.

[0023] As a preferred technical solution, the first pressure reducing valve RV1 is a rubber diaphragm reverse pressure reducing valve.

[0024] As a preferred technical solution, the dual-thrust module includes:

[0025] Thrust A and thruster B, which serves as a cold backup for each other;

[0026] The third solenoid valve SV3 is connected at one end to the dual-path pressure regulating module and at the other end to the thruster A.

[0027] The fourth solenoid valve SV4 is connected at one end to the dual-path pressure regulating module and at the other end to the thruster B.

[0028] As a preferred technical solution, the power supply and control module includes:

[0029] The control unit includes a CAN control interface and a sensor interface, wherein the sensor interface is electrically connected to the pressure sensor in the dual-channel pressure regulation module.

[0030] Power supply unit A and power supply unit B are respectively connected to the satellite's onboard primary power supply and the control unit, and are also connected to the dual-thrust module through a drive power line, generating an electric arc for the ammonia propellant through arc discharge between the anode and cathode.

[0031] As a preferred technical solution, the control module is housed within an aluminum alloy casing.

[0032] As a preferred technical solution, the control unit is connected to the dual-channel pressure regulation module.

[0033] As a preferred technical solution, the storage tank is a titanium alloy metal gas cylinder.

[0034] Compared with the prior art, the present invention has at least one of the following beneficial effects:

[0035] (1) Reliable Redundant Architecture: The pressure regulation module of this utility model is designed with two independent channels. Each channel includes a solenoid valve, a preheater, a pressure reducing valve, and a pressure sensor. When one channel fails, it can switch to the other channel, reducing the system failure rate. The miniature rubber diaphragm reverse pressure reducing valve ensures small output pressure fluctuations. In conjunction with the pressure sensor, it monitors the upstream pressure of the thruster in real time, ensuring the stability of the thrust output. The two thrusters serve as cold backups for each other. The gas input is independently controlled by the solenoid valve. Even if either thruster fails, the track control capability can still be maintained, improving the success rate of track climbing missions and solving the problem of low reliability of existing resistance-heated thrusters.

[0036] (2) Lightweight and low cost: This utility model uses titanium alloy metal gas cylinders as storage tanks, which are lighter than traditional steel storage tanks. At the same time, it meets the safety factor requirements under 2MPa pressure conditions. The liquid ammonia storage method reduces the storage tank volume compared to the gas working fluid, reducing the weight and volume of the overall propulsion system to ≤1U, effectively releasing space for micro-nano satellite payloads and reducing satellite launch costs.

[0037] (3) Status monitoring: The present invention is equipped with a pressure sensor P1 in the storage tank to monitor the ammonia balance in real time. The pressure sensors P2 and P3 of the pressure regulation module monitor the upstream pressure. The data is transmitted to the control module through the CAN bus. It can be integrated with the existing spaceborne monitoring system to realize status monitoring. Attached Figure Description

[0038] Figure 1 This is a schematic diagram of the satellite electric propulsion system in the embodiment. Detailed Implementation

[0039] The technical solutions of the present utility model will be clearly and completely described below with reference to the accompanying drawings of the embodiments. Obviously, the described embodiments are only some, not all, of the embodiments of the present utility model. Based on the embodiments of the present utility model, all other embodiments obtained by those skilled in the art without creative effort should fall within the protection scope of the present utility model.

[0040] In the description of this utility model, it should be noted that the terms "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings, or the orientation or positional relationship commonly used when the utility model product is in use. They are only for the convenience of describing this utility model and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this utility model. For example, "horizontal" simply means that its direction is more horizontal than "vertical," and does not mean that the structure must be completely horizontal, but can be slightly tilted.

[0041] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this application, "multiple" means two or more, unless otherwise explicitly specified.

[0042] To address the problems existing in the prior art, this embodiment provides a satellite electric propulsion system, see [link to previous document]. Figure 1The system mainly consists of a storage tank, a power supply and control module, a dual-path pressure regulation module, and a dual-thrust module. The storage tank is a spherical structure welded entirely from TC4 titanium alloy, with an inner diameter of 226mm, and stores liquid ammonia. The power supply and control module is integrally machined from aluminum alloy and includes a control module, power supply A, and power supply B. The dual-path pressure regulation module is connected to the storage tank via piping and includes an add / release valve and a dual-path pressure reduction sensor unit. The dual-thrust module is perpendicular to the bottom surface and positioned past the center of mass, and includes thruster A and thruster B. Each module will be described in detail below.

[0043] (1) Storage tank module.

