CN116101516B - Satellite with electrostatic electric propulsion system - Google Patents

Abstract

The invention belongs to the technical field of space navigation, and relates to a satellite with an electrostatic electric propulsion system, comprising: a body, one or more propulsion systems, wherein at least one of the propulsion systems uses an inertial electrostatic confinement ion thruster, and the inertial electrostatic The confined ion thruster includes: a gas supply device; a magnetic conduction device, the magnetic conduction device includes an electromagnetic coil and a magnetic conduction nozzle; an anode, the anode is a spherical shell; a cathode, the body of the cathode is in the form of a cage as a whole shape, and the cathode is provided with a first interface connected to the gas supply device and a second interface connected to the magnetic nozzle; the cathode is accommodated inside the anode, and the gas supply The device is used to introduce the gas located outside the anode into the inside of the cathode; a part of the magnetic conduction nozzle is located outside the anode. The invention adopts the brand-new ionization mode of inertial electrostatic restraint, and can adjust the attitude and orbit of the satellite.

Description

A satellite with an electrostatic electric propulsion system

technical field

The invention belongs to the technical field of space navigation, and in particular relates to a satellite with an electrostatic electric propulsion system.

Background technique

Cosmonautics, also known as space flight, space flight, or aerospace flight, refers to the general term for various activities of entering, exploring, developing and utilizing space outside the earth's atmosphere and celestial bodies outside the earth. Vehicles used in space travel include spacecraft and satellites.

The satellite itself has a propulsion system, because when the satellite enters the predetermined orbit through the launch vehicle or other means, it still needs to realize the functions of orbit keeping, precise positioning and resistance compensation, so as to complete application tasks such as network operation and precise observation of the earth . Therefore, the propulsion system of the satellite has an important impact on improving the operating life of the satellite and ensuring the normal operation of the satellite. For micro-nano satellites, in addition to using traditional chemical propulsion devices, electric propulsion devices can also be used. Electric propulsion is a kind of advanced propulsion method that uses electric energy to directly heat the propellant or ionizes and accelerates the propellant to obtain propulsion power. It has high specific impulse, thrust and efficiency. It has broad application prospects in space missions such as space navigation and interstellar navigation.

For example, Patent Document 1 proposes a satellite using chemical propulsion and electric propulsion at the same time, and when it is detected that the electric propulsion fails, corresponding control is performed. For example, controlling the firing of the remaining electric thrusters close to the orbital intersection, and controlling the firing of the chemical thrusters at specific locations. The solution in Patent Document 1 is only for post-event remediation after the failure of the electric propulsion, compared with the situation where no failure occurs, it does not improve the life of the electric propulsion or even the satellite.

Patent Document 1:

Application publication number: CN 106275506 A, applicant: Boeing Company, classification number: B64G1/10, B64G1/24, B64G1/40, title of invention: effective position-keeping design for hybrid fuel system in response to electric propulsion failure.

Specific to the type of electric propulsion, the ion electric propulsion system and the Hall electric propulsion system are currently the most widely used in orbit. The structure of the Hall thruster is relatively simple, the volume is small, the specific impulse is lower than that of the ion thruster, and the required propellant injection is larger, but its thrust is larger and the orbit change time is shorter; the ion structure is relatively complex and the volume is larger. The specific impulse is higher than that of the Hall thruster, and the required propellant filling amount is small, but the thrust is small and the orbit change time is long.

For this reason, Patent Document 2 proposes a satellite platform compatible with ion and Hall electric propulsion configurations. The electric propulsion subsystem mainly includes two thruster modules, a storage tank supply module and a power supply module. The electric propulsion system adopted can Two forms of thrusters, ion and Hall, are used to complete satellite orbit change and satellite in-orbit north-south position maintenance, east-west position maintenance and angular momentum unloading tasks, and ion or Hall can also be used alone to complete the above tasks, in a separate configuration For ion or Hall electric propulsion, the unconfigured functional modules can be removed directly, and the layout of other instruments and equipment on the platform does not need to be adjusted. However, this structure is still combined based on known thrusters, and does not overcome the shortcomings of each thruster itself.

