36 results about "Spacecraft propulsion" patented technology

The invention relates to the technical field of spacecraft propulsion system measurement, in particular to a device and a method for measuring response time and thrust of a microminiature thruster. A thrust frame of the device is arranged in a vacuum chamber, a to-be-tested thruster and a laser transmitter are installed on the thrust frame, and a gas source provides gas for the to-be-tested thruster. The laser receiver and the laser transmitter are arranged at an interval, and the laser receiver is used for receiving laser transmitted by the laser transmitter. Computer calculates thruster opening response time and thruster closing response time through transmission light intensity change time and control signal emission time difference and calculates thrust according to thrust data acquired by force sensor in time period from thruster starting control signal fluctuation first tending to stabilize to thruster starting control signal termination and the force device stabilizes the thrust. According to the device, laser is used for irradiating plume of the miniature thruster, the body structure of the thruster is not damaged, rapid non-contact measurement of the response time of the thruster is achieved, and steady-state thrust measurement of the thruster can also be achieved.

The invention concerns a spacecraft propellant tank designed for storing and supplying fuel to the propulsion system, as well as its application method, which is related to space technology and particularly to the design of tanks for storing liquid propellants within rocket engines. The tank in question comprises a body featuring at least one convex edge that creates a corner capillary and an intake port positioned on one of these convex edges. Additionally, the tank's interior surface is treated to be hydrophilic. The objective of this invention is to develop a spacecraft propellant tank characterized by its lightweight and straightforward design. It facilitates the creation of various propellant tank shapes while minimizing the quantity of internal components that reduce useful volume. It also ensures consistent and reliable functionality of the propellant system with continuous supply of propellant to the intake port, even in the absence of gravity and during multiple activations of the propulsion system. The claimed invention achieves its objective by implementing technical results that include enhancing the efficiency of propellant supply within the tank to the intake port, improving reliability, augmenting the useful internal volume, and offering a variety of tank shapes. The stated objectives are met in particular through the hydrophilic properties of the propellant tank's inner surface, the design of the propellant tank shape featuring at least one convex edge that forms a corner capillary, and positioning of the intake port on one of the convex edges that forms the corner capillary. The application method for the spacecraft propellant tank involves initially filling the tank with liquid propellant and pressurizing gas through the intake port prior to spacecraft launch. Subsequently, once the spacecraft is in orbit, propellant is conveyed from the tank to the propulsion system through the intake port. The movement of liquid propellant towards the intake port is facilitated by the pressure of the pressurizing gas and a localized pressure reduction generated during propellant discharge. This reduction occurs in the corner capillary near the intake port relative to the pressure of the propellant in other sections of at least one corner capillary, thus establishing a flow of propellant through the corner capillaries towards the intake port area.

The present application relates to the technical field of spacecraft propulsion system, and more particularly to a micro-newton hybrid mode propulsion system and method based on gas working substance. The system comprises a gas storage unit, a gas supply unit, a thruster unit and a propulsion system driving and control unit. The gas storage unit adopts symmetrically arranged high-pressure gas cylinders connected through high-pressure isolation valves, and in combination with the automatic balancing strategy of the control unit, effectively solves the problem of on-orbit centroid drift. The gas supply unit includes parallel cold gas propulsion branches and electric propulsion branches, and each branch adopts a mechanical pressure reducer to eliminate vibration and noise caused by electromagnetic valve action from the source, realizing super-static operation of the system. The propulsion system driving and control unit controls the micro-newton cold gas thruster and the micro-newton electric thruster to work alone or in cooperation according to the distribution strategy based on the thrust critical point, realizing the optimal comprehensive efficiency of high precision and high specific impulse in the full thrust range, and being suitable for ultra-high precision space scientific exploration missions.

