Hall plasma propulsion device for spacecrafts

Abstract

This invention relates to a Hall plasma propulsion device for spacecraft, belonging to the field of spacecraft propulsion technology. It solves the problems of existing Hall thrusters being large in size, high in power, and unable to meet the needs of low-Earth orbit satellites and micro / nano satellites. The invention includes a base with a central permanent magnet fixed at its center. A central magnetic conductive component is inserted into the end face of the base. An inner magnetic screen, a ceramic sleeve, and an outer magnetic screen are sequentially fitted onto the outside and snapped into the base. An annular groove is formed at the distal end of the ceramic sleeve to create a discharge channel. An anode component is fixed within the channel, and a gas working fluid inlet is provided. An inner magnetic pole is fixed inside the distal end of the ceramic sleeve, and an outer magnetic pole is fitted outside. A shell is snapped between the outer magnetic pole and the base. Multiple sets of peripheral permanent magnets are evenly distributed circumferentially on the outer side of the shell. The central permanent magnet has opposite magnetic poles to the peripheral permanent magnets, forming a lens-shaped magnetic field convex to the anode within the discharge channel. This magnetic field is orthogonal to the anode electric field, causing Hall drift electrons to ionize the working fluid, and the accelerated ejection of ions generates thrust. This invention is suitable for orbit maintenance and attitude control of low-Earth orbit satellite constellations and micro / nano satellites.

Description

Technical Field

[0001] This invention belongs to the field of spacecraft propulsion technology, specifically relating to a Hall plasma propulsion device for spacecraft. Background Technology

[0002] With the rapid development of low-Earth orbit satellite constellations, microsatellites, and deep space exploration missions, the demand for efficient, low-power, miniaturized electric propulsion systems with high specific impulse is becoming increasingly urgent. Traditional chemical propulsion methods, due to their high propellant consumption, low specific impulse, and system complexity, are unable to meet the requirements of current space missions for long lifespan, high-precision orbit control, and maximum payload. Plasma propulsion technology, especially Hall effect-based propulsion devices, has become one of the most researched and widely used electric propulsion solutions in the international aerospace field due to its compact structure, high thrust density, excellent specific impulse, long service life, and high thrust adjustment precision.

[0003] Among existing Hall effect plasma propulsion devices, various configurations have been developed, including annular Hall thrusters, cylindrical Hall thrusters, tangential field Hall thrusters, permanent magnet Hall thrusters, and channelless Hall thrusters. A typical device forms an annular cylindrical plasma discharge channel using two coaxially arranged ceramic sleeves of different radii. The magnetic field is radially distributed with an axial component, while the electric field is axially distributed, forming an orthogonal electromagnetic field with the radial magnetic field component. Electrons emitted from the cathode enter the channel and undergo Hall drift motion, colliding with the injected atomic working fluid (such as xenon gas). This results in avalanche ionization, generating a large number of ions and electrons. The ions are accelerated and ejected under the influence of a self-consistent axial electric field, thus generating continuous thrust. This type of device has been widely used in space missions such as satellite orbit control, repositioning, attitude control, position holding, and deep space exploration, significantly saving propellant mass, reducing launch costs, and improving the payload capacity of spacecraft.

[0004] However, with the explosive growth of low-Earth orbit satellite constellation networking and micro / nano satellite technology, space missions are placing more stringent demands on electric propulsion systems: power output is typically in the hundreds of watts or even lower, size is strictly limited, and both high specific impulse and high reliability are required. Existing conventional Hall thrusters are mostly designed for large satellites or medium-to-high power missions, making it difficult to directly meet the on-orbit requirements of miniaturization, low power consumption, and high integration. Therefore, there is an urgent need to develop a Hall electric propulsion device that is more compact, has lower power output, still maintains a high specific impulse, and is easily integrated with microsatellite platforms for actual on-orbit applications, in order to fill the current technological gap and promote the lightweight and intelligent development of spacecraft propulsion systems. Summary of the Invention

[0005] The present invention addresses the problem that existing Hall effect plasma propulsion devices are too large and have too high power, making it difficult to meet the on-orbit requirements of low-Earth orbit satellite constellations and micro-nano satellites for miniaturized, low-power, and high-specific-impulse electric propulsion systems. The invention provides a Hall plasma propulsion device for spacecraft.

