CN211397783U - Ultra-low rail variable thrust air suction type magnetic plasma thruster - Google Patents

Abstract

The utility model provides an ultra-low orbit variable thrust suction-type magnetic plasma thruster, which comprises a suction duct, an air inlet and an air outlet, which are used for compressing thin air and sucking it into the thruster; The cavity, communicated with the air outlet, is composed of a cathode body, an anode body and a power source. The cathode body and the anode body are all electrically connected to the power source, and are used to ionize the compressed thin air into plasma and convert it into plasma under the action of an electric field. The plasma accelerates ejection, and the discharge chamber is provided with a magnetic field to further enhance the effect of the plasma acceleration ejection; the restrictor valve is located between the air inlet and the discharge chamber, and the intake end of the restrictor valve is connected to the outlet The air port is in communication, and the air outlet end of the restrictor valve is in communication with the discharge chamber, and is used for controlling the mass flow of the compressed air mass entering the discharge chamber. The process of capturing, storing, ionizing and accelerating the thin air in the environment by the thruster is realized, thereby stably generating thrust. The utility model is applied to the fields of aerospace technology and plasma.

Description

Ultra-low orbit variable thrust air-breathing magnetic plasma thruster

technical field

The utility model relates to the field of aerospace technology and plasma, in particular to an ultra-low orbit variable thrust suction type magnetic plasma thruster.

Background technique

With the increasing number of satellites in space orbits, ultra-low orbits have become a new choice for satellites to expand their operating range and enhance their mission capabilities. Compared with other orbits, the operation of satellites in ultra-low orbits can significantly reduce the launch cost of satellites and improve their navigation and positioning accuracy and response speed. Positioning and navigation, remote sensing and other fields have broad application prospects. However, due to the particularity and complexity of the ultra-low orbit space environment, the persistent presence of satellites will face the problems of short lifespan, difficult propellant replenishment, and high on-orbit maintenance costs, which seriously restrict the development of space ultra-low orbit satellites.

Pursuing a propulsion system with stable performance, long life, light weight and low cost is the frontier research focus in the current aerospace field. The development of new propulsion schemes based on the substances existing in the space environment as much as possible can effectively reduce the costs related to satellites, increase their lifespans, and provide a new scheme for orbit control of ultra-low orbit satellites. However, there are few related studies, and it is impossible to achieve accurate and variable thrust, and it cannot meet the power requirements of multiple flight missions of ultra-low orbit satellites.

Utility model content

In view of the prior art, the purpose of this utility model is to provide an ultra-low orbit variable thrust air-breathing magnetic plasma thruster.

The technical solutions it adopts are:

Ultra-low orbit variable thrust air-breathing magnetic plasma thruster, including:

The suction duct, located at the head end of the thruster, is provided with an air inlet and an air outlet for compressing the rarefied air and sucking it into the thruster;

The discharge chamber is located at the rear end of the thruster and communicated with the air outlet. It consists of a cathode body, an anode body and a power source. The cathode body and the anode body are all electrically connected to the power source and are used to ionize the compressed thin air into plasma. Under the action of the electric field, the plasma is accelerated and ejected, and an accelerating magnetic field is arranged in the discharge chamber to further enhance the effect of the accelerated ejection of the plasma;

The restrictor valve is located between the intake duct and the discharge chamber, the intake end of the restrictor valve communicates with the air outlet, and the air outlet end of the restrictor valve communicates with the discharge chamber for controlling the compressed air entering the discharge chamber mass flow of the group.

Further preferably, it also includes a shunt, the shunt is located between the restrictor valve and the discharge chamber;

The cathode body and the anode body are both hollow cylindrical structures, the cathode body is located in the cavity of the anode body, an annular cavity is enclosed between the outer wall of the cathode and the inner wall of the anode, and the discharge cavity is composed of the annular cavity and the anode. The remaining cavity in the body is formed, the outer wall of the anode body is surrounded by a magnetic coil, and the accelerating magnetic field is generated by the magnetic coil;

One end of the shunt is provided with a shunt inlet, and the other end of the shunt is provided with a first shunt outlet and a second shunt outlet surrounding the first shunt outlet, the first shunt outlet and the second shunt outlet are both. It is communicated with the shunt inlet through the shunt structure in the shunt;

The shunt inlet is communicated with the gas outlet end of the restrictor valve, the first shunt outlet is communicated with the cavity of the cathode body, and the second shunt outlet is communicated with the annular cavity.