Patent Document 2:

Application publication number: CN 115503984 A, applicant: China Academy of Space Technology, classification number: B64G1/40, invention name: a satellite platform compatible with ion and Hall electric propulsion configurations.

Therefore, research on thrusters is still needed to provide better propulsion for satellites.

Contents of the invention

Combining the inventor's research and practical experience in this field, the following improved technical solutions are proposed here.

A satellite comprising:

ontology,

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:

air supply device;

Magnetic conduction device, described magnetic conduction device comprises electromagnetic coil and magnetic conduction nozzle;

an anode, the anode is a spherical shell;

The cathode, the body of the cathode is generally cage-shaped, and the cathode is provided with a first interface connected to the gas supply device and a second interface connected to the magnetic nozzle;

The cathode is accommodated inside the anode, and the gas supply device is used to introduce gas located outside the anode into the inside of the cathode;

A part of the magnetic nozzle is located outside the anode.

According to the satellite of the present invention, it can adopt electric propulsion to realize the change of attitude or orbit. Specifically, by using an inertial electrostatic thruster, a positive voltage is formed on the anode and a negative voltage is formed on the cathode, so that a strong electric field will be formed between the cathode and the anode, and a small amount of electrons will escape from the cathode under the action of the electric field. Under the action of the electric field, the electrons will move toward the anode at a higher speed, and during this process, they will collide with the gas atoms supplied to the cathode through the gas supply device, thereby generating positive ions and new electrons, and starting the electron impact ionization process. Furthermore, as the positive ions accelerate toward the cathode along the direction of the electric field lines, they will further collide with gas atoms and ionize during the process, generating a large number of positive ions and electrons. Positive ions fly back and forth near the cathode until they collide. These reciprocating oscillating ions encounter the compression effect of the three-dimensional electrostatic field, and gradually form a virtual anode in the center of the cathode, that is, a high potential electric barrier, and electrons will also start to enter the virtual anode to form discharge ionization on the incoming gas atoms. Finally, the plasma beam is drawn out through the magnetic field formed by the electromagnetic coil and the magnetic nozzle to provide power for the satellite.

According to an aspect of the invention, the satellite further includes one or more solar panels. Thus, the satellite can be powered by solar energy during operation, and the service life of the satellite can also be improved.

Further, the plurality of propulsion systems all adopt electric propulsion systems, or use a combination of electric propulsion systems and chemical propulsion systems. For micro-nano satellites, it is appropriate to choose an electric propulsion system, but considering factors such as cost, reliability, and service life, it is reasonable to choose a combination of an electric propulsion system and a chemical propulsion system.

According to an aspect of the invention, the body has a plurality of hexagonal material-removed regions. Thus, the transmittance of the bulk of the cathode can be increased as much as possible.

According to one aspect of the present invention, the body is made of 304 stainless steel. Thereby, the strength of the main body of the cathode can be ensured.

According to one aspect of the present invention, the anode is made of 304 stainless steel.

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

According to one aspect of the present invention, the supporting device adopts boron nitride ceramics.

According to an aspect of the present invention, a sealing device is further provided for sealing between the support device, the gas supply device and the anode.

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

The present invention also proposes a method for scaling down the inertial electrostatic confinement ion thruster of the satellite, the method comprising:

Determine the radius of the discharge area;

The radius of the cathode region is determined such that the radius of the cathode region is larger than the radius of the discharge region.

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

1. The new ionization method of inertial electrostatic confinement is adopted.

2. The cathode adopts a hexagonal cage structure, and the transmittance is as high as 90%. It has better electric field uniformity and mechanical properties under the condition of ensuring the passage of ions.

3. The magnetic nozzle design is adopted, and the electromagnetic coil is used to provide an adjustable magnetic field. The magnetic nozzle of silicon steel material forms a convergent and expanding magnetic field and a magnetic induction line structure. Finally, the convergent and expanding magnetic field configuration can accelerate the extraction of ions to form plasma Compared with the traditional ion thruster acceleration grid, the volume beam has the advantages of corrosion resistance, long life and high power.