[0006] The Hall plasma propulsion device for spacecraft described in this invention includes: a disk-shaped inner magnetic pole, a cylindrical ceramic sleeve, an annular outer magnetic pole, a cylindrical outer magnetic screen, a base, a cylindrical inner magnetic screen, a central permanent magnet, peripheral permanent magnets, and a central magnetic conductive component.

[0007] One end of the central permanent magnet is fixed to the center of the base; the other end of the central permanent magnet is connected to a central magnetic conductive component.

[0008] One end of the cylindrical inner magnetic screen, the cylindrical ceramic sleeve, and the cylindrical outer magnetic screen are snapped onto the upper surface of the base and are coaxially sleeved on the outside of the central magnetic conductive component in sequence.

[0009] An annular groove is coaxially formed on the end face of the other end of the ceramic sleeve to form an annular discharge channel. An anode component is fixed inside the discharge channel, and a working gas inlet is provided.

[0010] The inner side of the ceramic sleeve is fitted with a disc-shaped inner magnetic pole, and the outer side is fitted with an annular outer magnetic pole; and both the disc-shaped inner magnetic pole and the annular outer magnetic pole are located at the other end of the cylindrical ceramic sleeve.

[0011] A cylindrical outer shell is snapped between the annular outer magnetic pole and the base, and the cylindrical outer shell is disposed on the outside of the cylindrical outer magnetic screen;

[0012] The peripheral permanent magnets are evenly distributed along the circumference of the cylindrical outer shell;

[0013] The central permanent magnet has magnetic poles that are opposite to those of the peripheral permanent magnets, forming a lens-shaped magnetic field that bends towards the anode within the discharge channel. This magnetic field is orthogonal to the electric field applied to the anode, causing electrons emitted from the cathode to undergo Hall drift and ionize the working gas. The ions generated by the ionization of the working gas are accelerated and ejected under the influence of the electric and magnetic fields, thereby generating thrust.

[0014] This invention employs a central and peripheral permanent magnet with opposite magnetic poles. It utilizes the magnetic circuit structure of the magnetically conductive material to form a lens-shaped magnetic field, a compact straight-cylinder annular discharge channel, and a low-voltage, low-current operating mode. Electrons are effectively confined at the discharge channel outlet and undergo sufficient collisions with the working gas, resulting in ionization. The generated ions are accelerated and ejected under the action of orthogonal electromagnetic fields. Thus, under the constraints of low power (tens of watts) and small size (discharge channel outer diameter not exceeding 24 mm), it achieves high specific impulse (up to 750 s), small plume divergence angle (≤15°), high ionization efficiency, and long working life. This significantly reduces spacecraft power consumption and launch costs, and meets the stringent requirements of low-Earth orbit satellite constellations and micro / nano satellites for orbit maintenance, attitude control, and maneuver propulsion. Attached Figure Description

[0015] Figure 1 This is a cross-sectional view of the miniaturized Hall electric propulsion device for spacecraft described in this invention;

[0016] Figure 2 An enlarged cross-sectional view showing the position of the central permanent magnet;

[0017] Figure 3 An enlarged cross-sectional view showing the position of the peripheral permanent magnet;

[0018] Figure 4 This is a partially enlarged cross-sectional view of the discharge channel;

[0019] Figure 5 The simulation results of the magnetic induction intensity along the axial direction of the discharge channel are shown in the figure.

[0020] Figure 6 This is a simulation diagram of the lens-type magnetic field configuration at the discharge channel exit.