Further preferably, the power supply includes:

The ignition circuit is electrically connected with the cathode body and the anode body for igniting the air in the discharge chamber;

The main discharge circuit is electrically connected with the cathode body and the anode body for supplying an electric field to the discharge chamber.

Further preferably, the ignition circuit includes:

a first charging power source for charging the first capacitor;

The first capacitor includes a first terminal and a second terminal, the first terminal of the first capacitor is coupled with the anode of the first charging power source, the second terminal of the first capacitor is coupled with the cathode of the first charging power source, and the cathode body is coupled to the anode of the first charging power source. The second terminal of the first capacitor and the cathode of the first charging power source are coupled;

The first silicon controlled rectifier includes a first terminal and a second terminal. The first terminal of the first silicon controlled rectifier is coupled with the first terminal of the first capacitor and the anode of the first charging power supply. The second terminal is coupled to the anode body.

Further preferably, the main discharge circuit includes a second charging power supply, a second silicon controlled rectifier, a diode, a protection resistor, a relay, n second capacitors C ₁ ˜C _n and n inductors L ₁ ˜L _n , wherein , n is a natural number greater than 1;

The second silicon controlled rectifier, the protection resistor, the relay and each second capacitor include a first terminal and a second terminal;

The first terminal of the first second capacitor C ₁ is coupled to the anode of the second charging power supply, and the first terminal of the i-th second capacitor C _i is connected to the first terminal of the i+1-th second capacitor C _i+1 Through the coupling of the ith inductance Li, the second terminal of each of the second capacitors C ₁ ˜C _n is coupled with the cathode of the second charging power source, where 1≤i<n;

The first terminal of the _nth second capacitor Cn is also coupled with the input end of the diode through the nth inductance _Ln , and the output end of the diode is coupled with the anode body through a matching resistor;

The first terminal of the second silicon controlled rectifier is respectively coupled to the cathode of the second charging power supply and the second terminal of each of the second capacitors C ₁ ˜C _n , and the second terminal of the second silicon controlled rectifier is connected to Cathode body coupling;

The first terminal of the first second capacitor _C1 is also coupled to the first terminal of the protection resistor, the second terminal of the protection resistor is coupled to the first terminal of the relay, and the second terminal of the _second second capacitor C2 is coupled to the second terminal of the relay and grounded.

Further preferably, the air intake duct is a trumpet-shaped structure, the air inlet is located at the large end of the trumpet-shaped structure, and the air outlet is located at the small end of the trumpet-shaped structure.

Further preferably, part of the suction duct near the air outlet on the suction duct is made of nitrogen storage and oxygen storage solid solution materials, the remaining part of the suction duct is made of foamed silicon carbide material, and the foam on the suction duct is made of solid solution material. The silicon carbide material is filled with carbon molecular sieves.

Further preferably, the filling ratio of carbon molecular sieve in the foamed silicon carbide material on the air intake channel gradually increases along the direction from the air inlet to the air outlet.

Further preferably, a reinforced coating is provided on the air intake duct at the position of the air intake.

Further preferably, the cathode body is made of tungsten metal material, and the anode body is made of titanium metal material.

Beneficial technical effects of the present utility model:

1. The structure of the present utility model is simple. By ionizing and accelerating the compressed thin air, thrust is generated without carrying propellant, which can not only avoid the limitation of propellant exhaustion on the life of the thruster, but also save complicated The thruster supply device and ground assembly can effectively reduce the overall weight and cost of the thruster and further improve the performance.

