603results about "Using plasma" patented technology

The invention relates to a propulsion system, in particular to an air-breathing radio frequency ion electric propulsion system with mixed working medium power and a multi-mode working method thereof, and aims to solve the problems that an existing air-breathing electric propulsion system only takes atmosphere as a single working medium, the atmosphere trapping efficiency is low, and the performance of a spacecraft is affected. The device comprises a propelling and gas trapping unit, a radio frequency neutralizer, a power and control unit and a xenon storage and supply unit, the propelling and gas trapping unit comprises a radio frequency ion thruster and an atmosphere trapping gas inlet channel; the power and control unit comprises a direct current module, a control module and a radio frequency module; the xenon storage and supply unit comprises a xenon bottle and a valve and sensor assembly; the control module and the valve are electrically connected with the sensor assembly. The radio frequency ion thruster is electrically connected with the radio frequency neutralizer through the radio frequency module. The first output end of the xenon storage and supply unit is connected with the gas path input end of the radio frequency neutralizer, and the second output end of the xenon storage and supply unit is connected with the gas path input end of the radio frequency ion thruster.

The invention relates to a starting time sequence control method and device for an ion electric propulsion system. According to the neutralizer ignition control time sequence, the main cathode ignition control time sequence and the discharge chamber arcing control time sequence, neutralizer arcing, main cathode arcing and discharge chamber arcing controlled by the ion electric propulsion system starting time sequence are carried out; after arc starting of the discharge chamber succeeds, beam extraction of ion electric propulsion system starting time sequence control is carried out according to the beam extraction control time sequence; and after the beam current is successfully led out, power supply out-of-tolerance judgment for starting time sequence control of the ion electric propulsion system is carried out, and if the power supply out-of-tolerance judgment passes, the ion electric propulsion system is successfully started. According to the invention, on the premise of ensuring normal starting of the ion electric propulsion system, optimization of the ignition time sequence of the thruster is realized; the purposes that the neutralizer starts arcing preferentially, the ignition time of the main cathode is shortest, the gas supply balance time of the anode is shortest, the discharge chamber starts arcing stably, the length and width range of the ignition starting time of the thruster is adjustable, and the whole starting time sequence has high robustness control are achieved.

An electrostatic thruster for producing thrust by charging gas and / or liquid particles using a high voltage potential, the electrostatic thruster comprises at least two thruster cells, wherein each thruster cell comprises: - at least one emitter, - at least one collector, wherein the at least one collector is spaced from the at least one emitter by a thruster cell gap, wherein the electrostatic thruster further comprises a power source electrically connected to the at least one emitter and the at least one collector of each thruster cell, wherein the power source is configured to apply a voltage between the at least one emitter and the at least one collector of each thruster cell, wherein the at least two thruster cells are positioned in parallel to form a thruster cell array, wherein the thruster cell array further comprises an intermediate electrode powered by a current source that generates a substantially constant current.

The invention discloses a Hall thruster and a Hall thruster regulation and control method, and relates to the technical field of power devices, and the Hall thruster comprises a hollow cathode and a Hall accelerator. The hollow cathode is used for emitting electrons, the Hall accelerator comprises an anode, a first magnet exciting coil, a second magnet exciting coil, a discharge channel and a regulation coil, the first magnet exciting coil and the second magnet exciting coil are used for generating a first magnetic field, and the first magnetic field is at least distributed in a discharge area defined by the discharge channel; the anode is used for attracting electrons emitted by the hollow cathode into a magnetic field of a discharge area for ionization and generating discharge current, and under the condition that the oscillation amplitude of the discharge current is not lower than the target amplitude, the regulation and control coil is used for generating a second magnetic field which is used for adjusting magnetic field distribution in the discharge area. The oscillation amplitude of the discharge current is lower than the target amplitude, and the excitation power supplies of the excitation coil and the regulation coil are different. The Hall thruster has a good low-frequency oscillation suppression effect.

The invention belongs to the technical field of spacecraft electric propulsion systems, and relates to an on-satellite autonomous recovery method for a non-resident fault induced by ion electric propulsion ignition, which comprises the following steps of: judging whether a fault occurring in an ignition process of an ion electric propulsion system is an ignition fault, if so, controlling the ion electric propulsion to start a discharge chamber ignition mode, and if not, starting the discharge chamber ignition mode; in the mode, after waiting for a first preset time t1, judging whether an ion electric propulsion acceleration current is within a preset range or not, if yes, successful autonomous recovery is achieved, ignition of ion electric propulsion is stopped, and after waiting for a third preset time t3, the ion electric propulsion is autonomously recovered to a normal ignition mode; otherwise, stopping ignition of the ion electric propulsion, waiting for a second preset time t2, and repeating the process for multiple times after the ion thruster is cooled.