[0021] In the diagram, 1 is the disc-shaped inner magnetic pole, 2 is the cylindrical ceramic sleeve, 3 is the annular outer magnetic pole, 4 is the cylindrical outer magnetic screen, 5 is the cylindrical outer shell, 6 is the base, 7 is the cylindrical inner magnetic screen, 8 is the central permanent magnet, 10 is the cylindrical permanent magnet protective shell, 11 is the peripheral permanent magnet, and 12 is the central magnetic conductive component. Detailed Implementation

[0022] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of them. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention. It should be noted that, unless otherwise specified, the embodiments and features in the embodiments of the present invention can be combined with each other.

[0023] Specific implementation method one: Refer to Figures 1 to 4This embodiment describes a miniaturized Hall electric propulsion device for spacecraft, comprising: a disc-shaped inner magnetic pole 1, a cylindrical ceramic sleeve 2, an annular outer magnetic pole 3, a cylindrical outer magnetic screen 4, a base 6, a cylindrical inner magnetic screen 7, a central permanent magnet 8, a peripheral permanent magnet 11, and a central magnetic guiding component 12.

[0024] One end of the central permanent magnet 8 is fixed to the center of the base 6; the other end of the central permanent magnet 8 is connected to the central magnetic conductive component 12.

[0025] One end of the cylindrical inner magnetic screen 7, the cylindrical ceramic sleeve 2, and the cylindrical outer magnetic screen 4 are snapped onto the upper surface of the base 6, and are coaxially sleeved on the outside of the central magnetic conductive component in sequence.

[0026] An annular groove is coaxially formed on the end face of the other end of the ceramic sleeve 2 to form an annular discharge channel. An anode component is fixed in the discharge channel, and a working gas inlet is provided.

[0027] The inner side of the ceramic sleeve 2 is fitted with a disc-shaped inner magnetic pole 1, and the outer side is fitted with an annular outer magnetic pole 3; and the disc-shaped inner magnetic pole 1 and the annular outer magnetic pole 3 are both located at the other end of the cylindrical ceramic sleeve 2.

[0028] A cylindrical outer shell 5 is snapped between the annular outer magnetic pole 3 and the base 6, and the cylindrical outer shell 5 is disposed on the outside of the cylindrical outer magnetic screen 4;

[0029] The peripheral permanent magnets 11 are evenly distributed along the circumference of the cylindrical outer shell 5;

[0030] The central permanent magnet 8 and the peripheral permanent magnet 11 have opposite magnetic pole directions, forming a lens-shaped magnetic field that bends towards the anode within the discharge channel. This magnetic field is orthogonal to the electric field applied by the anode, causing the electrons emitted by the cathode to undergo Hall drift and ionize the working gas. The ions generated by the ionization of the working gas are accelerated and ejected under the action of the electric and magnetic fields, thereby generating thrust.

[0031] Furthermore, in this invention, the base 6 has a circular protrusion at its center and steps that decrease sequentially outward. The circular protrusion is used to install the central permanent magnet 8, and the steps are used from the inside out to engage the cylindrical inner magnetic screen 7, the cylindrical outer magnetic screen 4, and the cylindrical outer shell 5, respectively.

[0032] Furthermore, in this invention, the central magnetic conductive component 12, the cylindrical inner magnetic screen 7, the cylindrical outer magnetic screen 4, the disc-shaped inner magnetic pole 1, and the annular outer magnetic pole 3 are all made of magnetic conductive material.

[0033] Furthermore, in this invention, both the central permanent magnet 8 and the peripheral permanent magnet 11 are samarium cobalt permanent magnets.

[0034] Furthermore, in this invention, the central permanent magnet 8 is a cylindrical permanent magnet with a diameter of 7mm±1mm and a height of 6.5mm±1mm, and the peripheral permanent magnet 11 is a cylindrical permanent magnet with a diameter of 3mm±1mm.

[0035] Furthermore, in this invention, there are three or six sets of permanent magnets 11 evenly spaced on the periphery of the cylindrical shell 5, and the axis of the peripheral permanent magnets 11 is parallel to the axis of the cylindrical shell 5.

[0036] Furthermore, in this invention, the thickness of both the cylindrical inner magnetic screen 7 and the cylindrical outer magnetic screen 4 is 1.2±0.02mm.