2. By absorbing, compressing and storing the rarefied air, the present utility model realizes the flow control of the compressed air entering the discharge chamber by using the current limiting valve, and at the same time uses the power supply to control the discharge power of the thinned air in the discharge chamber, which can effectively realize the impact on the thrust. The precise control of the satellite can effectively meet the needs of different missions of the satellite.

3. In the present invention, by arranging an accelerating magnetic field in the discharge chamber, combined with the electric field generated between the anode body and the cathode body, the plasma generated by the air discharge has a stable acceleration effect, thereby obtaining thrust.

Description of drawings

1 is a cross-sectional view of a thruster in this embodiment;

Fig. 2 is the schematic diagram of the shunt process of the shunt in the present embodiment;

3 is a schematic structural diagram of the restrictor valve in this embodiment;

FIG. 4 is a schematic circuit diagram of the power supply in this embodiment.

Detailed ways

In order to make the objectives, technical solutions and advantages of the present disclosure more clear, the present utility model will be further described in detail below with reference to the specific embodiments and the accompanying drawings. It should be noted that, in the drawings or descriptions of the specification, the undescribed contents and some English abbreviations are the contents well known to those skilled in the art. Some specific parameters given in this embodiment are only exemplary, and the values can be changed to appropriate values accordingly in different implementations.

As shown in FIG. 1 , the ultra-low orbit variable thrust air-breathing magnetic plasma thruster includes a suction duct 1 , a restrictor valve 2 , a shunt 3 , a cathode body 4 , an anode body 5 , a magnetic coil 6 and a power source 7 and other sections, which specifically:

The air intake duct 1 has a trumpet-shaped structure with a gradually decreasing cross-sectional area, and is located at the head end of the thruster. An air inlet 11 and an air outlet 12 are arranged on it. 12 is located at the small end of the horn-shaped structure, which can effectively compress the thin air and increase the bulk density of the incoming air; the part of the air inlet 1 on the air inlet 1 near the air outlet 12 is made of nitrogen storage and oxygen storage solid solution materials. The remaining part of the air intake duct 1 is made of foamed silicon carbide material, and the foamed silicon carbide material on the air intake duct 1 is filled with carbon molecular sieve. Among them, the filling ratio of carbon molecular sieve in the foamed silicon carbide material on the intake duct 1 gradually increases along the direction from the air inlet 11 to the air outlet 12, and finally achieves the directional adsorption of the lean air, which is convenient for nitrogen storage in the air intake 1. , The oxygen storage solid solution material stores the air; at the same time, a reinforced coating made of C-SiC nanocomposite material is provided on the air intake duct 1 at the position of the air intake 11, which effectively strengthens the air intake duct 1 Collision protection performance against high-speed particles.

In this embodiment, part of the air intake duct 1 made of nitrogen storage and oxygen storage solid solution materials on the air intake duct 1 accounts for 1/3 to 1/2 of the total length of the air intake duct 1, and the part of the air intake duct 1 is made of The solid solution material for nitrogen storage and oxygen storage is specifically Ce _0. Zr _0. O.xBaO, so that the compressed rarefied air can be stored in the air intake duct 1 and also in the solid solution material of the air intake duct 1 itself.