Other features and advantages of the present invention will be described below with reference to the accompanying drawings.

Description of drawings

Exemplary embodiments of the invention are described with reference to the accompanying drawings, in which:

Fig. 1 is a schematic diagram of the operation of the satellite of the present invention in space.

Fig. 2 is a schematic diagram of the basic structure of the satellite of the present invention.

Fig. 3 is a schematic diagram of the propulsion system adopted by the satellite of the present invention.

Fig. 4 is a schematic structural view of the hexagonal cage-shaped grid cathode of the propulsion system adopted by the satellite of the present invention.

Detailed ways

Specific embodiments of the present invention will be described below in conjunction with the accompanying drawings.

Fig. 1 is a schematic diagram of the operation of the satellite of the present invention in space. Referring to Fig. 1, a schematic diagram of the operation of a satellite orbiting the earth is shown therein. In FIG. 1 , the sun in the solar system is on the left, the earth is on the right, and the satellite 100 of the present invention is in the middle. In FIG. 1, a satellite 100 of the present invention orbits the earth in an orbit. The satellite 100 includes a main body 101 and solar panels 102 located on both sides of the main body 101 for powering the satellite. Of course, the solar panel 102 is only a preferred structure, not a necessary component. The solar panel shown in FIG. 1 includes two panels located on both sides of the satellite body 101 , but it can also be designed in other forms.

Also shown in Fig. 1 is a three-dimensional orbital coordinate system at the center of mass of the satellite 100, wherein the X direction represents the running speed direction of the satellite 100, the Z axis points to the earth, and the Y axis is perpendicular to the XZ formed by the X axis and the Z axis. flat. According to the operation of the satellite, the operation of the satellite around the X axis is called roll, the operation around the Y axis is called pitch, and the operation around Z is called yaw. During the operation of the satellite, the above three motions of the satellite are adjusted by the propeller to make the satellite in the required operating state.

Fig. 2 is a schematic diagram of the basic structure of the satellite of the present invention. The specific structure of the satellite 100 of the present invention can be further seen from FIG. 2 . Referring to FIG. 2 , it includes the main body 101 shown in FIG. 1 , two solar panels 102 , two propellers 103 on the right side of the main body 101 and an adjustment mechanism 104 for adjusting the thrust direction of the propellers 103 . It should be noted that the arrangement of the thrusters 103 relative to the satellite body 101 in FIG. 2 is schematic, and the number and positions of the thrusters 103 can be set in other forms according to actual needs. For example, propellers may also be provided on other surfaces of the body 101 . In addition to setting electric thrusters, traditional chemical thrusters can also be set.

Fig. 3 shows a schematic diagram of the electric propulsion system applied to the satellite of the present invention, and Fig. 4 is a schematic structural diagram of the hexagonal cage grid cathode used in the propulsion system of the satellite of the present invention. The electric propulsion system adopted by the satellite of the present invention adopts an inertial electrostatic confinement ion thruster. As shown in Figure 3, the inertial electrostatic confinement ion thruster includes: a gas supply device 2; a magnetic conduction device, the magnetic conduction device includes an electromagnetic coil 6 and a magnetic conduction nozzle 7; an anode 11, and the anode is spherical Housing; cathode 10, as can be seen from FIG. 4 , the body 202 of the cathode is cage-shaped as a whole, and the cathode is provided with a first interface 201 connected to the gas supply device 2 and connected to the magnetic spray The second interface 203 connected to the tube 7; the cathode 10 is accommodated inside the anode 11, and the gas supply device 2 is used to introduce the gas outside the anode 11 into the cathode 10; A part of the magnetic nozzle 7 is located outside the anode 11 .

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

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 better mechanical properties; the magnetic nozzle can be made of magnetic material , such as silicon steel.

Magnetic nozzle 7 becomes tapered, and one end of its small diameter is fixed with electromagnetic coil 6, and one end of large diameter is then fixed with anode 11, and electromagnetic coil 6 and magnetic nozzle 7 are all insulated with anode 11 and negative electrode 10.

