472results about "Using plasma" patented technology

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.

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.

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 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 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 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 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.

The invention discloses a buffering and flow stabilizing device for an iodine working medium space propulsion system, which comprises a buffering cavity, the input end of the buffering cavity is connected with the output end of an iodine storage and supply module through a pipeline, the output end of the buffering cavity is connected with the input end of a thruster module, at least one group of rectifying components is arranged in the buffering cavity, and the rectifying components are connected with the thruster module through a pipeline. A porous medium plate covering an inner side port of the output end of the buffer cavity is installed at the output end of the buffer cavity, a plurality of temperature sensors and pressure sensors are further installed in the buffer cavity, and the temperature sensors and the pressure sensors are in communication connection with a control module of the iodine working medium space propelling system; and the control module feeds back and adjusts heating parameters of the iodine storage and supply module through data collected by the temperature sensor and the pressure sensor. By means of the mode, the iodine vapor supply stability can be improved, solid particles entering the propeller are reduced, and the thrust stability is improved.

The application discloses a microwave magnetic nozzle thruster without an exciting antenna, belongs to the field of space electric propulsion, and aims at solving the problems of antenna ablation and performance drift of a conventional microwave magnetic nozzle thruster. The microwave magnetic nozzle thruster comprises a discharge chamber and a resonance chamber which are coaxially arranged, a gas inlet pipe is arranged through the center of the thruster in an axial direction, the resonance chamber is coaxially arranged outside the gas inlet pipe, the gas inlet pipe is in communication with the discharge chamber in a gas path, and the gas inlet pipe is isolated from the resonance chamber in a gas path; microwaves are fed into the resonance chamber, repeatedly reflected through a short-circuit end and an open-circuit end of the resonance chamber to generate a standing wave, so that a strong standing wave electric field is formed at the open-circuit end; the strong standing wave electric field ionizes working gas in the discharge chamber to form plasma; the plasma is axially accelerated under a characteristic magnetic field of a divergent magnetic nozzle which is externally applied, forms a plume jet and is ejected out of the thruster, so that a thrust is formed.

It is presented an electric thruster (1) comprising an inlet conduit (2) defining an inlet section (1b) of atmospheric gas; an expulsion section (1c) of atmospheric gas from said electric thruster (1); an ioniser (3) configured to generate a first magnetic field and an accelerator (4) configured to generate a second magnetic field; and an assembly (5) generating a stream of electrons which, by interacting with said magnetic fields, ionises and accelerates the atmospheric gas defining the thrust of said electric thruster.

The invention relates to the technical field of launching propulsion, in particular to an electric heating supercritical CO2 launching propulsion device.A flying body comprises an inner shell and a Laval nozzle arranged in the inner shell, an external gas storage space is formed between the narrowing section of the Laval nozzle and the inner wall of the inner shell, and an internal gas storage space is formed in the Laval nozzle; a throat pipe is arranged on an inner shell of an inlet section of the Laval nozzle and communicated with a CO2 gas source, a plug is arranged on an inner shell of an outlet section of the Laval nozzle and matched with an external gas storage space in a contact mode, and a heating resistance wire is arranged between the inner shell and the Laval nozzle. The guide structure comprises a launching cylinder or a track, and the flying body is in friction fit with the guide structure. The invention provides two electric heating supercritical CO2 propelling and launching propelling devices, the two electric heating supercritical CO2 propelling and launching propelling devices heat the supercritical CO2 gas through the electric heating system, thrust is rapidly generated through high-pressure energy of the supercritical CO2 gas, and an object can be efficiently and rapidly pushed.

The invention discloses a coil excitation Hall thruster magnetic circuit structure suitable for a krypton working medium, relates to the technical field of aerospace technology and plasma propulsion, and solves the problems of low ionization efficiency and insufficient characteristic length of an existing Hall thruster after krypton is used. A magnetic conductive inner core is arranged at one end of a bottom plate, and a magnetic conductive cover is arranged at the other end; an inner magnetic pole is arranged at the upper end of the magnetic conductive inner core; an inner coil is arranged between the inner magnetic pole and the bottom plate and corresponds to the outer side of the magnetic conductive inner core; an outer coil is arranged between the outer magnetic pole and the bottom plate and corresponds to the inner side of the magnetic conductive cover; a ceramic channel is arranged between the inner coil and the outer coil; the bottom plate comprises an inner bottom plate and an outer bottom plate, the inner bottom plate and the outer bottom plate are coaxially arranged, and a bottom plate gap is formed between the inner bottom plate and the outer bottom plate. The characteristic length of a magnetic field in a discharge channel can be remarkably prolonged, the krypton ionization efficiency is improved, and meanwhile the structure of the thruster is simplified.