The cathode body 4 and the anode body 5 are both hollow cylindrical structures. Specifically, the cathode body 4 is a hollow cylindrical structure made of tungsten metal material, and the anode body 5 is a hollow expansion ring structure made of titanium metal material. The cathode body 4 is a hollow cylindrical structure. It is located in the cavity of the anode body 5, so that an annular cavity 81 is enclosed between the outer wall of the cathode body 4 and the inner wall of the anode body 5, and the annular cavity 81 and the remaining cavity in the anode body form a discharge cavity together, wherein the remaining cavity in the anode body forms a discharge cavity. The cavity refers to the remaining part of the cavity of the anode body after removing the part occupied by the annular cavity 81 and the cathode body 4; preferably, the axis of the cathode body 4 and the axis of the anode body 5 are parallel to each other; The axis of the body 4 coincides with the axis of the anode body 5. Specifically, the length of the cathode body 4 is 1/3 of the length of the anode body 5, and one end of the cathode body 4 and one end of the anode body 5 are located on the same cross section. The other end of the body 4 is located in the cavity of the anode body 5 . The air compressed and sucked into the thruster through the air inlet 1 enters the discharge chamber and is ionized into plasma and accelerated and ejected under the action of the electric field; the power source 7 is also located in the cavity 83 of the anode body, and the cathode body 4 and the power source The cathode of 7 is electrically connected, and the anode body 5 is electrically connected to the anode of the power source 7. The power source 7 not only provides energy for the electric field generated in the discharge chamber, but also ignites the air ionization in the electric field. At the same time, the outer wall of the anode body 5 is surrounded by a magnetic coil 6. After the magnetic coil 6 is energized, an accelerating magnetic field along the axial direction of the anode body 5 is generated in the discharge chamber. Combined with the electric field, the plasma generated by the air discharge has a stable acceleration effect. Further, thrust is obtained, wherein the hollow structure of the cathode body 4 and the anode body 5 can also effectively increase the ionization rate and acceleration efficiency. In this process, adjusting the output voltage of the power supply 7 to the cathode body 4 and the anode body 5 can adjust the electric field, or adjust the current in the magnetic field line to adjust the accelerating magnetic field, which can achieve the effect of adjusting the thrust. Effect.

The restrictor valve 2 is located between the intake duct 1 and the discharge chamber, the intake end of the restrictor valve 2 is connected to the air outlet 12, and the air outlet end of the restrictor valve 2 is communicated with the discharge chamber through the flow divider 3 to control the discharge The mass flow of the compressed air mass of the cavity. 2, one end of the diverter 3 is provided with a diverter inlet 31, and the other end of the diverter 3 is provided with a first diverter outlet 32 and a second diverter outlet 33 surrounding the first diverter outlet 32. The first diverter outlet 32. The second shunt outlet 33 is connected to the shunt inlet 31 through the shunt structure in the shunt 3; the shunt inlet 31 is communicated with the air outlet of the restrictor valve 2, the first shunt outlet 32 is communicated with the cavity 82 of the cathode body, and the first shunt outlet 32 is communicated with the cavity 82 of the cathode body. The second branch outlet 33 is communicated with the annular cavity 81. The arrow in FIG. 2 is the air flow direction. The air mass is separated into the discharge chamber 81 through the first branch outlet 32 and the second branch outlet 33, so that the air can be more uniformly distributed. Distributed in the discharge chamber 81, thereby improving the efficiency of ionization and acceleration.

Referring to FIG. 3 , the restrictor valve 2 in the present embodiment adopts a valve disclosed in the patent CN 105840904B, which has an annular structure as a whole, and consists of an air intake port 20 of the restrictor valve, a helical drive coil 21 and a coil bobbin 22 , Truncated cone-shaped reed 23, gasket 24, limit block 25, main valve body 26, valve cavity 27 and O-ring 28 to control the mass flow of compressed air mass entering the discharge cavity and adjust the thrust. In the restrictor valve 2 , the helical drive coil 21 and the truncated cone-shaped reed 23 constitute the valve driving mechanism, and the truncated-cone shaped reed 23 is an actuating member. When the pulse current flows through the helical drive coil 21 , according to the principle of electromagnetic induction, the transient magnetic field generated by the pulse current induces an annular current in the truncated cone-shaped reed 23 in the opposite direction. The interaction between the induced current and the radial component of the magnetic field generated by the coil current will generate an axial Lorentz force in the truncated cone reed 23, when the axial Lorentz force is much larger than the initial value of the truncated cone reed 23 The outer edge of the truncated cone-shaped reed 23 will rise rapidly under the action of the axial Lorentz force. When the axial Lorentz force is less than the initial elastic force of the reed, the truncated cone-shaped reed 23 will reverse. Therefore, during the working process of the restrictor valve 2, by adjusting the current of the helical drive coil 21 to control the opening size of the electromagnetic force drive valve port, the mass flow of the compressed air mass entering the discharge chamber can be precisely controlled, and finally the variable thruster can be realized. The thrust function has the characteristics of fast response, long life, strong anti-interference and high durability.