The gas supply device 2 may take the form of a conduit. As shown in FIG. 3 , gas 1 is introduced from one end of gas supply device 2 into the interior of the thruster. Preferably, the gas 1 is supplied to the inside of the cathode 10 of the thruster. The gas 1 can be any known suitable gas, for example, argon Ar can be selected.

Specifically, a part of the gas supply device 2 is located radially outside the anode 11, so as to introduce gas located outside the anode. Between the anode 11 and the gas supply device 2, a support device 3 is arranged. The support device 3 can be made of insulating ceramics, such as boron nitride ceramics. The air supply device can be made of stainless steel. Between the gas supply device 2 , the supporting device 3 and the anode 11 , a sealing device is used to seal the three. The sealing device can adopt rubber material.

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

Hereinafter, the working process of the electric thruster adopted by the satellite of the present invention will be described by taking argon gas as an example. In operation, a positive voltage is developed on the anode 11 and a negative voltage is developed on the cathode 10 . For example, 100V can be applied to the anode 11 and -1200V can be applied to the cathode 10 at the same time. Thereby, a strong electric field will be formed between the anode 11 and the cathode 10 . At the same time, the cathode made of metal material will escape a small amount of electrons 5 under the action of an electric field. Under the action of a strong electric field, the electrons 5 will move to the anode 11 at a high speed. Then, when the gas 1 passes through the gas supply device 2 passes into the inside of the anode 11, and a small amount of electrons 5 escaping from the cathode 10 will collide with argon atoms, thereby generating Ar+ ions and new electrons, and starting the electron impact ionization process. At this time, Ar+ will accelerate toward the cathode along the direction of the electric field line, and will further collide with Ar atoms and ionize during the process, generating a large number of Ar+ ions and electrons. Ar+ is driven by the electric field and accelerates toward the center of the cathode.

As shown in FIG. 4 , preferably, a plurality of hexagonal material removal regions are provided on the body 202 to improve the transmittance and ensure good mechanical properties. With this design, accelerated ions can pass through the cathode grid and, if no more collisions occur on their trajectories, reach the other side of the inter-electrode region. The ions fly back and forth near the cathode until they collide. These reciprocating oscillating ions encounter the compression effect of the three-dimensional electrostatic field, and gradually form a dummy anode in the center of the cage cathode, that is, a very high electric potential barrier, and electrons will also start to enter the dummy anode to form a Discharge ionization.

Under the action of inertial electrostatic constraints, high-efficiency discharge ionization has been formed. It is necessary to accelerate the extraction of Ar+ ions to form thruster plasma beams to generate thrust to complete the propulsion task. The electromagnetic coil 6 and the magnetic nozzle 7 form a magnetic field 8, and the electromagnetic coil 6 provides an adjustable magnetic field, and the magnetic nozzle 7 of silicon steel material forms a convergent and dilated magnetic field and a magnetic induction line structure, and finally the convergent and dilated magnetic field configuration can Accelerate the extraction of ions to form the effect of plasma beam 9 . In order to constrain the divergence angle of the plasma beam and further accelerate the plasma, the magnetically permeable material silicon steel is used as the extraction device of the plasma beam, and the direction of the magnetic induction line points to the outside of the anode casing. According to the electron temperature and ion energy of the ionized plasma, it belongs to the category of low-temperature plasma, so a weaker magnetic field can be used to extract part of the plasma in the central area of the cathode, and the electromagnetic coil is energized with 0-5A DC to form a 100G-500G magnetic field. Ar+ ions and electrons will move along the lines of magnetic induction under the action of a magnetic field, and the radius of motion follows the Larmor radius of gyration:

r is the radius of rotation of ions and electrons, m is the particle mass, v is the particle velocity, q is the charge, and B is the local magnetic field strength.