[0044] The storage tank is a full titanium alloy metal cylinder, consisting of upper and lower hemispheres (including cylindrical sections) and a gas-liquid inlet, and is fixed to the satellite platform using a strap-on mounting structure. Its internal storage density is 0.58 g / cm³. 3 The liquid ammonia gas has a pressure ≤2MPa at a high temperature of 45℃. A filler / release valve 1 (FDV1) is installed at the gas-liquid port of the storage tank to control the filling, release and flow rate of ammonia gas; a pressure sensor 1 (P1) is installed on the side wall to monitor the remaining ammonia gas in the storage tank in real time and transmit the signal to the power supply and control module.

[0045] (2) Dual-channel pressure regulation module.

[0046] This module is connected to the outlet of the tank's fill / drain valve 1 via a pipeline and includes two symmetrical pressure-reducing sensing units. One branch sequentially houses a solenoid valve 1 (SV1), a preheater 1 (PH1), a pressure-reducing valve 1 (RV1), and a pressure sensor 2 (P2). The solenoid valve 1 controls the gas output from the tank; the preheater 1 heats the gas to its operating temperature; the pressure-reducing valve 1 uses a miniature rubber diaphragm reverse structure to stably reduce the high-pressure gas to the operating pressure; and the pressure sensor 2 monitors the upstream pressure of the thruster in real time and feeds it back to the control module. The other branch has the same structure as the first branch, including a solenoid valve 2 (SV2), a preheater 2 (PH2), a pressure-reducing valve 2 (RV2), and a pressure sensor 3 (P3). This dual-redundancy design ensures automatic switching in case of single-path failure, maintaining system pressure stability.

[0047] (3) Dual thrust module.

[0048] The thrust module comprises thruster A and thruster B, both oriented perpendicular to the base and passing through the satellite's center of mass. Each unit can provide 30 mN of thrust and serves as a cold backup for the others. Each thruster inlet is connected to solenoid valve 3 (SV3) and solenoid valve 4 (SV4), respectively. These solenoid valves control the gas input to allow for independent operation or cold backup switching of the thruster. Ammonia gas, after being depressurized by the pressure regulating module, enters the thruster through the solenoid valves. Under the action of an electric arc discharge between the cathode and anode, it is heated to a high-temperature plasma, which is then ejected through the anode to generate thrust.

[0049] (4) Power supply and control module.

[0050] The module features a machined aluminum alloy casing and integrates a control unit, power supply A, and power supply B. The control unit connects to the power module via side lugs and includes two CAN control interfaces for satellite communication and one peripheral interface for connecting solenoid valves, pressure sensors, etc., connected to the power module via wires. Within the power unit, power supplies A and B are connected to the satellite's primary power supply via connectors. Two sets of wires serve as the drive power lines for the thrust module, generating a high-voltage, high-temperature electric arc to heat the ammonia working fluid.

[0051] The control unit is implemented using a commercially available existing spaceborne control unit.

[0052] The working principle of this system is as follows:

[0053] 1. Working fluid storage and release. Liquid ammonia in the storage tank vaporizes at room temperature to produce 2MPa high-pressure gas, and the output flow is controlled by the vent valve 1. Pressure sensor 1 monitors the tank pressure in real time, and sends an alarm signal to the control module when the pressure is lower than 1.8MPa, indicating that the working fluid is insufficient.

[0054] 2. Pressure regulation. High-pressure gas enters the dual-path pressure regulation module through the expansion and contraction valve 1.

[0055] When the system is working normally, solenoid valve 1 and solenoid valve 2 open simultaneously. The gas is heated to 60℃±5℃ by preheaters 1 and 2 respectively, and then reduced to 0.5MPa working pressure by pressure reducing valves 1 and 2. Pressure sensors 2 and 3 monitor the output pressure in real time and feed it back to the control module to ensure that the pressure fluctuation is relatively stable.

[0056] If a branch circuit fails, such as solenoid valve 1 malfunctioning, the control module will automatically close the solenoid valve in that branch circuit and open solenoid valve 2 in another branch circuit. The second branch circuit will then independently complete the pressure reduction process, ensuring continuous operation of the system.

[0057] 3. Thrust generation and control.

[0058] After depressurization, the ammonia gas enters thrusters A and B via solenoid valves 3 and 4. The power supply and control module receives the satellite's orbit-climbing command via the CAN bus and controls power supplies A and B to output high-voltage electricity, generating an electric arc between the cathode and anode of the thrusters. This arc heats the ammonia gas to a plasma state, which is then ejected through the anode nozzle, generating 30 mN of thrust. The control module adjusts the arc power to achieve a rapid response in thrust output. Combined with the dual-thruster cold backup design, this ensures the successful execution of the orbit-climbing mission.

[0059] 4. Redundancy and monitoring.