The invention relates to the technical field of aerospace propulsion, in particular to an ion rocket engine topological structure based on a carbon fiber composite material, which comprises a grid shell, a discharge shell, a gas circuit shell, a cover plate, a discharge cylinder and a neutralizer outer cover, wherein the grid shell, the discharge shell, the gas circuit shell, the cover plate, the discharge cylinder and the neutralizer outer cover are all prepared based on a carbon fiber composite material; the grid shell is fixedly connected with the front end of the discharge shell; the rear end of the discharge shell is fixedly connected with the front end of the gas circuit shell; the cover plate is fixed at the rear end of the gas circuit shell; the discharge cylinder is arranged in the discharge shell; and the neutralizer outer cover is fixed on the side surface of the discharge shell. The outstanding advantages of good mechanical property, excellent conductivity, small expansion coefficient and low density of the carbon fiber composite material are fully utilized, and the technical problems of high-temperature working reliability, performance level satisfaction and stability of the medium-high power ion rocket engine are solved.

A gas distributor for improving ionization efficiency of a radio frequency ion thruster, comprising a gas supply pipeline, a cavity, a flow guide ball, radial rotational flow holes, axial diffusion holes and a flow guide ball limiter, the gas supply pipeline and the cavity are integrally formed, the radial rotational flow holes and the axial diffusion holes are dispersedly arranged in the cavity, the flow guide ball is filled into the cavity through the gas supply pipeline, the flow guide ball limiter is arranged in the cavity to prevent relative movement of the flow guide ball in the cavity, and the gas distributor composed of the gas supply pipeline, the cavity, the flow guide ball, the radial rotational flow holes, the axial diffusion holes and the flow guide ball limiter is arranged at the bottom of a discharge chamber of the thruster.

A conductive liquid-fed pulsed plasma thruster includes a first electrode having a conductive solid portion and a conductive liquid portion, a second electrode separated from the first electrode to define an ignition space therebetween, at least one electric insulator separating the first and second electrodes, and a conductive-liquid passage extending within the conductive solid portion through which the conductive liquid portion flows from an inlet to an outlet located at the ignition space. The first and second electrodes are configured so that a drop of the conductive liquid portion forms and grows at the outlet when the conductive liquid portion flows through the conductive liquid passage until the drop of the conductive liquid causes an arc discharge between the drop and the second electrode that ignites the drop to produce a plasma cloud that generates thrust when exhausted.

The invention discloses a magnetic circuit structure for adjusting the gradient of a magnetic field in a Hall thruster channel, relates to the technical field of Hall thrusters, and solves the problems that a krypton working medium Hall thruster is insufficient in krypton working medium ionization and low in thruster efficiency. The device comprises magnetic poles, an inner iron core, a bottom plate, a magnetic screen, an outer magnetic conductive cover, a coil and an annular channel, the inner iron core is arranged at one end of the upper surface of the bottom plate, the outer magnetic conductive cover is arranged at the other end of the upper surface of the bottom plate, and the magnetic poles are arranged at the upper ends of the inner iron core and the outer magnetic conductive cover; a magnetic screen is arranged on the bottom plate and comprises an inner magnetic screen and an outer magnetic screen, and an annular channel is arranged between the inner magnetic screen and the outer magnetic screen; a coil is arranged between the inner magnetic screen and the inner iron core, and a coil is arranged between the outer magnetic screen and the outer magnetic conductive cover; and grooves are formed in the inner magnetic screen and the outer magnetic screen. By adjusting the thicknesses of different positions of the magnetic screen, the magnetic flux density is saturated, so that the magnetic field distribution in the channel is changed, the magnetic field gradient is reduced, the length of an ionization area is increased, and the krypton working medium is ionized more sufficiently in the channel.

