159results about "Using plasma" patented technology

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.

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 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 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 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.

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.

Disclosed is a Hall-effect propulsion system comprising a Hall-effect thruster (10), a cathode (40), and an electrical power supply (80) including an engine power supply (81) delivering an engine discharge current (Id) to the cathode (40) allowing an engine discharge (DM) to be created in a gas contained in the Hall-effect thruster (10), and a secondary power supply (82) delivering a secondary discharge current (Ik) to the cathode (40). The Hall-effect propulsion system comprises means (90) for controlling the electrical power supply (80), which are configured to command delivery of a secondary discharge current (Ik) by the secondary power supply (82) to the cathode (40) in order to keep the cathode (40) at a minimum operating temperature between two activations of the engine power supply (81).

The invention discloses an integrated electrospray thruster transmitting module based on a 3D printing technology, and belongs to the technical field of spaceflight. The integrated electrospray thruster transmitting module is an integrated component which is integrally formed through a micro three-dimensional photoetching 3D printing technology, and a continuous fluid transmission channel from a liquid supply connector at the bottom end to a transmitting nozzle at the top end is arranged in the integrated electrospray thruster transmitting module; in the fluid flowing direction, the integrated electrospray thruster transmitting module sequentially comprises a liquid supply interface layer, a multi-layer flow channel structure and a transmitting layer from bottom to top. Integrated forming and high integration of the complex micro-channel structure are achieved, the array liquid supply uniformity is improved through the forked micro-channel structure, the size of sprayed liquid drops and the single-hole flow are reduced, and therefore emitter miniaturization is achieved, and the overall emission performance and working stability of the electrospray thruster are improved.

A system for heat conditioning an area of Earth includes an Earth-orbiting satellite. The satellite includes a power supply, a precursor gas supply, and one or more double helicon plasma beam generators coupled to the power supply and the gas supply and configured to generate a plasma and further configured with a magnetic nozzle to maintain a shape of the beam; therefore, the top surface area of the beam is maximized. The generated plasma provides enhanced electromagnetic waves absorption, reflection, and deflection of incoming solar light and electromagnetic radiation, thereby reducing the heat striking the area of the Earth.

The invention relates to a method (100) for controlling an ion thruster, the ion thruster comprising an emission electrode, an extraction electrode and a conductive liquid which is deposited on the emission electrode, the ion thruster being suitable for emitting an ion beam when an electric field is applied to the conductive liquid, the ion beam being suitable for providing thrust to the thruster, the thrust depending on an emission current l em and an ion emission speed, characterised in that the method comprises the following steps: - adjusting (102) the emission current to a setpoint value l c by applying a threshold emission potential Vthresh to the emission electrode by means of a current generator; and - when the setpoint value l c of the emission current is reached, adjusting (104) the emission speed by applying an extraction potential Vext to the extraction electrode by means of a voltage generator in order to bring the emission potential Vem to a predetermined value Vempr = Vthresh + Vext. The invention also relates to an ion thruster system (1) which is suitable for implementing the method (100) according to the invention.

The invention discloses a contact pole gap additional ventilation type hollow cathode, which belongs to the technical field of spaceflight propulsion, and comprises a cathode assembly and a contact pole assembly, and the core is improved as follows: the contact pole assembly consists of a porous contact pole tube, a gas distributor and a contact pole ventilation ring which are mutually nested to form an independent gas supply system; after being subjected to pressure equalizing by the gas distributor and rectification by the contact pole ventilation ring, a working medium is uniformly injected into an annular gap between the contact pole and the cathode tube through circumferential ventilation holes of the contact pole tube with holes. The downstream gas density of the throttling hole is remarkably improved, discharge oscillation and ion bombardment corrosion are restrained, the gas supplementing amount can be dynamically adjusted and controlled, and the device is suitable for different working conditions. The cathode is simple and reliable in structure, highly compatible with an existing cathode, capable of prolonging the service life of the cathode and widening the stable operation interval, and suitable for long-service-life aerospace tasks of various electric propulsion systems.