Since the air in the space is relatively thin, although the air intake duct 1 simply compresses the air, it is still difficult to discharge and ignite the air in the discharge chamber. Therefore, this embodiment designs a power supply 7 compatible with the ignition circuit 71 and the main discharge circuit 72 as shown in FIG. 4 for the characteristics of lean atmosphere ignition. The power source 7 can not only effectively improve the ignition success rate and promote the ionization rate, but also have the function of controlling the discharge power and adjusting the thrust change. The power supply 7 mainly includes an ignition circuit 71 and a main discharge circuit 72 . The ignition circuit 71 is electrically connected with the cathode body 4 and the anode body 5 for igniting the air in the discharge chamber; the main discharge circuit 72 is electrically connected with the cathode body 4 and the anode body 5 for providing the discharge chamber with electric field.

Referring to FIG. 4 , the ignition circuit 71 includes: a first charging power source 711 , a first capacitor 712 and a first silicon controlled rectifier 713 . The first charging power source 711 of the ignition circuit 71 is a low-power high-voltage charging power source, the first capacitor 712 is a low-capacity capacitor, and the first charging power source 711 is used for charging the low-capacity, high-voltage capacitor; the first silicon controlled rectifier 713 is used for The conduction between the ignition circuit 71 and the thruster is controlled while preventing reverse current from flowing into the ignition circuit 71 . The first capacitor 712 and the first silicon controlled rectifier 713 are both provided with a first terminal and a second terminal.

The specific structure of the ignition circuit 71 is as follows: the first terminal of the first capacitor 712 is coupled to the anode of the first charging power source 711, the second terminal of the first capacitor 712 is coupled to the cathode of the first charging power source 711, and the cathode body 4 is coupled to the first charging power source 711. The second terminal of the capacitor 712 is coupled with the cathode of the first charging power source 711; the first terminal of the first silicon controlled rectifier 713 is coupled with the first terminal of the first capacitor 712 and the anode of the first charging power source 711, and the first controllable The second terminal of the silicon rectifier 713 is coupled to the anode body 5 .

The main discharge circuit 72 includes: a second charging power source 721 , n second capacitors C ₁ ˜C _n , n inductors L ₁ ˜L _n , a diode 722 , a second silicon controlled rectifier 723 , a protection resistor 724 and a relay 25 , Among them, n is a natural number greater than 1. The second charging power source 721 of the main discharge circuit 72 is a high-power and high-current charging power source, the second capacitor is a large-capacity capacitor, and the second charging power source 721 is used to charge the large-capacity capacitor; the matching combination of the second capacitor and the inductance is a thruster Provide the required discharge waveform; the diode 722 is used to prevent the high voltage charging power source 7 of the ignition circuit 71 from charging the capacitor of the main discharge circuit 72; the second thyristor rectifier 723 is used to control the conduction between the main discharge circuit 72 and the thruster. The protection resistor 724 is used to release the electrical energy stored in the main discharge circuit 72 through the protection resistor 724 when the thruster discharge fails; the relay 25 is used to control the protection resistor 724 to communicate with the main discharge circuit 724. The connection and disconnection of the discharge circuit 72 . The matching resistor 73 in FIG. 4 is used to perform load matching between the discharge circuit impedance and the thruster discharge impedance, so as to improve the thruster energy utilization efficiency. Wherein, the second silicon controlled rectifier 723, the protection resistor 724, the relay 25 and each second capacitor are provided with a first terminal and a second terminal.

