Satellite having electrostatic electric propulsion system

Abstract

This invention belongs to the field of aerospace technology and relates to a satellite with an electrostatic electric propulsion system, comprising: a satellite body; one or more propulsion systems, wherein at least one of the propulsion systems employs an inertial electrostatic confinement ion thruster, the inertial electrostatic confinement ion thruster comprising: a gas supply device; a magnetic guiding device, the magnetic guiding device comprising an electromagnetic coil and a magnetic guiding nozzle; an anode, the anode being a spherical shell; a cathode, the cathode body being generally cage-like, and the cathode having a first interface connected to the gas supply device and a second interface connected to the magnetic guiding nozzle; the cathode is housed inside the anode, and the gas supply device is used to introduce gas located outside the anode into the cathode; a portion of the magnetic guiding nozzle is located outside the anode. This invention employs the novel ionization method of inertial electrostatic confinement, enabling adjustments to the satellite's attitude and orbit.

Description

[0001] This application is a divisional application, the parent application of which is Chinese patent application No. 202310388646.0 filed on April 13, 2023, entitled "A Satellite with an Electrostatic Electric Propulsion System". Technical Field

[0002] This invention belongs to the field of spaceflight technology, and specifically relates to a satellite with an electrostatic electric propulsion system. Background Technology

[0003] Spaceflight, also known as space travel, space exploration, or spaceflight, refers to the general term for activities involving the entry, exploration, development, and utilization of outer space beyond Earth's atmosphere and celestial bodies beyond Earth. The spacecraft used in spaceflight include spacecraft and satellites.

[0004] Satellites possess their own propulsion systems because after entering their predetermined orbit via launch vehicle or other means, they still need to maintain their orbit, maintain precise positioning, and compensate for drag to complete application tasks such as network operation and precise Earth observation. Therefore, the satellite's propulsion system has a significant impact on improving its operational lifespan and ensuring its normal operation. For micro and nanosatellites, in addition to traditional chemical propulsion devices, electric propulsion devices can also be used. Electric propulsion is an advanced propulsion method that uses electrical energy to directly heat the propellant or uses electromagnetic effects to ionize and accelerate the propellant to obtain propulsion power. It has high specific impulse, thrust, and efficiency, and has broad application prospects in space missions such as orbit control of large spacecraft, deep space exploration, and interplanetary travel.

[0005] For example, Patent Document 1 proposes a satellite that simultaneously employs chemical and electric thrusters, and implements corresponding controls when an electric thruster failure is detected. For instance, it controls the combustion of the remaining electric thrusters near the orbital intersection point, and controls the combustion of the chemical thrusters at specific locations. Patent Document 1's solution is merely a reactive measure after an electric thruster failure; compared to a scenario where no failure occurs, it does not improve the lifespan of the electric thrusters or even the satellite.

[0006] Patent Document 1:

[0007] Application Publication Number: CN 106275506 A, Applicant: Boeing, Classification Number: B64G1 / 10, B64G1 / 24, B64G1 / 40, Invention Title: Effective Position Holding Design for Mixed-Fuel Systems in Response to Electric Thruster Failure.

[0008] Specifically, regarding the types of electric propulsion, the most widely used in orbit are ion electric propulsion systems and Hall thruster systems. Hall thrusters have a relatively simple structure and small size, with a lower specific impulse than ion thrusters, requiring a larger propellant loading, but they have greater thrust and shorter orbit change times. Ion thrusters have a relatively complex structure and larger size, with a higher specific impulse than Hall thrusters, requiring a smaller propellant loading, but they have lower thrust and longer orbit change times.

[0009] To address this, Patent Document 2 proposes a satellite platform compatible with both ion and Hall thruster configurations. The electric propulsion subsystem mainly comprises two thruster modules, a tank supply module, and a power supply module. This electric propulsion system can utilize both ion and Hall thrusters to perform satellite orbit changes and on-orbit position maintenance (north-south, east-west, and angular momentum unloading). It can also perform these tasks using either ion or Hall thrusters independently. When configuring ion or Hall thrusters independently, the unconfigured functional modules can be simply removed, without requiring adjustments to the layout of other instruments and equipment on the platform. However, this structure is still based on a combination of known thrusters and does not overcome the inherent limitations of each individual thruster.

