785results 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.

The invention relates to a Hall-effect propulsion device (100) comprising: - a support body (110) extending between a first end (102) and a second end (103), the body comprising an annular outer wall (112) and an annular inner wall (113) arranged coaxially around a longitudinal axis AA' to form an annular channel (111), the first end (102) having an annular opening forming an ion outlet; - a magnetic circuit (120) arranged in the annular channel (111) and configured to generate a radial magnetic barrier in the annular channel close to the first end of the channel (102); - at least one magnetic compensation magnetic element (130) arranged in the magnetic circuit and configured to compensate for at least a fraction of an absolute value of the residual magnetic moment generated by the magnetic circuit (120).

The invention describes an air plasma propulsion system and method comprising: an air turbine plant adapted to receive supply air from an external environment, and a plasma propulsion unit adapted to increase a speed of said air after a corresponding ionization and passage into air plasma phase. The plant comprises a turbine compressing air which transits in the plasma propulsion unit which ionizes it strongly and accelerates it through a frequency or microwaves in magnetic field produced by a laser beam.

The invention discloses a pulse type electromagnetic thruster performance detection system and an analysis method. A direct-current power supply, an ignition trigger controller, a pulse current detector, a pulse high-voltage detector, a discharge waveform collector, a pulse type electromagnetic thruster, a polar plate inductance intensity measuring device, a thruster impulse measuring device and a plasma movement speed measuring device are adopted. The polar plate inductance intensity measuring device is used for measuring the inductance gradient of a polar plate discharge channel of the pulse type electromagnetic thruster under different polar plate configurations; the thruster impulse measuring device is positioned below the pulse type electromagnetic thruster and is used for acquiring an output impulse value of the thruster when the pulse type electromagnetic thruster discharges and works; the plasma motion speed measuring device is arranged at the front end of a nozzle of the pulse type electromagnetic thruster and used for measuring the motion speed of plasma plume jetted by the pulse type electromagnetic thruster. According to the invention, measurement of macroscopic performance parameters of the pulse type electromagnetic thruster is realized, and quantitative characterization of microscopic characteristics of the thruster can be realized at the same time.

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

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.

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 present application relates to a method and device for secondary autonomous restart of an ion electric propulsion system. Collect screen grid current or acceleration current; execute a single flash and hardware autonomous restart process according to the screen grid current or acceleration current, determine whether the single flash and hardware autonomous restart process is successfully executed, if not, execute the continuous flash and software autonomous restart process, if so, return to collect screen grid current or acceleration current; determine whether the continuous flash and software autonomous restart process is successfully executed, if so, the restart ends, and returns to collect screen grid current or acceleration current; if not, and the number of successful executions is no does not exceed a predetermined number of times, return to collect screen grid current or acceleration current, if not, and the number of successful executions is no exceeds a predetermined number of times, execute the continuous flash fault reporting and shutdown process of the secondary autonomous restart of the ion electric propulsion system. Realize the autonomous recovery function of the ion thruster without human intervention in orbit; try self-recovery to the maximum extent, and protect the product failure from expanding.

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.