The invention provides an array type composite emitter ionic liquid electrospray thruster containing a far-end electrode, and relates to the technical field of spaceflight thrusters, an insulating material is adopted to process a composite emitter assembly, so that a double charge layer is generated at an ionic liquid electrode component, electrochemical reaction cannot be generated at an inner flow channel of a composite emitter, and the composite emitter assembly is made to be uniform. The problem that an electrochemical reaction product is directly generated in the composite emitter to block an internal flow channel of the emitter is avoided, and through the arrangement of a transverse flow channel, a vertical flow channel and a liquid collecting tank, a zigzag liquid flow channel is formed, so that the electrochemical reaction product generated at the electrode component is more uniform and uniform. When the ionic liquid flows, the ionic liquid is limited at the position of the electrode component by the flow channel wall surface and cannot enter the inlet of the base flow channel of the composite emitter assembly along with the ionic liquid, so that the influence of electrochemical reaction on the work of the composite emitter is prevented.

The invention discloses a microminiature modularized electrospray thruster, and belongs to the field of application of electrospray thrusters. The thrusters are installed in the thruster module frame, each thruster works independently, the thrusters have a simultaneous cooperative working mode, the emission polarities of the adjacent thrusters in the thrusters are opposite, and each thruster periodically changes the emission polarity. And controlling the high-voltage power supply connected with each thruster to alternately output positive and negative voltages according to a preset frequency, and due to positive and negative alternation of the voltages, the charge quantities of the charged particles ejected by the thrusters are the same, and the ejected plume is self-neutralized. Under the condition that the performance of an original thruster is not lost, the microminiature electrospray thrusters are matched with one another to be assembled in a modularized and arrayed mode, each microminiature thruster is independently powered to work, the working range of the thruster can be further widened, the adjusting precision of the thruster can be further improved, and the thrust requirement of the electrospray thruster is met. Under the same inventive concept, the invention provides two technical schemes of four thruster layouts and nine thruster layouts.

A direct converter for an axisymmetric mirror confinement system provides a set of radially symmetric vanes charged to capture ions escaping along the axis of the confinement system and to convert their energy to electrical power. An electron trap positioned before the charged vanes uses a magnetic field to divert and collect electrons, separating them from the ions, and may support a radial electric field providing plasma control. The charged vanes may be constructed of or have a coating of a getter material absorbing neutrals derived from those ions after capture.

The invention discloses a pulse type electromagnetic thruster plume characteristic analysis system and method, and the system employs a vacuum chamber, a current monitoring system, a multi-probe data collection device, an ion current measurement circuit, and a data collection and processing system. The multi-probe data acquisition device is used for acquiring plume plasma saturated ion current of the pulse type electromagnetic thruster and outputting the plume plasma saturated ion current as a voltage signal; the ion current measuring circuit is used for converting the voltage signal acquired by the multi-probe data acquisition device into an ion current signal; and the data acquisition and processing system is used for collecting the voltage signal output by the multi-probe data acquisition device and the ion current signal converted by the ion current measuring circuit, converting the ion current signal into a digital signal, performing fitting processing on the ion current signal, and analyzing plume characteristics of the pulse type electromagnetic thruster with multiple ion groups superposed. According to the method, a more accurate theoretical basis is provided for performance optimization of the thruster.

The invention belongs to the technical field of ionic wind propellers, and particularly relates to a multi-stage ionic wind propelling device with a common electrode structure, which comprises a plurality of ionic wind generating units which are sequentially arranged along the airflow direction, each unit consists of a conductive support structure and an ionization needle tip on the conductive support structure, and the ionic wind generating units are used for ionizing air under the action of a high-voltage electric field; the multiple units are isolated from one another through the insulation supporting structures and are connected in series, the power supply access module alternately provides positive and negative high-voltage power supplies for all the units, all the conductive supporting structures sequentially serve as ionization electrodes and receiving electrodes in a multi-stage electric field, and the electrode sharing type continuous acceleration structure is achieved. An independent receiving end does not need to be arranged for each stage, the number of electrodes and supporting parts is remarkably reduced, the ion acceleration path is prolonged, the thrust density is improved, a frame-type or non-frame-type structure can be adopted for flexible construction according to requirements, and the advantages of being light in structure, high in thrust-weight ratio and high in expansibility are achieved; the invention is suitable for the fields of micro aircrafts and directional air supply.