[0010] Patent Document 2:

[0011] Application Publication Number: CN 115503984 A, Applicant: China Academy of Space Technology, Classification Number: B64G1 / 40, Invention Title: A Satellite Platform Compatible with Ion and Hall Electric Propulsion Configuration.

[0012] Therefore, further research on thrusters is still needed to provide better propulsion for satellites. Summary of the Invention

[0013] Based on the inventor's research and practical experience in this field, the following improved technical solutions are proposed.

[0014] A satellite, comprising:

[0015] ontology,

[0016] One or more propulsion systems, characterized in that at least one of the propulsion systems employs an inertial electrostatic confinement ion thruster, the inertial electrostatic confinement ion thruster comprising:

[0017] Gas supply device;

[0018] A magnetic guiding device, the magnetic guiding device comprising an electromagnetic coil and a magnetic guiding nozzle;

[0019] Anode, wherein the anode is a spherical shell;

[0020] The cathode has a cage-like body and is provided with a first interface for connecting to the gas supply device and a second interface for connecting to the magnetic nozzle.

[0021] The cathode is housed inside the anode, and the gas supply device is used to introduce gas located outside the anode into the cathode.

[0022] A portion of the magnetic nozzle is located outside the anode.

[0023] According to the present invention, the satellite can use electric propulsion to change its attitude or orbit. Specifically, an inertial electrostatic electric thruster is used. By creating a positive voltage at the anode and a negative voltage at the cathode, a strong electric field is formed between the cathode and the anode. A small number of electrons are emitted from the cathode under the influence of the electric field. Under the influence of the electric field, these electrons move towards the anode at a high speed, colliding with gas atoms supplied to the cathode through a gas supply device, thus generating positive ions and new electrons, initiating an electron collision ionization process. Further, as the positive ions accelerate towards the cathode along the direction of the electric field lines, they further collide with gas atoms, generating a large number of positive ions and electrons. The positive ions oscillate back and forth near the cathode until they collide. These reciprocating oscillating ions encounter the compression effect of the three-dimensional electrostatic field, gradually forming a virtual anode at the center of the cathode—a high potential barrier. Electrons also begin to enter the virtual anode, ionizing the introduced gas atoms. Finally, a plasma beam is drawn out by a magnetic field formed by an electromagnetic coil and a magnetic nozzle, thereby providing power to the satellite.

[0024] According to one aspect of the invention, the satellite further includes one or more solar panels. This allows the satellite to utilize solar energy for power during operation, thereby extending its lifespan.

[0025] Furthermore, all of the aforementioned propulsion systems employ electric propulsion, or a combination of electric and chemical propulsion. While electric propulsion is suitable for microsatellites and nanosatellites, a combination of electric and chemical propulsion can be rationally chosen when considering factors such as cost, reliability, and lifespan.

[0026] According to one aspect of the invention, the body has a plurality of hexagonal material removal regions. This allows for maximizing the transmittance of the cathode body.

[0027] According to one aspect of the invention, the body is made of 304 stainless steel. This ensures the strength of the cathode body.

[0028] According to one aspect of the invention, the anode is made of 304 stainless steel.

[0029] According to one aspect of the invention, a support device is provided between the gas supply device and the anode.

[0030] According to one aspect of the invention, the support device is made of boron nitride ceramic.

[0031] According to one aspect of the invention, a sealing device is also provided for sealing the support device, the gas supply device, and the anode.

[0032] According to one aspect of the invention, the satellite further includes an adjustment mechanism for adjusting the thrust direction of the thruster.

[0033] This invention also proposes a method for scaled-down design of an inertial electrostatic confinement ion thruster for a satellite, the method comprising:

[0034] Determine the radius of the discharge region;

[0035] The radius of the cathode region is determined such that the radius of the cathode region is greater than the radius of the discharge region.

[0036] Based on the above technical solutions, it can be seen that the present invention has at least the following beneficial technical effects:

[0037] 1. It adopts a novel ionization method: inertial electrostatic confinement.

[0038] 2. The cathode adopts a hexagonal cage structure with a transmittance of up to 90%, which ensures better electric field uniformity and mechanical properties while ensuring ion passage.