A plasma generation system for generating a multilayer plasma includes a plasma generator that includes an inner electrode, an intermediate electrode surrounding the inner electrode and defining therebetween an inner plasma channel having an inner plasma outlet, and an outer electrode surrounding the intermediate electrode and defining therebetween an outer plasma channel having an outer plasma outlet; a process gas unit configured to provide a first and a second process gas inside the inner and outer plasma channels, respectively; and a power supply unit configured to energize the first process gas into a first plasma that flows along the inner plasma channel and out through the inner plasma outlet as an inner layer of the multilayer plasma, and to energize the second process gas into a second plasma that flows along the outer plasma channel and out through the outer plasma outlet as an outer layer of the multilayer plasma.

Provided are a method for controlling an artificial object, a device for controlling an artificial object, and an artificial object provided with the device for controlling an artificial object, whereby an artificial object is controlled with a high degree of freedom without requiring a special propellant. Provided is a control method for controlling an artificial object (100) in the vicinity of a planet or satellite including the earth located in a cosmic space, the control method accelerating the injection of electrons, increasing the potential of the artificial object (100) compared to a state in which no electron is emitted, promoting the capture of electrons present in the vicinity of the artificial object (100) compared to the state in which no electron is emitted, and increasing the potential of the artificial object (100) in the vicinity of a planet or satellite including the earth located in the cosmic space. The electrons captured by the accelerated injection change the momentum or / and angular momentum of the artificial object (100) by a reaction generated in accordance with the momentum imparted to the electrons after the accelerated injection.

The invention relates to a hydrazine arc thruster support in split design, and relates to the field of spacecraft supports, the hydrazine arc thruster support comprises a generator support and a thruster support, the generator support is fixedly connected with a fuel gas generator, the thruster support is fixedly connected with an arc thruster, and the generator support is made of an aluminum alloy material and is in heat conduction connection with a satellite mounting surface; the thruster bracket is made of a titanium alloy material and is in heat insulation connection with a satellite mounting surface; and one side of the generator bracket is fixedly connected with one end of the thruster bracket through a blocky bracket connecting piece. The electromagnetic valve has the advantage that the use temperature and the mounting surface temperature of the electromagnetic valve are effectively reduced.

The application provides a self-cooling air-breathing radio frequency plasma electric thruster, which comprises an air inlet channel, an ionization chamber and a nozzle, the air inlet channel is divided into a main air inlet channel and a secondary air inlet channel, the main air inlet channel is communicated with the ionization chamber, a cooling flow channel is arranged on the periphery of the ionization chamber, the secondary air inlet channel is communicated with a gas guide channel, the gas guide channel is communicated with the cooling flow channel through a cooling flow channel inlet, and the cooling flow channel is communicated with the ionization chamber through a cooling flow channel outlet; the gas taken by the main air inlet channel directly enters the ionization chamber to be ionized, the gas taken by the secondary air inlet channel is guided to the cooling flow channel inlet through the gas guide channel, enters the cooling flow channel through the cooling flow channel inlet, cools the ionization chamber, and then enters the ionization chamber through the cooling flow channel outlet to be ionized, and the two gas flows entering the ionization chamber are accelerated through the nozzle after being ionized, so that the thruster generates thrust. The cooling design provided by the application has the advantages of simple structure and high energy utilization rate.

A Hall effect thruster is provided having one or more components fabricated using additive manufacturing techniques. Additive manufacturing can be used to fabricate the magnetic flux guide and the discharge channel of the thruster. In addition, the propellant distributor for the thruster can be incorporated into the base of the discharge channel. The magnetic flux guide can be fabricated with a filament impregnated with a magnetic material and the discharge channel can be fabricated with a filament impregnated with a ceramic material. Permanent magnets can be positioned in the magnetic flux guide at locations that surround the discharge channel.

The invention belongs to the field of spaceflight plasma electric propulsion, and relates to a microwave magnetic control plasma thruster and a working method thereof. A permanent magnet is arranged on the outer side of a discharge chamber of the thruster, and the permanent magnet is arranged close to one side of a spray pipe; the permanent magnet is composed of a plurality of same strip-shaped magnet units, the strip-shaped magnet units are evenly distributed in the circumferential direction of the outer wall of the discharge chamber, and each strip-shaped magnet unit is formed by regularly arranging a plurality of permanent magnets; the permanent magnet forms a multiplexing magnetic field in the discharge chamber, and the multiplexing magnetic field is composed of an embedded magnetic mirror, an envelope magnetic mirror, a magnetic trap and a magnetic spray pipe; the plurality of embedded magnetic mirrors cling to the wall surface of the discharge chamber and are tightly arranged along the axial direction, and the outer sides of the embedded magnetic mirrors are surrounded and coated by the enveloping magnetic mirror to form a multi-magnetic mirror structure; the magnetic trap is arranged at the downstream of the multiple magnetic mirror structure; the magnetic nozzle is located downstream of the magnetic trap. A plurality of tightly arranged embedded magnetic mirrors exist in the magnetic field configuration, the exterior of the embedded magnetic mirrors is surrounded by the enveloping magnetic mirror, the axial and radial restraining capacity of the plasma is enhanced through magnetic field distribution, and the thrust performance is remarkably improved.