The specific structure of the main discharge circuit 72 is as follows: the first terminal of the first second capacitor C ₁ is coupled with the anode of the second charging power source 721 , and the first terminal of the i-th second capacitor C _i is connected to the i+1-th first terminal. The first terminals of the two capacitors C _{i +1} are coupled through the i-th inductor Li, and the second terminals of each of the second capacitors C ₁ to C _n are coupled to the cathode of the second charging power source 721 , where 1≤i<_n; the first terminal of the nth second capacitor Cn is also coupled to the input end of the diode 722 through the nth inductance _Ln , and the output end of the diode 722 is coupled to the anode body 5 through the matching resistor 73; the second thyristor The first terminal of the rectifier 723 is respectively coupled to the cathode of the second charging power source 721 and the second terminal of each of the second capacitors C ₁ ˜C _n , and the second terminal of the second silicon controlled rectifier 723 is coupled to the cathode body 4 ; The first terminal of a second capacitor _C1 is also coupled to the first terminal of the protection resistor 724, the second terminal of the protection resistor 724 is coupled to the first terminal of the relay 25, and the second terminal of the _second second capacitor C2 It is coupled to the second terminal of the relay 25 and grounded.

The working process of this embodiment is as follows: the suction duct 1 is located at the head end of the thruster, collects air as the ultra-low orbit satellite moves forward, and captures, compresses and stores the rare air at the tail of the suction duct 1; Valve 2 controls the flow of air entering the discharge chamber from suction duct 1, thereby achieving the effect of controlling the amount of ionized air, thereby changing the magnitude of the thrust generated, and realizing the purpose of variable thrust; The outgoing air is split, and the anode and the cathode are simultaneously inhaled to make the air enter the cavity 82 of the cathode body and the cavity 83 of the anode body at the same time, and the structural dimensions of the first split outlet 32 and the second split outlet 33 can be used. The design realizes the requirements of the air intake ratio of the cavity of the anode body 5 and the cavity of the cathode body 4 under different working conditions; the cathode body 4, the anode body 5 and the power source 7 form a discharge chamber, which ionizes the air mass to form plasma; the power source 7 can The size of the discharge power can be controlled, and the size of the thrust can be further controlled. The dual control of the limiting valve 2 and the power supply 7 can effectively improve the control accuracy of the variable thrust; the plasma is under the action of the accelerating magnetic field of the magnetic coil 6 and the electric field of the discharge chamber. Accelerate the ejection to obtain thrust.

The description of the preferred embodiments of the present invention is included above, which is to describe the technical features of the present invention in detail, and is not intended to limit the content of the present invention to the specific forms described in the embodiments. Other modifications and variations of this patent are also protected. The gist of the content of the present invention is defined by the claims, rather than by the specific description of the embodiments.

Claims (10)