A Hall effect thruster system (12) includes a thruster body (52) and a diffuser (1) configured to apply a torque (Td) to the thruster body (52) during operation. The diffuser (1) applies the torque (Td) by ejecting the propellant in a non-axial direction, such as a direction tangent to helical or curvilinear channels (8) formed within a body (10) of the diffuser (1). The applied torque (Td) can be used to counteract a swirl torque (Tswirl) that is induced on the thruster body (52) by the ionizing Hall current flowing in an annular channel (16) of the thruster body (52). This effective counter-torque (Td) is useful in deep space applications outside the Earth's magnetic field.

The invention discloses a Beta radioactive isotope coating and magnetic confinement enhanced air-breathing electric propulsion ionization method. The method is characterized by comprising the following steps: selecting a plasma generation assembly without electrode discharge; the ionization is enhanced by adopting Beta radioactive isotope; arranging a magnetic field in the beta ray radiation direction to carry out magnetic field constraint; by dynamically adjusting the magnetic induction intensity and regulating and controlling the plasma state and the ionization rate, the Beta radioactive isotope (such as nickel-63) is creatively adopted, and the device has the advantages that the service life is long, the protection requirement is low, and mechanical compression and heating cannot induce nuclear reaction. And the requirement on the activity of the radioactive source is fundamentally reduced, so that the material cost and the safety risk are greatly reduced, and the practicability and the safety of the electric propulsion system are improved.

The present invention relates to an ion source (1), in particular an ion thruster for propelling a spacecraft, comprising a reservoir (2) for a propellant (3) that has a solid state and can be liquefied into a liquid state, a heater (4) for liquefying the propellant (3) in the reservoir (2), an emitter (5) in fluid communication with the reservoir (2) for receiving a stream (S) of liquefied propellant (3) from the reservoir (2), an extractor (7) facing the emitter (5) for extracting ions (3') of the liquefied propellant (3) from the emitter (5) and accelerating the extracted ions (3') away from the emitter (5), wherein at least one baffle (11) is arranged upstream of the emitter (3) in the stream (S) of the propellant (3) to the emitter (5).

The present invention relates to the use of natural uranium hexafluoride UF6 as propellant in an electric propulsion system.

The utility model discloses a Hall thruster ceramic channel heat dissipation structure, and relates to the field of Hall thruster heat dissipation. The problem that the temperature of the ceramic wall face of a discharge channel of an existing Hall thruster is too high is solved. The device comprises a ceramic discharge channel, an annular heat pipe, an inner magnetic pole, a bottom plate, an inner magnetic screen, an outer magnetic pole, an outer coil frame and an outer magnetic screen. The heat pipe is installed between the ceramic channel and the inner magnetic screen, one end of the heat pipe is directly contacted with the wall surface of the ceramic channel, and the other end is directly contacted with the bottom plate. According to the utility model, the temperature of the ceramic wall surface is conducted to the bottom plate by using the principle of phase change in the heat pipe, and then the bottom plate directly radiates and transfers the heat to space, thereby achieving the purpose of reducing the temperature of the ceramic wall surface.

The application relates to the technical field of electric propulsion, in particular to a high-structural-stability grid assembly, which comprises a screen grid, a ceramic grid sheet, an accelerating grid and a mounting assembly, wherein the ceramic grid sheet is arranged between the screen grid and the accelerating grid; the grid holes of the ceramic grid sheet are arranged in correspondence with the grid holes of the screen grid and the accelerating grid; the screen grid, the ceramic grid sheet and the accelerating grid are fixedly installed together through the mounting assembly to form a solid grid assembly. The application adopts the ceramic grid sheet with high strength and high density, combines the grid assembly into a solid whole, improves the mechanical resistance, and sets the ceramic grid sheet between the grids to prevent the opposite deformation due to thermal stress during the working process of the grid assembly. Compared with the existing products, the force and thermal environmental stability of the grid assembly is remarkably improved.