[0037] Furthermore, in this invention, the flow rate of the working gas entering the discharge channel is 2 sccm to 5 sccm.

[0038] Furthermore, in this invention, the peripheral permanent magnet 11 is placed inside the cylindrical permanent magnet protective shell 10, and the cylindrical permanent magnet protective shell 10 is disposed between the annular outer magnetic pole 3 and the base 6.

[0039] Furthermore, in this invention, the anode element is implemented using a gas distributor.

[0040] This invention further reduces the size and operating power of Hall thrusters, solving the problems of decreased ionization efficiency, reduced thrust performance, and increased plume divergence angle in plasma propulsion devices under constraints of small size and low power. This meets the on-orbit propulsion and attitude control requirements of low-Earth orbit satellite constellations and micro / nano satellites. Furthermore, while ensuring heat dissipation, a compact structural design reduces the size of the Hall thruster. It abandons the traditional method of using excitation coils to generate a magnetic field, instead employing permanent magnets as the magnetic field source, thus reducing power consumption and expenses. Simultaneously, while ensuring the thruster can ignite normally and discharge continuously, a discharge voltage of 200V-300V and a low current of 0.2A-0.3A are applied to the anode of the device to reduce power requirements. To address the issues of decreased ionization efficiency and thrust performance caused by the size constraints of the discharge channel, the width and length of the discharge channel are optimized, two magnetic screens are used, the distance between the magnetic screens and the magnetic poles is optimized, and the magnetic field gradient distribution within the discharge channel is optimized. This reduces the collision frequency of electrons / ions with the discharge channel, minimizes wall energy loss and sputtering corrosion, ensures the utilization rate of the magnetic field and working fluid, and improves ionization efficiency. To reduce the plume divergence angle and achieve better plume focusing, a lens-type magnetic field configuration is designed, where magnetic field lines are symmetrically distributed along the center of the discharge channel, and the magnetic field lines inside the discharge channel convex towards the anode. The magnetic field lines exhibit a Gaussian or bell-shaped curve. This achieves miniaturization, low power consumption, and high specific impulse of the spacecraft electric propulsion device, which can significantly improve the satellite's on-orbit lifespan and payload, reduce launch costs, and is suitable for low-Earth orbit satellite constellations, micro-nano satellites, and other space platforms.

[0041] To address the requirements of lightweight and low-power magnetic field sources in spacecraft propulsion systems, samarium cobalt permanent magnets are used as the magnetic field source. Figure 2 and Figure 3 The enlarged view illustrates the selection of magnets and their placement within the thruster: The optimized selection of samarium cobalt permanent magnets demonstrates their superior thermal stability compared to neodymium iron boron (NdFeB) magnets (700-800℃ Curie point), while also exhibiting a higher magnetic energy product than AlNiCo magnets (AlNiCo), generating a stronger magnetic field within the same volume. Finite element simulation was used to optimize the magnet size and arrangement. Cylindrical samarium cobalt permanent magnets with an 8mm diameter are positioned rearward along the thruster's central axis. Three or six sets of circumferentially distributed cylindrical samarium cobalt permanent magnets with a diameter of 4mm, placed parallel to the central axis, serve as auxiliary magnets, achieving optimal magnetic field distribution and ensuring adequate confinement of electrons. The use of permanent magnets eliminates the need for and consumption of electrical energy for the magnetic field source, further reducing the thruster's operating energy consumption.