[0039] 3. It adopts a magnetic nozzle design, which uses an electromagnetic coil to provide an adjustable magnetic field. The magnetic nozzle made of silicon steel forms a convergent and expanding magnetic field and magnetic field line structure. Finally, the convergent and expanding magnetic field configuration can accelerate the extraction of ions to form a plasma beam. Compared with the traditional ion thruster acceleration grid, it has the advantages of corrosion resistance, long life and high power.

[0040] Other features and advantages of the present invention will be described below in conjunction with the accompanying drawings. Attached Figure Description

[0041] Exemplary embodiments of the present invention are described with reference to the accompanying drawings, wherein:

[0042] Figure 1 This is a schematic diagram of the satellite of the present invention operating in space.

[0043] Figure 2 This is a schematic diagram of the basic structure of the satellite of this invention.

[0044] Figure 3This is a schematic diagram of the propulsion system used in the satellite of this invention.

[0045] Figure 4 This is a schematic diagram of the hexagonal cage-shaped gate cathode structure of the propulsion system used in the satellite of this invention. Detailed Implementation

[0046] The specific embodiments of the present invention will be described below with reference to the accompanying drawings.

[0047] Figure 1 This is a schematic diagram illustrating the operation of the satellite of this invention in space. See also... Figure 1 The diagram shows the orbit of a satellite orbiting the Earth. Figure 1 In the image, the Sun is located on the left, the Earth on the right, and the satellite 100 of this invention is located in the center. Figure 1 In this invention, the satellite 100 orbits the Earth in an orbit. The satellite 100 includes a body 101 and solar panels 102 located on both sides of the body 101 to power the satellite. Of course, the solar panels 102 are only a preferred structure and not an essential component. Figure 1 The solar panels shown include two on either side of the satellite body 101, but they can also be designed in other forms.

[0048] exist Figure 1 The diagram also shows a three-dimensional orbital coordinate system located at the center of mass of satellite 100, where the X-axis represents the direction of satellite 100's orbital velocity, the Z-axis points towards the Earth, and the Y-axis is perpendicular to the XZ plane formed by the X and Z axes. Depending on the satellite's orbit, its motion around the X-axis is called roll, its motion around the Y-axis is called pitch, and its motion around the Z-axis is called yaw. During satellite operation, thrusters are used to adjust these three motions to maintain the satellite in the desired orbital state.

[0049] Figure 2 This is a schematic diagram of the basic structure of the satellite of this invention. From... Figure 2 The specific structure of the satellite 100 of the present invention can be further seen in the image. See also... Figure 2 including Figure 1 The diagram already shows a main body 101, two solar panels 102, and also includes two thrusters 103 located on the right side of the main body 101, and an adjustment mechanism 104 for adjusting the thrust direction of the thrusters 103. It should be noted that... Figure 2 The arrangement of the thrusters 103 relative to the satellite body 101 is schematic; the number and position of the thrusters 103 can be configured in other ways according to actual needs. For example, thrusters can also be installed on other surfaces of the body 101. In addition to electric thrusters, conventional chemical thrusters can also be installed.

[0050] Figure 3 A schematic diagram of the electric propulsion system for a satellite used in this invention is shown. Figure 4 This is a schematic diagram of the hexagonal cage-like grid cathode structure of the propulsion system used in the satellite of this invention. The electric propulsion system used in the satellite of this invention employs an inertial electrostatic confinement ion thruster. (As shown...) Figure 3 As shown, the inertial electrostatic confinement ion thruster includes: a gas supply device 2; a magnetic guiding device, which includes an electromagnetic coil 6 and a magnetic nozzle 7; an anode 11, which is a spherical shell; and a cathode 10, combined with... Figure 4 It can be seen that the body 202 of the cathode is generally cage-shaped, and the cathode is provided with a first interface 201 connected to the gas supply device 2 and a second interface 203 connected to the magnetic nozzle 7; the cathode 10 is housed inside the anode 11, and the gas supply device 2 is used to introduce gas located outside the anode 11 into the cathode 10; a part of the magnetic nozzle 7 is located outside the anode 11.

[0051] Preferably, the cathode 11 can be fixed on the gas supply device 2 and is insulated from the anode 11.