The invention provides liquid metal and application thereof, a cathode material and a micro cathode arc thruster, and belongs to the technical field of alloy materials. The liquid metal provided by the invention comprises a liquid gallium alloy and metal particles, the mass ratio of the liquid gallium alloy to the metal particles is (3.78-4.62): (0.3-6); the metal particles comprise one or more of titanium particles, silver particles, gold particles, copper particles and aluminum particles. The surface tension of the liquid gallium alloy is reduced and the viscosity of the liquid gallium alloy is improved by adopting the metal particles, the metal particles can overcome slight vibration, are adsorbed on the surface of the metal feeding pipe and do not flow willfully, the metal particles serve as a cathode working medium of the micro-cathode arc thruster, stable work can be achieved, and the service life of the thruster is prolonged.

The invention provides a Hall and ion hybrid electric propulsion system. The Hall and ion hybrid electric propulsion system comprises a composite material gas cylinder, an inflation valve, a pressure adjusting module, a Hall flow adjusting module, an ion flow adjusting module, a Hall thruster, an ion thruster, an electric propulsion control unit, a power processing unit and a power supply processing and control unit, the pressure adjusting module supplies gas to the Hall flow adjusting module and the ion flow adjusting module in a time-sharing manner; the Hall flow regulation module and the ion flow regulation module are used for carrying out flow control on the Hall thruster and the ion thruster respectively; the electric push control unit performs time-sharing power supply control on the power processing unit and the power supply processing and control unit; the power processing unit and the power supply processing and control unit respectively provide electric energy for the Hall thruster and the ion thruster; the Hall thruster and the ion thruster provide thrust or impulse for the spacecraft; according to the Hall and ion hybrid electric propulsion system, the integrated design of air supply, power supply and autonomous process control of the Hall and ion thrusters is achieved, system air path and circuit configuration is simplified, and the requirement for a spacecraft with high specific impulse and high thrust electric propulsion systems at the same time is met.

The invention provides a wide-adjustment high-reliability electric propulsion system and a control method. The wide-adjustment high-reliability electric propulsion system comprises a pressure adjustment module, a flow adjustment module and a Hall thruster. The flow adjusting module is composed of an electromagnetic valve and a flow distributor. According to the Hall thruster flow distributor, through different previous series-parallel connection relations, on the premise that flow distributors are not added, the Hall thrusters can work in pairs, meanwhile, the backup number of the cathode flow distributor and the anode flow distributor is increased, and the system reliability is improved. By arranging the flow resistance step of the flow distributor, the large-range flow adjusting capacity of the cathode and anode inlets of the Hall thruster can be obtained, and therefore the large-range variable thrust working mode, the cathode large-flow ignition mode and the anode low-flow starting mode of the Hall thruster are achieved.

A vehicle with vehicle traction enhancement system providing increased tractive force is disclosed. The traction enhancement system includes a thruster. The thruster includes a prime mover and an air pressure generator. The traction enhancement system further includes an energy storage to provide energy for the prime mover. The thruster is mounted on vehicle chassis structure, body, and / or suspension, substantially vertical to the vehicle to provide an upward thrust to the vehicle to increase the reactive downward normal force to provide the enhanced tractive force.

The invention belongs to the field of space electric propulsion, and particularly relates to a working medium-free hot cathode for electric propulsion, which comprises an emitter material, an emitter, a support structure, N heat shields, an extraction grid, a base and a shell, the emitter is used for generating heat and heating the emitter material, and the emitter material is used for emitting electrons after being heated; the supporting structure is used for reducing heat dissipation of the emitter; the N heat shields are used for reducing heat radiated outwards by the emitter; a cathode emitter heater and an emitter material are integrated, and the emitter material can reach the temperature enough to meet the emission current under the heating power as much as possible through the heat shield heat insulation design; and a leading-out gate structure is designed, so that a certain electric field can be applied to the surface of the emitter, the emission current density of the emitter is improved, the space charge effect is relieved, and the hindrance that electrons emitted by the emitter are led out to the plume of the thruster to participate in the neutralization process is reduced.

One or more embodiments relates to an air-breathing plasma thruster including a thruster wall, an anode, a cathode, and at least one ring electrode. The thruster wall defines a cylindrical channel, the cylindrical channel having a first end and an opposing second end in fluid communication with the first end, where the cylindrical channel is adapted to receive incoming airflow. The anode is at the first end of the channel and the cathode is at the second end of the channel opposite the first end. The at least one ring electrode is positioned on the thruster wall.