The invention provides a Hall thruster hollow cathode ignition circuit and a control method, Hall thruster hollow cathode ignition adopts two selectable cathode ignition modes of a cathode power supply and an anode power supply, the cathode power supply comprises a constant current source for cathode heating and a voltage source for cathode ignition, the anode power supply is mainly used for supplying power to an electric thruster anode, and the cathode power supply is mainly used for supplying power to an electric thruster cathode. Cathode ignition power supply can be achieved by connecting the change-over switch and the cathode power source in parallel. Through time sequence design, automatic switching of two cathode ignition modes can be achieved, and timing output of thrust of the Hall thruster is achieved. According to the ignition circuit and the control method, the reliability of cathode ignition of the Hall thruster is improved on the premise that a power module is not added.

The invention provides a Hall electric propulsion controller based on dual redundancy and a control method thereof. The controller comprises a first main control unit, a second main control unit, a first power processing unit, a second power processing unit, a first flow control unit and a second flow control unit; the first main control unit and the second main control unit are respectively provided with two paths of independent control signals which are respectively connected with and control the first power processing unit and the second power processing unit; the first main control unit and the second main control unit are respectively connected with and control the first flow control unit and the second flow control unit; when one of the first main control unit and the second main control unit breaks down, the other main control unit takes over and controls the first power processing unit, the second power processing unit, the first flow control unit and the second flow control unit.

The present invention relates to an ion source (1), in particular an ion thruster for propelling a spacecraft, comprising a reservoir (2) for a propellant (3), a heater (4) for heating the propellant (3), an emitter (6) having one or more projections (7) for emitting ions (3') of the propellant (3), wherein the one or more projections (7) are in fluid communication with the reservoir (2), an extractor (8) facing the emitter (6) for extracting ions (3') of the propellant (3) from the emitter (6) and for accelerating the extracted ions (3') in a direction of emission (R), wherein the heater (4) comprises a projection heating unit (5) and is configured to liquefy solid propellant (3) at the one or more projections (7) by the projection heating unit (5) before the solid propellant (3) in the reservoir (2) is liquefied.

The invention provides an ignition starting method and system for an 80mN Hall electric propulsion system, and the method comprises the steps: firstly, powering on a power processing unit, and automatically carrying out the power path self-inspection; whether ignition is conducted or not is determined according to the self-inspection result, if the self-inspection result is normal, execution continues, buffer gas tank closed-loop pressure control is started, and at the moment, the control pressure is the initial pressure; starting a cathode gas path to supply gas, and heating the cathode; whether the starting mode is soft starting or hard starting is judged, if the starting mode is soft starting, cathode ignition is conducted and then anode ignition is conducted, if the starting mode is hard starting, cathode ignition and anode ignition are conducted at the same time, and low voltage is adopted for ignition during anode ignition in the two starting modes; judging whether ignition succeeds or not after anode ignition, if not, quitting the ignition process, and if yes, executing the next step; turning off a cathode heating power supply and a cathode ignition power supply; the anode voltage is gradually increased to a rated value, and the buffer gas tank pressure is gradually increased to a rated value; enabling an ignition fault monitoring program, and completing ignition starting.

The application discloses a kind of inductively coupled plasma micro-propeller of double antenna structure, and its discharge system includes gas source, air inlet valve, antenna, discharge tube and radio frequency power supply.Gas source is directly connected with discharge tube by valve, wherein the valve can control the air intake during discharge;Then two identical helical antennas are connected in series at a certain distance and are wound on the same discharge tube, the output end of radio frequency power supply is connected to the end of antenna away from air inlet through impedance matcher, and the other end of radio frequency power supply and antenna is connected to the same ground.The application has simple structure, small space size and mass, and only needs to work stably at low power and low air intake, realizing micro-newton to millinewton level thrust.It is suitable for equipping in most of the current constellation micro-satellites without propulsion system of nanosatellite and cubic satellite.In addition, the length and spacing of the two antennas in the system are adjustable, and the size of the propeller can be adjusted according to the specific requirements of micro-satellite.