[0042] Combination Figure 4 This document describes the optimized design of the discharge channel. To address the space constraints and efficiency improvements of miniaturized electric propulsion devices for aerospace applications, the discharge channel structure was optimized. The discharge channel of this device is a straight annular shape, with its cross-sectional area matched to the working gas flow rate. For low-power annular channel plasma accelerators, size limitations restrict the magnet volume and the magnetic field strength provided at the discharge channel outlet. Reducing the width of the discharge channel helps achieve effective electron confinement under low magnetic field strength. A 4mm width was chosen. Narrow discharge channels tend to increase the sheath ratio (sum of inner and outer sheath thicknesses / discharge channel width), leading to significant wall scattering. Wall loss can be reduced by appropriately shortening the discharge channel length. However, an excessively short discharge channel length results in insufficient ionization of working gas atoms. Through optimization, the width-to-length ratio of the discharge channel was set to 0.3, and the designed channel length was 13mm to balance ionization efficiency and wall loss rate. A gas supply flow rate of 2 sccm to 5 sccm was used, with a corresponding channel cross-sectional area of ​​250mm². 2 That is, the inner diameter is 16mm and the outer diameter is 24mm. For low-power plasma accelerators, the plasma density is low, and an extended channel would cause the ionization region to shift outward, resulting in a decrease in working efficiency. The channel is set as a straight cylindrical annular shape with parallel inner and outer walls and no expansion design.

[0043] Combination Figure 5 and Figure 6 Explanation of the optimized design of the magnetic circuit:

[0044] To further improve the propulsion efficiency of spacecraft electric propulsion systems, the magnetic circuit design is optimized. The inner cylindrical magnetic screen 7 is a cylinder with an outward-extending edge at the bottom, and the outer cylindrical magnetic screen 4 is a cylinder with an inward-extending edge at the bottom. The two magnetic screens face each other, with the inner magnetic screen having an outward-extending edge and the outer magnetic screen having an inward-extending edge. Both edges are located within the cylindrical ceramic sleeve 2.

[0045] The connection end to the base is positioned opposite to it, and the thickness of both the inner and outer magnetic screens is uniformly designed to be 1.2mm to balance the overall weight of the device and the magnetic conductivity of the magnetic circuit. The distance between the magnetic poles and the magnetic screen is optimized to improve the gradient of magnetic induction intensity in the axial direction inside the discharge channel, such as... Figure 5 Simulation results show that a larger magnetic flux density gradient can concentrate the ionization region, improve ionization efficiency, suppress axial diffusion of electrons, and enhance the directionality and flatness of ion acceleration, thereby increasing the specific impulse of the device. Optimizing the magnetic circuit configuration and magnet placement, and designing a lens-type magnetic field configuration, such as... Figure 6 Simulation results show that at the discharge outlet, the lens-type magnetic field configuration exhibits magnetic field lines convex towards the anode.

[0046] Low-cost design:

[0047] To address the low-cost and rapid production requirements of miniaturized electric propulsion devices for aerospace applications, a comprehensive cost control system was established. At the design stage, the DFMA (Design for Manufacturing and Assembly) approach was adopted to control costs from the source and simplify assembly. At the material selection stage, all materials were domestically produced, with priority given to national standard grades. At the procurement stage, standard parts were used to leverage a mature supply chain to reduce costs. At the machining stage, tolerances for non-critical dimensions were relaxed to reduce machining difficulty. At the assembly stage, modular design was adopted to simplify the assembly process. Secondly, key material substitution schemes were implemented, including replacing precision ceramics with BN ceramics to reduce costs; replacing IJ22 with DT4C electrical pure iron as the magnetic shield material to reduce costs; replacing NdFeB with Samarium Cobalt permanent magnets to reduce costs; replacing special stainless steel with 316L stainless steel for gas distributors and housings to reduce costs; and replacing imported seals with domestically produced standard seals to reduce costs. Furthermore, in terms of process simplification design, complex curved surfaces were simplified into simple, machinable curved surfaces, parts with similar functions were merged to reduce the number of parts, standard threads and standard hole spacing were used to facilitate assembly, and an assembly guide structure was designed to reduce assembly difficulty and cost. Specific component material lists are as follows:

[0048] Table 1 Specific Component Materials List

[0049]

[0050] While the invention has been described herein with reference to specific embodiments, it should be understood that these embodiments are merely examples of the principles and applications of the invention. Therefore, it should be understood that many modifications can be made to the exemplary embodiments, and other arrangements can be designed without departing from the spirit and scope of the invention as defined by the appended claims. It should be understood that different dependent claims and features described herein can be combined in ways different from those described in the original claims. It is also understood that features described in conjunction with individual embodiments can be used in other described embodiments.