[0052] In terms of material selection, the gas supply device 2 can be made of stainless steel; the anode 11 can be made of non-magnetic material, such as 304 stainless steel; the cathode can be made of 304 stainless steel with good mechanical and structural properties; and the magnetic nozzle can be made of magnetic material, such as silicon steel.

[0053] The magnetic nozzle 7 is conical, with its small-diameter end fixed to the electromagnetic coil 6 and its large-diameter end fixed to the anode 11. Both the electromagnetic coil 6 and the magnetic nozzle 7 are insulated from the anode 11 and the cathode 10.

[0054] The gas supply device 2 can be in the form of a conduit. For example... Figure 3 As shown, gas 1 is introduced into the interior of the thruster from one end of the gas supply device 2. Preferably, gas 1 is supplied into the interior of the cathode 10 of the thruster. Gas 1 can be any known suitable gas, for example, argon (Ar).

[0055] Specifically, a portion of the gas supply device 2 is located radially outside the anode 11 to introduce gas located outside the anode. A support device 3 is provided between the anode 11 and the gas supply device 2. The support device 3 can be made of insulating ceramic, such as boron nitride ceramic. The gas supply device can be made of stainless steel. A sealing device is also used to seal the gas supply device 2, the support device 3, and the anode 11. The sealing device can be made of rubber.

[0056] The magnetic guiding device and the gas supply device 2 are located on both sides of the cathode 10 in the radial direction.

[0057] The following describes the operation of the electric thruster used in the satellite of this invention, using argon as an example. During operation, a positive voltage is formed on the anode 11 and a negative voltage is formed on the cathode 10. For example, 100V can be applied to the anode 11 while -1200V is applied to the cathode 10. This creates a strong electric field between the anode 11 and the cathode 10. Simultaneously, the cathode, made of metallic material, will release a small number of electrons 5 under the influence of the electric field. Under the strong electric field, these electrons 5 move towards the anode 11 at a high speed. Subsequently, when gas 1 is introduced into the anode 11 through the gas supply device 2, the small number of electrons 5 released from the cathode 10 will collide with argon atoms, generating Ar+ ions and new electrons, initiating the electron collision ionization process. At this time, the Ar+ ions accelerate towards the cathode along the direction of the electric field lines, further ionizing with Ar atoms during the process, generating a large number of Ar+ ions and electrons. Driven by the electric field, the Ar+ ions accelerate towards the center of the cathode.

[0058] like Figure 4 As shown, preferably, multiple hexagonal material removal regions are provided on the body 202 to improve transmittance and ensure good mechanical properties. With this design, accelerated ions can pass through the cathode grid and, if no further collisions occur on their trajectory, can reach the other side of the inter-electrode region. Ions will fly back and forth near the cathode until a collision occurs. These reciprocating oscillating ions encounter the compression effect of the three-dimensional electrostatic field, gradually forming a virtual anode at the center of the cage-like cathode—a high potential barrier—while electrons begin to enter the virtual anode to discharge and ionize the introduced argon gas (Ar).

[0059] Under the influence of inertial electrostatic confinement, high-efficiency discharge ionization has been formed. Ar+ ions need to be accelerated and extracted to form a thruster plasma beam to generate thrust and complete the propulsion task. Electromagnetic coil 6 and magnetic nozzle 7 form magnetic field 8, with electromagnetic coil 6 providing an adjustable magnetic field. The silicon steel magnetic nozzle 7 forms a convergent-expanding magnetic field and magnetic field line structure. Finally, the convergent-expanding magnetic field configuration accelerates the extraction of ions, thus forming the plasma beam 9. To constrain the divergence angle of the plasma beam and further accelerate it, silicon steel is used as the plasma beam extraction device, with the magnetic field lines pointing towards the outside of the anode shell. Based on the electron temperature and ion energy of the ionized plasma, it belongs to the category of low-temperature plasma. Therefore, a relatively weak magnetic field can be used to extract part of the plasma from the central region of the cathode. A DC current of 0-5A is applied to the electromagnetic coil to form a magnetic field of 100G-500G. Ar+ ions and electrons move along the magnetic field lines under the influence of the magnetic field, with the radius of motion following the Larmor cyclotron radius.