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 application discloses an electric spray thruster, relates to the technical field of electric thrusters, and can quickly position an emitter and an extraction grid, saves the mounting time of the thruster, and improves the consistency of the positions of the emitter and the extraction grid. The electric spray thruster comprises a shell, an emitter, a high-voltage electrode and an extraction grid, the shell has a containing cavity for containing propellant, and the inner wall of the shell is provided with a first limiting surface. The emitter is arranged in the containing cavity, and one end of the emitter is located below the liquid level of the propellant. The high-voltage electrode is fixedly sleeved outside the emitter, and the high-voltage electrode is arranged on the inner wall of the shell. The outer wall of the high-voltage electrode is provided with a second limiting surface which is limitedly matched with the first limiting surface and is used for limiting the relative position of the high-voltage electrode and the shell. The extraction grid is arranged at the outlet end of the shell, and one side of the extraction grid close to the containing cavity is provided with a protruding portion which extends into the containing cavity, the protruding portion is provided with a third limiting surface which is limitedly matched with the first limiting surface and is used for limiting the relative position of the extraction grid and the shell.

The present subject matter provides an electromagnetic propulsion system comprising: at least one electromagnetic thrustor configured to intake air, ionize the air to produce ionized air, pass the ionized air through at least one electromagnetic field and emit the ionized air in a first direction, to reach acceleration, thereby creating a thrust force in a second direction that is opposite to the first direction, and an electrostatic repulser surrounding the at least one electromagnetic field, configured to pass at least one inducing element therethrough and control a velocity and acceleration rate of the at least one inducing element, wherein a velocity and acceleration rate of the ionized air through the at least one electromagnetic thrustor is induced by the velocity and acceleration rate of the at least one inducing element in the electrostatic repulser. Additional embodiments of the electromagnetic propulsion system are disclosed herein.

Ionic thruster methods and apparatus for aircraft are disclosed. An example ionic thruster for aircraft includes a nozzle. The nozzle includes an outlet and an inlet, the inlet to receive fluid and containing an electrode mount. A ground electrode is disposed within the nozzle. Conducting pins are coupled to the electrode mount, each of the pins having a first end coupled to the electrode mount and a second end positioned closer to the ground electrode than the first end, the pins spaced apart from the ground electrode.

The invention provides a plug-in mounting type integrated gas circuit module for an electric propulsion system. The plug-in mounting type integrated gas circuit module aims at solving the problems that an existing gas circuit system is loose in structure, large in weight and low in reliability. The module adopts a highly integrated design, and a high-pressure sensor, a high-pressure self-locking valve, a two-stage redundant pressure reducing valve, a pressure stabilizing cavity, a low-pressure sensor, a four-way low-pressure electromagnetic valve and a four-way proportional valve are all integrated on an integrated board. A gas circuit sequentially flows through the elements and is divided into four parallel branches after pressure reduction and pressure stabilization, each branch is controlled by a low-pressure electromagnetic valve and a proportional valve, and finally, three outlets are formed by converging, so that redundancy control of the cathode and anode gas circuits is realized. Through the plug-in mounting structure and the standardized sealing of the O-shaped ring, the structure is compact and light, the assembly is simple and convenient, the reliability and the integration degree of the system are remarkably improved, and the manufacturing cost is reduced.

Device (52) for regulating the flow rate of a propellant fluid for an electric thruster, of the type of a hot capillary conduit device, comprising at least one electrically conductive capillary conduit and capable of regulating the flow rate of the propellant fluid under the action of a change in conduit temperature, characterized in that said at least one capillary conduit comprises a nickel-based alloy.