[0060] Power supplies A and B serve as cold backups for each other. When a single power supply module fails, the control module automatically switches to the other power supply module to maintain continuous arc discharge. Pressure sensors 1, 2, and 3 monitor the tank level and the upstream pressure of the thruster in real time. The data is transmitted to the control module via the CAN bus. When the pressure is abnormal, such as exceeding 2 MPa or falling below 0.4 MPa, the control module automatically activates the protection mechanism and sends an alarm signal to the satellite, supporting manual intervention from the ground.

[0061] This invention generates a high-voltage, high-temperature electric arc through a power supply and control module. The ammonia propellant is heated by the arc discharge between the cathode and anode, significantly improving the electrothermal conversion efficiency compared to traditional resistance-heated thrusters. The ammonia propellant, after being heated by the arc, has a higher specific impulse, increasing its utilization efficiency and extending the satellite's on-orbit operating time. Through modular integrated design, dual-redundancy architecture, and arc heating, a lightweight, low-cost, and highly reliable micro / nano satellite propulsion system is achieved. This significantly improves propellant utilization efficiency and control precision compared to traditional chemical thrusters and resistance-heated thrusters, making it suitable for large-scale deployment in low-Earth orbit satellite constellations.

[0062] It should be noted that the control unit in this utility model is an existing product, and this utility model does not involve any methodological improvements such as algorithms or control methods.

[0063] The above description is merely a specific embodiment of this utility model, but the protection scope of this utility model is not limited thereto. Any person skilled in the art can easily conceive of various equivalent modifications or substitutions within the technical scope disclosed in this utility model, and these modifications or substitutions should all be covered within the protection scope of this utility model. Therefore, the protection scope of this utility model should be determined by the scope of the claims.

Claims

1. A satellite electric propulsion system, characterized in that, include: The storage tank contains liquid ammonia gas. The power supply and control module is used to generate a high-voltage, high-temperature electric arc and control the operation of the thruster. A dual-path pressure regulation module, connected to the storage tank, is used to convert the high-pressure gas in the storage tank into working pressure gas and output it to the thrust module in two paths; The dual-thrust module, connected to the dual-path pressure regulation module, is used to generate thrust by heating ammonia propellant through electric arc discharge.

2. The satellite electric propulsion system according to claim 1, characterized in that, The dual-channel pressure regulation module includes: The FDV1 valve is connected at one end to the storage tank and is used to control the injection and release of ammonia. A dual-path pressure reducing sensor unit is connected to the other end of the FDV1 valve and is used to convert the high-pressure gas in the tank into gas at the working pressure. The first pressure sensor P1 is connected to the storage tank.

3. A satellite electric propulsion system according to claim 2, characterized in that, In the aforementioned dual-channel pressure reduction sensing unit, one branch includes: The first solenoid valve SV1 is connected to the feed / discharge valve FDV1; The first preheater PH1 is connected to the first solenoid valve SV1 and is used to preheat the gas to the working temperature. The first pressure reducing valve RV1 is connected to the first preheater PH1 and is used to output the pressure-reduced gas. The second pressure sensor P2 is located at the first pressure reducing valve RV1 and is electrically connected to the power supply and control module, and is used to detect the pressure of the output gas.

4. A satellite electric propulsion system according to claim 3, characterized in that, In the dual-path pressure reducing sensing unit, the two branches have the same structure. The other branch includes a second solenoid valve SV2, a second preheater PH2, and a second pressure reducing valve RV2 connected in sequence, as well as a third pressure sensor P3 located in the second pressure reducing valve RV2.

5. A satellite electric propulsion system according to claim 3, characterized in that, The first pressure reducing valve RV1 is a rubber diaphragm reverse pressure reducing valve.

6. A satellite electric propulsion system according to claim 1, characterized in that, The dual-thrust module includes: Thrust A and thruster B, which serves as a cold backup for each other; The third solenoid valve SV3 is connected at one end to the dual-path pressure regulating module and at the other end to the thruster A. The fourth solenoid valve SV4 is connected at one end to the dual-path pressure regulating module and at the other end to the thruster B.

7. A satellite electric propulsion system according to claim 1, characterized in that, The power supply and control module includes: The control unit includes a CAN control interface and a sensor interface, wherein the sensor interface is electrically connected to the pressure sensor in the dual-channel pressure regulation module. Power supply unit A and power supply unit B are respectively connected to the satellite's onboard primary power supply and the control unit, and are also connected to the dual-thrust module through a drive power line, generating an electric arc for the ammonia propellant through arc discharge between the anode and cathode.

8. A satellite electric propulsion system according to claim 7, characterized in that, The control module is housed within an aluminum alloy casing.

9. A satellite electric propulsion system according to claim 7, characterized in that, The control unit is connected to the dual-channel pressure regulation module.

10. A satellite electric propulsion system according to claim 1, characterized in that, The storage tank is a titanium alloy metal gas cylinder.

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