It can be seen from the formula that the mass of electrons is extremely small, and the mass of Ar+ ions is relatively large. Therefore, during the process of accelerating and moving along the magnetic field lines to the exit, there will be a large difference in momentum, and the small radius of gyration of electrons will not Freed from the constraints of the magnetic field, the mass and momentum of the ions are large, and the radius of gyration increases all the time. There will be a phenomenon of crossing or even breaking away from the magnetic field line, that is, the ions and electrons will separate as the downstream magnetic field weakens. In the end, Ar+ breaks away from the shackles of the magnetic field and is accelerated to be ejected, so that the plasma beam 9 is drawn out by the magnetic nozzle at high speed, and the reaction force obtained by the thruster is the thrust, thus providing the power for the satellite to adjust its orbit in space.

Furthermore, since the present invention adopts a brand-new ionization method of inertial electrostatic confinement, the power span of this ionization method is large, and the volume scaling is easier to realize, and the electric thruster based on this will have a high degree of ionization, The advantage of a large scale. The scale design of the electric thruster of the present invention is described below.

Ions are generated in the sheath with approximately the same thermal velocity as the background gas:

The drift speed of electrons is as follows:

Among them, v _i is the ion velocity in the ionization region, k is the Boltzmann constant, Tn is the neutral gas temperature, m _i is the ion mass, v _e is the electron drift velocity, e is the element charge, and V _DL is the potential in the ionization region barrier potential, me _is the mass of the electron. Introducing the Langmuir condition, the radius of the ionization region where the electric potential barrier is located can be expressed as:

Vs is the plasma space potential outside the ionization region, V _DL is the barrier potential of the ionization region, R is the resistance value between the two ends of the plasma double layer after the discharge is formed, and β is a dimensionless proportional coefficient, which can be obtained through experiments , Nn is the number density of neutral gas. Thus, the exact radius r _SDL of the discharge area can be obtained, while the radius of the cathode area needs to be higher than the radius of the discharge area, so the radius r _c >r _SDL of the cathode area. Under the discharge ionization prerequisite, the size of the cathode and anode of the thruster can be scaled according to the required discharge power. The scale range is large, and the upper limit of the power depends on the sputtering resistance of the cathode material.

That is to say, the present invention also proposes a design method based on the scaling of the above-mentioned thruster, which can calculate the radius of the cathode region according to the above-mentioned parameters.

What have been described above are merely exemplary embodiments pertaining to the spirit and principles of the invention. It will be apparent to those skilled in the art that various changes may be made to the examples described without departing from the spirit and principles described, and these changes and their various equivalents are envisioned by the inventor and fall within the scope of the present invention. within the scope defined by the claims of the present invention.

Claims (5)

1. A satellite comprising: ontology(101), 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: Air supply device (2); The magnetic conduction device, the magnetic conduction device includes an electromagnetic coil (6) and a magnetic conduction nozzle (7), the electromagnetic coil (6) provides an adjustable magnetic field, and forms a convergent expansion through the magnetic conduction nozzle (7) made of silicon steel material Type magnetic field and magnetic induction line structure; an anode (11), the anode is a spherical shell; Cathode (10), the body (202) of the cathode is generally in the shape of a cage, and the cathode is provided with a first interface (201) connected to the gas supply device (2) and connected to the magnetic nozzle The second interface (203) connected to the pipe (7); The cathode (10) is accommodated inside the anode (11), and the gas supply device (2) is used to introduce gas outside the anode (11) into the cathode (10); A part of the magnetically conductive nozzle (7) is located outside the anode (11); The body (202) of the cathode has multiple hexagonal material removal areas, and the body (202) of the cathode is made of 304 stainless steel; Between the gas supply device (2) and the anode (11), a support device (3) is provided, and a sealing device is also provided to realize the support device (3), the gas supply device (2) and the seal between the anode (11). 2. The satellite according to claim 1, characterized in that: The satellite also includes one or more solar panels (102). 3. The satellite according to claim 1, characterized in that: All of the multiple propulsion systems use electric propulsion systems, or a combination of electric propulsion systems and chemical propulsion systems. 4. The satellite according to any one of claims 1-3, characterized in that: The anode (11) is made of 304 stainless steel. 5. The satellite according to claim 1, characterized in that: The supporting device (3) adopts boron nitride ceramics.