1. Ultra-low orbit variable thrust air-breathing magnetic plasma thruster, is characterized in that, comprises: The suction duct, located at the head end of the thruster, is provided with an air inlet and an air outlet for compressing the rarefied air and sucking it into the thruster; The discharge chamber is located at the rear end of the thruster and communicated with the air outlet. It consists of a cathode body, an anode body and a power source. The cathode body and the anode body are all electrically connected to the power source and are used to ionize the compressed thin air into plasma. Under the action of the electric field, the plasma is accelerated and ejected, and the discharge chamber is provided with an accelerating magnetic field; The restrictor valve is located between the intake duct and the discharge chamber, the intake end of the restrictor valve communicates with the air outlet, and the air outlet end of the restrictor valve communicates with the discharge chamber for controlling the compressed air entering the discharge chamber mass flow of the group. 2. The ultra-low orbit variable thrust air-breathing magnetic plasma thruster according to claim 1, further comprising a flow divider, the flow divider being located between the flow restrictor valve and the discharge chamber; The cathode body and the anode body are both hollow cylindrical structures, the cathode body is located in the cavity of the anode body, an annular cavity is enclosed between the outer wall of the cathode and the inner wall of the anode, and the discharge cavity is composed of the annular cavity and the anode. The remaining cavity in the body is formed, the outer wall of the anode body is surrounded by a magnetic coil, and the accelerating magnetic field is generated by the magnetic coil; One end of the shunt is provided with a shunt inlet, and the other end of the shunt is provided with a first shunt outlet and a second shunt outlet surrounding the first shunt outlet, the first shunt outlet and the second shunt outlet are both. It communicates with the shunt inlet through the shunt structure in the shunt; The shunt inlet is communicated with the gas outlet end of the restrictor valve, the first shunt outlet is communicated with the cavity of the cathode body, and the second shunt outlet is communicated with the annular cavity. 3. The ultra-low orbit variable thrust air-breathing magnetic plasma thruster according to claim 1, wherein the power supply comprises: The ignition circuit is electrically connected with the cathode body and the anode body for igniting the air in the discharge chamber; The main discharge circuit is electrically connected with the cathode body and the anode body for supplying an electric field to the discharge chamber. 4. The ultra-low orbit variable thrust air-breathing magnetic plasma thruster according to claim 3, wherein the ignition circuit comprises: a first charging power source for charging the first capacitor; The first capacitor includes a first terminal and a second terminal, the first terminal of the first capacitor is coupled with the anode of the first charging power source, the second terminal of the first capacitor is coupled with the cathode of the first charging power source, and the cathode body is coupled to the anode of the first charging power source. The second terminal of the first capacitor and the cathode of the first charging power source are coupled; The first silicon controlled rectifier includes a first terminal and a second terminal. The first terminal of the first silicon controlled rectifier is coupled with the first terminal of the first capacitor and the anode of the first charging power supply. The second terminal is coupled to the anode body. 5. The ultra-low orbit variable thrust air-breathing magnetic plasma thruster according to claim 3, wherein the main discharge circuit comprises a second charging power supply, a second silicon controlled rectifier, a diode, a protection resistor, a relay, n second capacitors C ₁ ˜C _n and n inductances L ₁ ˜L _n , where n is a natural number greater than 1; The second silicon controlled rectifier, the protection resistor, the relay and each second capacitor include a first terminal and a second terminal; The first terminal of the first second capacitor C ₁ is coupled to the anode of the second charging power source, and the first terminal of the i-th second capacitor C _i is coupled to the first terminal of the i+1-th second capacitor C _i+1 Through the coupling of the ith inductance Li, the second terminal of each of the second capacitors C ₁ ˜C _n is coupled with the cathode of the second charging power source, where 1≤i<n; The first terminal of the _nth second capacitor Cn is also coupled with the input end of the diode through the nth inductance _Ln , and the output end of the diode is coupled with the anode body through a matching resistor; The first terminal of the second silicon controlled rectifier is respectively coupled to the cathode of the second charging power supply and the second terminal of each of the second capacitors C ₁ ˜C _n , and the second terminal of the second silicon controlled rectifier is connected to Cathode body coupling; The first terminal of the first second capacitor _C1 is also coupled to the first terminal of the protection resistor, the second terminal of the protection resistor is coupled to the first terminal of the relay, and the second terminal of the _second second capacitor C2 is coupled to the second terminal of the relay and grounded. 6. The ultra-low orbit variable thrust air-breathing magnetic plasma thruster according to any one of claims 1 to 5, wherein the air inlet is a trumpet-shaped structure, and the air inlet is located in a trumpet-shaped The large end of the structure, the air outlet is located at the small end of the trumpet-shaped structure. 7. The ultra-low orbit variable thrust air-breathing magnetic plasma thruster according to any one of claims 1 to 5, wherein the part of the air-suction duct near the air outlet position on the air-suction duct is made of nitrogen storage, The oxygen storage solid solution material is made of material, and the remaining part of the air intake channel is made of foamed silicon carbide material, and the foamed silicon carbide material on the air intake channel is filled with carbon molecular sieves. 8 . The ultra-low orbit variable thrust air-breathing magnetic plasma thruster according to claim 7 , wherein the filling ratio of carbon molecular sieve in the foamed silicon carbide material on the air inlet is along the air inlet to the air outlet. 9 . direction gradually increases. 9 . The ultra-low orbit variable thrust air-breathing magnetic plasma thruster according to any one of claims 1 to 5 , wherein an enhanced coating is provided on the air intake duct at a position located at the air inlet. 10 . 10. The ultra-low orbit variable thrust air-breathing magnetic plasma thruster according to any one of claims 1 to 5, wherein the cathode body is made of tungsten metal material, and the anode body is made of titanium Made of metal material.

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