Claims

1. A Hall plasma propulsion device for spacecraft, characterized in that, include: The components include a disc-shaped inner magnetic pole (1), a cylindrical ceramic sleeve (2), an annular outer magnetic pole (3), a cylindrical outer magnetic screen (4), a base (6), a cylindrical inner magnetic screen (7), a central permanent magnet (8), an outer permanent magnet (11), and a central magnetic conductive component (12). One end of the central permanent magnet (8) is fixed to the center of the base (6); the other end of the central permanent magnet (8) is connected to a central magnetic conductive component (12). One end of the cylindrical inner magnetic screen (7), the cylindrical ceramic sleeve (2) and the cylindrical outer magnetic screen (4) are snapped onto the upper surface of the base (6) and are coaxially sleeved on the outside of the central magnetic conductive component in sequence. The ceramic sleeve (2) has an annular groove coaxially formed on the end face of the other end to form an annular discharge channel. An anode component is fixed in the discharge channel and a working gas inlet is provided. The inner side of the ceramic sleeve (2) is fitted with a disc-shaped inner magnetic pole (1), and the outer side is fitted with an annular outer magnetic pole (3); and the disc-shaped inner magnetic pole (1) and the annular outer magnetic pole (3) are both located at the other end of the cylindrical ceramic sleeve (2); A cylindrical outer shell (5) is snapped between the annular outer magnetic pole (3) and the base (6), and the cylindrical outer shell (5) is disposed on the outside of the cylindrical outer magnetic screen (4); The peripheral permanent magnets (11) are evenly distributed along the circumference of the cylindrical shell (5); The central permanent magnet (8) has the opposite magnetic pole direction to the peripheral permanent magnet (11).

2. The Hall plasma propulsion device for spacecraft according to claim 1, characterized in that, The base (6) has a circular protrusion at its center and steps that decrease in a radial direction. The circular protrusion is used to install the central permanent magnet (8), and the steps are used to engage the cylindrical inner magnetic screen (7), the cylindrical outer magnetic screen (4), and the cylindrical outer shell (5) from the inside out.

3. The Hall plasma propulsion device for spacecraft according to claim 1, characterized in that, The central magnetic conductive component (12), the cylindrical inner magnetic screen (7), the cylindrical outer magnetic screen (4), the disc-shaped inner magnetic pole (1), and the annular outer magnetic pole (3) are all made of magnetic conductive material.

4. The Hall plasma propulsion device for spacecraft according to claim 3, characterized in that, Both the central permanent magnet (8) and the peripheral permanent magnet (11) are made of samarium cobalt permanent magnets.

5. The Hall plasma propulsion device for spacecraft according to claim 1, characterized in that, The central permanent magnet (8) is a cylindrical permanent magnet with a diameter of 7mm±1mm and a height of 6.5mm±1mm, and the outer permanent magnet (11) is a cylindrical permanent magnet with a diameter of 3mm±1mm.

6. The Hall plasma propulsion device for spacecraft according to claim 5, characterized in that, The permanent magnets (11) are arranged at equal intervals on the periphery of the cylindrical shell (5) in three or six groups, and the central axis of the peripheral permanent magnets (11) is parallel to the central axis of the cylindrical shell (5).

7. The Hall plasma propulsion device for spacecraft according to claim 1, characterized in that, The thickness of both the cylindrical inner magnetic screen (7) and the cylindrical outer magnetic screen (4) is 1.2 ± 0.02 mm.

8. The Hall plasma propulsion device for spacecraft according to claim 1 or 2, characterized in that, The flow rate of the working gas into the discharge channel is 2 sccm to 5 sccm.

9. The Hall plasma propulsion device for spacecraft according to claim 1 or 2, characterized in that, The peripheral permanent magnet (11) is placed inside the cylindrical permanent magnet protective shell (10), which is located between the annular outer magnetic pole (3) and the base (6).

10. The Hall plasma propulsion device for spacecraft according to claim 1, characterized in that, The anode component is achieved using a gas distributor.

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