The invention belongs to the field of space electric propulsion, and particularly relates to a working medium-free cathode neutralization circuit for a microwave ion thruster. The circuit comprises an ion thruster, a bias power supply and a cathode power supply circuit, the ion thruster is used for coupling discharge to generate ion plume, and the cathode power supply circuit is used for releasing electrons to neutralize the ion plume generated by the ion thruster; the bias power supply is used for adjusting the potential difference between the cathode and the ion thruster; the positive electrode of the bias power supply is connected with the ion thruster, and the negative electrode of the bias power supply is connected with the cathode power supply circuit, so that high-precision neutralization of ion plume of the thruster is realized, the problem of thrust eccentricity of a reasoning system caused by insufficient neutralization precision of the ion plume of the thruster is solved, and meanwhile, the service life of the cathode is prolonged.

The invention belongs to the field of space electric propulsion, and particularly relates to a working medium-free cathode heating method, which comprises the following steps: heating a working medium-free cathode by adopting a constant voltage mode and a constant power mode, dividing the heating process into a plurality of stages, calculating the heating voltage according to the expected power of each stage, and heating for a period of time under corresponding heating; after multi-stage constant-voltage heating is completed, a constant-power heating mode is switched to, the working-medium-free cathode heating voltage is adjusted, and the adjusted working-medium-free cathode heating power is made to be equal to the set rated heating power; according to the method, the thermal instability of working medium-free cathode heating can be effectively reduced, and the working performance of the cathode is enhanced; in the constant-power heating mode, as the input heating power is always a constant value, the temperature of the working-medium-free cathode tends to be stable due to the zeroth law of thermodynamics, namely the thermal equilibrium law, as time goes on, and then thermal instability and noise caused by the thermal performance of the working-medium-free cathode are reduced.

The invention discloses an ion wind aircraft manufactured in a paper folding mode, an ion propulsion flying micro-robot is designed and manufactured by innovatively applying the paper folding mode, the assembling time is greatly shortened, a middle layer of the structure adopts a pre-impregnated unidirectional carbon fiber plate, and the pre-impregnated unidirectional carbon fiber plate can be converted into an extremely light and extremely hard substance after being thermally cured; and a PI film is used as a connecting material, so that the paper folding ionic wind aircraft which is flexible, high in strength, light in weight, low in cost and high in pushing ratio is constructed.

The invention relates to the technical field of superconducting magnets, and discloses a high-reliability and low-heat-leakage propeller magnet system which comprises a Dewar, a coil assembly and a plurality of sets of heat insulation supporting structures, the coil assembly is arranged in the Dewar and connected with the Dewar through the heat insulation supporting structures, the coil assembly comprises a coil framework, and the coil framework is arranged in the Dewar and connected with the Dewar through the heat insulation supporting structures. The coil body is wound on the surface of the coil framework, the heat insulation supporting structure comprises a hot ring, a cold ring and a pull rod, the two ends of the pull rod are connected with the hot ring and the cold ring respectively, the hot ring is connected with the Dewar, and the cold ring is connected with the coil framework. Through the integrally designed heat insulation supporting structure, the number of parts of the magnet structure is remarkably reduced, the supporting stability is improved, it is guaranteed that the center line of a magnetic field coincides with the center line of plume of the propeller, and heat leakage of the magnet is further reduced.

The invention relates to the technical field of spaceflight electric propulsion, in particular to a miniature light-weight electric thruster gas circuit insulation device which comprises a first gas circuit connector, a second gas circuit connector and a ceramic gas circuit main insulator. A spiral slender air supply channel is arranged in an inner cavity of the shell; the first gas path joint is welded at the front end of the ceramic gas path main insulator, and the second gas path joint is welded at the rear end of the ceramic gas path main insulator; working medium gas sequentially passes through the first gas path connector, the gas supply channel of the ceramic gas path main insulator and the second gas path connector and enters the electric thruster. According to the electric thruster, the ceramic insulation material spiral slender inner cavity air supply channel can be directly integrally formed and manufactured, the structural complexity of the electric thruster is remarkably reduced, the insulation and pressure resistance of the air path insulation device is improved on the basis of light weight and miniaturization, and the on-orbit reliability of the electric thruster is improved.