The application discloses a capillary type ion liquid electric spray thruster with electrically controlled thrust, and belongs to the technical field of micro-nano satellite micro propulsion. The application comprises a propellant tank, an upper shell, a constant force spring, a limiting plate, a capillary and an extraction electrode. The inner and outer walls of the capillary are covered with a dielectric layer with a micron level thickness, and the inner wall is covered with a hydrophobic layer with a nanometer level thickness. The extraction electrode is fixed through an insulating screw; and the upper shell is fixed with the propellant tank. The application is based on the principle of electric wetting, uses an electric field to control the active supply of ion liquid and simultaneously adjusts the thrust size of the thruster. The application can not only save complicated fluid control components such as high-precision pumps, valves and pipelines, but also can couple the supply process and the emission process of the ion liquid, simplify the operation process of the thruster system and facilitate the miniaturization of the thruster system. The application is suitable for resistance compensation, position keeping and attitude control of low-power micro-nano satellites.

The application discloses a kind of plasma propulsion devices, belong to plasma electric propulsion technical field, comprising: discharge chamber, spiral antenna I, center arc structure, spiral antenna II, center compensation structure, radio frequency power supply and electromagnetic coil;Center arc structure is located in discharge chamber, spiral antenna I spiral is wound in the outer wall of discharge chamber, with the position opposite of center arc structure;Center compensation structure is located in discharge chamber, spiral antenna II spiral is wound in the outer wall of discharge chamber, with the position opposite of center compensation structure;Spiral antenna I and spiral antenna II are electrically connected with radio frequency power supply respectively;Two electromagnetic coils are wound and set in the outer wall of discharge chamber, respectively located in the two ends of spiral antenna II;Center compensation structure can release neutral particles under the action of heat;The application can improve ion number density, realize high specific impulse output of propulsion device, improve propulsion efficiency.

To provide ion thruster systems for spacecraft, and more particularly, to an integrated ion thruster unit that may be configured to accommodate payload envelope dimensions associated with a spacecraft.SOLUTION: An integrated ion thruster unit for a spacecraft can include a unit housing including a thrust face, and a plurality of tiles arranged in an array on the thrust face. Each of the tiles can include an emitter and an extractor configured to provide a reference voltage with respect to the emitter. The emitter can include a plurality of tips configured to emit ions in a thrust direction from an ionic liquid propellant in response to an applied voltage.SELECTED DRAWING: Figure 5

A plasma generation system for generating a multilayer plasma includes a plasma generator that includes an inner electrode, an intermediate electrode surrounding the inner electrode and defining therebetween an inner plasma channel having an inner plasma outlet, and an outer electrode surrounding the intermediate electrode and defining therebetween an outer plasma channel having an outer plasma outlet; a process gas unit configured to provide a first and a second process gas inside the inner and outer plasma channels, respectively; and a power supply unit configured to energize the first process gas into a first plasma that flows along the inner plasma channel and out through the inner plasma outlet as an inner layer of the multilayer plasma, and to energize the second process gas into a second plasma that flows along the outer plasma channel and out through the outer plasma outlet as an outer layer of the multilayer plasma.

The invention discloses a precise adjustable electrospray thruster device which is characterized in that an emitter mounting bracket is used for fixing a porous ceramic emitter, and the fixed porous ceramic emitter can move along with the mounting bracket; the special-shaped groove in the mounting bracket is used for quickly positioning the mounting bracket; the fine adjustment of the relative position between the porous ceramic emitting electrode and the extraction electrode can be realized by adjusting a spring jackscrew in a horizontal plane and a fastening screw in a vertical plane; the fine adjustment process can be operated under a microscope to improve the adjustment precision; the shell of the electrospray thruster is designed to be of a spiral twenty-four-face structure and is machined and formed through the 3D printing technology, and the weight of shell parts is further reduced by designing a hollow cavity in the shell.

An electrospray apparatus comprising a capillary emitter.

Disclosed herein is a propulsion system comprising: a solid conductive or semiconductive cathode (130); an anode (110) having a potential difference relative to said cathode (130), said potential difference creating an electric field between said anode (110) and said cathode (130); and an insulated trigger (150) adapted to trigger an arc discharge from a point on a upper surface of said cathode (130) in pulses, when said trigger (150) and cathode (130) are substantially in a vacuum, said trigger being bounded within the cathode so that the point at which the arc is triggered is located on the upper surface of said cathode.