[0060]

[0061] r is the rotation radius of the ion and electron, m is the particle mass, v is the particle velocity, q is the charge, and B is the local magnetic field strength.

[0062] As can be seen from the formula, electrons have extremely small mass, while Ar+ ions have relatively large mass. Therefore, during the acceleration along the magnetic field lines towards the exit, there will be a significant difference in momentum. Electrons, with their small gyration radius, will not escape the magnetic field's binding force, while ions, with their large mass and momentum, will have their gyration radius continuously increasing, eventually crossing or even escaping the magnetic field lines. That is, ions and electrons will separate as the downstream magnetic field weakens. Ultimately, Ar+ ions will escape the magnetic field's binding force and be accelerated out, thus the plasma beam 9 will be extracted at high speed by the magnetic nozzle. The reaction force obtained by the thruster is the thrust, which provides the power for the satellite to adjust its attitude and orbit in space.

[0063] Furthermore, since this invention employs a novel ionization method called inertial electrostatic confinement, which offers a wide power range and is easier to scale down in terms of volume, the resulting electric thruster will possess advantages such as high ionization degree and significant scaling capability. The scaling design of the electric thruster of this invention is described below.

[0064] Ions are generated in the sheath, and their thermal rate is approximately the same as that of the background gas.

[0065]

[0066] The drift velocity of electrons is as follows:

[0067]

[0068] Among them, v i Let m be the ion velocity in the ionization region, k be the Boltzmann constant, Tn be the temperature of the neutral gas, and m be the ion velocity in the ionization region. i v is the ion mass. e V is the electron drift velocity, e is the elementary charge, and V is the electron drift velocity. DL m is the potential barrier potential of the ionization region. e Let be the electron mass. Introducing the Langmuir condition, the radius of the ionization region containing the potential barrier can be expressed as:

[0069]

[0070] Vs is the plasma space potential outside the ionization region, V DL Let R be the barrier potential of the ionization region, β be the resistance between the two ends of the plasma bilayer after discharge, β be a dimensionless proportionality constant that can be obtained experimentally, and Nn be the neutral gas number density. From this, the accurate radius r of the discharge region can be determined. SDL The radius of the cathode region needs to be larger than the radius of the discharge region, therefore the radius r of the cathode region... c >rSDL Under the premise of satisfying this discharge ionization, the cathode and anode dimensions of the thruster can be scaled down according to the required discharge power. The scaling range is large, while the upper limit of power depends on the sputtering resistance of the cathode material.

[0071] In other words, the present invention also proposes a scaling design method based on the above-mentioned thruster, which can determine the radius of the cathode region based on the above-mentioned parameters.

[0072] The foregoing description is merely an exemplary embodiment relating to the spirit and principles of the present invention. Those skilled in the art will understand that various changes can be made to the described examples without departing from the spirit and principles, and such changes and their various equivalents are contemplated by the inventors and fall within the scope defined by the claims of the present invention.

Claims

1. An inertial electrostatic confinement ion thruster, comprising: Gas supply device (2); A magnetic guiding device, comprising an electromagnetic coil (6) and a magnetic guiding nozzle (7), wherein the electromagnetic coil (6) provides an adjustable magnetic field, and a convergent-expanding magnetic field and magnetic field line structure are formed through the magnetic guiding nozzle (7) made of silicon steel. Anode (11), wherein the anode is a spherical shell; The cathode (10) has a cage-like body (202) and is provided with a first interface (201) connected to the gas supply device (2) and a second interface (203) connected to the magnetic nozzle (7). The cathode (10) is housed inside the anode (11), and the gas supply device (2) is used to introduce gas located outside the anode (11) into the cathode (10); A portion of the magnetic nozzle (7) is located outside the anode (11); The cathode body (202) has multiple hexagonal material removal areas, and the cathode body (202) is made of 304 stainless steel; A support device (3) is provided between the gas supply device (2) and the anode (11), and a sealing device is also provided to achieve sealing between the support device (3), the gas supply device (2) and the anode (11).

2. The inertial electrostatic confinement ion thruster according to claim 1, characterized in that: The anode (11) is made of 304 stainless steel.

3. The inertial electrostatic confinement ion thruster according to claim 1, characterized in that: The support device (3) is made of boron nitride ceramic.

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