The present invention relates to a spacecraft electric thruster particularly suitable for both propellant-fed and air-breathing electric space propulsion comprising an ionization chamber (11) for plasma generation which develops along a central development axis (X), wherein the ionization chamber (11) has a chamber inlet (12) configured to allow gas inflow and a chamber outlet (13) configured to allow extraction of ionized gas particles; a power injection assembly (30) configured to perform electron heating for plasma generation inside the ionization chamber (11); a magnetic assembly (20) configured to generate a magnetic field for plasma confinement in the ionization chamber (11); and an ion extraction and acceleration assembly (40) configured to extract ionized gas particles from the ionization chamber (11) through the chamber outlet (13) and accelerate them along an ion acceleration direction (Z), wherein the power injection assembly (30) comprises a microwave power delivery system (31) configured to deliver an electric field oscillating at a selected microwave frequency (f) and to inject the generated oscillating electric field in the ionization chamber (11), and wherein the ionization chamber (11) comprises a plurality of walls (11a, 15,16,41) closed to the passage of microwaves apart from a portion of at least one lateral wall (11a) of the plurality of walls (11a, 15,16,41) at which the power delivery system (31) is coupled to inject the generated oscillating electric field in the ionization chamber (11), and is resonant for electromagnetic fields oscillating at the selected microwave frequency.

The invention relates to the technical field of space electric propulsion, in particular to a high-integration-level radio frequency ion electric propulsion system and a control method thereof. The system comprises an upper computer, a bus power supply, a working medium gas storage and supply unit, a power supply processing and control unit, a thruster impedance matcher, a radio frequency ion thruster, a neutralizer impedance matcher and a radio frequency neutralizer. According to the radio frequency ion electric propulsion system, the application requirements and the working characteristics of the radio frequency ion electric propulsion system are combined, all the single machines form a complete high-integration-level system, the problems that all the single machines are distributed in a scattered mode, and the system compatibility is poor due to interface conversion are solved, and the application reliability is enhanced; a pure radio frequency excitation scheme is adopted, so that the task response rate can be improved; the thruster impedance matcher and the neutralizer impedance matcher are used as an independent single machine, so that the control complexity of the system can be reduced.

The invention belongs to the field of space electric propulsion, and particularly discloses a field reflection plasma thruster which comprises a first coil and a second coil. The first coil and the second coil are coaxially and radially nested; the first coil and the second coil are sequentially discharged in different current directions to respectively generate a first magnetic field and a second magnetic field which are opposite in magnetic field direction, the first magnetic field and the second magnetic field are used for acting on pre-ionized plasmas to form field reverse plasmas, and the second magnetic field acts on the field reverse plasmas to enable the field reverse plasmas to be ejected in an accelerated manner; and the pre-ionized plasma is arranged in an in-axis gap of a coil nested on the inner side in the first coil and the second coil. According to the invention, the bidirectional discharge circuit topology is adopted, so that energy loss is avoided, and the efficiency of the propeller is improved.

The invention relates to a method and a device for controlling consistency of double ignition performance of an ion electric propulsion system. The method comprises the following steps: acquiring a thermal throttler, an anode power supply and a screen grid power supply which are used for controlling the double-firing ignition performance of the ion electric propulsion system; the heat throttlers are screened, so that the gas supply performance consistency control of the double-engine ignition performance control of the ion electric propulsion system is realized; through the anode power supply and the screen grid power supply, power supply performance consistency control and software consistency control of double-firing ignition performance control of the ion electric propulsion system are realized. According to the dual-engine thruster performance consistency control method and system, the purpose of dual-engine thruster performance consistency is comprehensively achieved by improving air supply performance consistency control, improving power supply performance consistency control and improving software consistency control, finally, the satellite disturbance torque is minimum, the propulsion efficiency is highest, and the whole satellite application efficiency is improved.

The invention discloses a method for starting a touch holding current of a hollow cathode by sparking a Hall thruster, and relates to the field of sparking of spaceflight electric thrusters. The problem that in the prior art, a technology for restraining the thruster from being extinguished due to ignition is lacked, so that the cathode needs to be heated and ignited again, and the thruster needs to be restarted is solved. The method is used for transferring pulse energy of abrupt change current generated during ignition between an anode of an anode power supply of the thruster and an anode of the thruster in a power supply system to a contact pole of the hollow cathode in a self-sustaining state of the hollow cathode of the thruster, so that extinguishing of the Hall thruster is inhibited. The method is mainly used for being applied to the hollow cathode self-sustaining state in the stable operation stage after successful ignition of the Hall thruster.