The invention relates to an airship propulsion system valve pressure prediction method fusing multiple physical field parameters, and the method comprises the steps: obtaining the current orbit ephemeris data, the combination attitude angle, the sun illumination intensity and the flange wall temperature of an airship propulsion system, and calculating the effective heat flux density acting on the outer surface of a key heated part according to the current orbit ephemeris data, the combination attitude angle, the sun illumination intensity and the flange wall temperature of the airship propulsion system; according to the effective heat flux density, the combination attitude angle, the flange wall temperature and the orbit height in the current orbit ephemeris data, the predicted valve pressure of the airship propulsion system is obtained through a pre-trained pressure prediction model. According to the method, the soft measurement model driven by the physical mechanism is constructed by fusing multiple physical field parameters of the orbit, the attitude, the illumination and the wall temperature, advanced prediction and active early warning of the pressure intensity of the valve are achieved, the safety and reliability of an airship propulsion system are improved, and dependence on hardware of an internal sensor is reduced.

The present application belongs to the technical field of on-orbit application of spacecraft propulsion system, and relates to an on-orbit evaluation method and system for an electric thruster and a terminal. The on-orbit evaluation method comprises the following steps: obtaining multi-dimensional on-orbit telemetry data of the electric thruster during on-orbit operation, including electric system, thermal system, propellant delivery, and attitude control related parameters; using at least two methods of basic index calculation, neural network prediction, and data feature analysis to calculate the on-orbit performance, reliability, stability, and health status of the electric thruster; and fusing the calculation results to obtain corresponding evaluation results. Through the cooperation of multi-source data fusion and multi-dimensional algorithms, the present application realizes all-around accurate characterization of the equipment state, effectively improves the evaluation accuracy and robustness, and has engineering practical value.

The present disclosure relates to systems, devices, and methods for spacecraft propulsion. In an embodiment, the present disclosure relates to an apparatus comprising a heat exchanger body defining a plurality of propellant channels configured to contain a propellant, a central cavity configured to contain a working fluid and fluidically connected to a plurality of working fluid channels that extend along a radial dimension of the apparatus, and a nozzle fluidically connected to the plurality of propellant channels and configured to expel the propellant.

The invention provides a design method of a spherical obliquely-mounted parallel tension storage tank for a spacecraft, which comprises the following steps of: comprehensively analyzing and determining the basic structure composition of a storage tank shell and a propellant management device according to the index requirements of a spacecraft liquid propulsion system; according to the index requirements of the spacecraft liquid propulsion system, structural design optimization is conducted on the storage box shell and the propellant management device in sequence; and carrying out modeling simulation on the optimized storage tank shell and the optimized propellant management device to obtain optimal structure parameters. According to the spacecraft liquid propulsion system index requirements, the storage tank shell and the propellant management device are optimally designed, the problem that an existing traditional storage tank design method cannot be suitable for a spherical obliquely-installed parallel tension storage tank is solved, and the design efficiency of the obliquely-installed storage tank is improved; and meanwhile, the requirements of a spacecraft propulsion system on the structural space utilization rate, light weight, long service life cycle and high reliability are met.

The invention relates to the technical field of spacecraft propulsion equipment, in particular to a connector suitable for common ground of an arc thruster anode structure and a whole satellite. A thruster connector of the connector comprises a threaded sleeve, a metal sleeve, an outer conductor tube, an insulating ceramic sleeve and a jack conductor tube, the outer conductor tube, the insulating ceramic sleeve and the jack conductor tube are sequentially sleeved, the threaded sleeve is coaxially connected with one end, facing a power source, of the outer conductor tube, and one end, facing the power source, of the insulating ceramic sleeve extends out of the outer conductor tube and is spaced from the threaded sleeve. The jack conductor tube is used for inserting an anode connecting column of the arc thruster and a cable core wire of the connector, the metal sleeve is fixedly connected with the outer conductor tube, and the middle layer and the outer layer of the cable are both fixedly connected with the metal sleeve. According to the connector, parasitic inductance and impedance are greatly reduced, arcing pulse current of the arc thruster is preferentially returned to a power supply through a cable, the requirement for high temperature resistance and high voltage resistance of the connector is met, meanwhile, interference is reduced, and the requirement of a spacecraft propulsion system under the condition that an anode structure of the arc thruster and a whole satellite are in common ground is met.

The invention provides a temperature closed-loop control method and system for a nitrous oxide propulsion self-pressurization storage box, and relates to the technical field of spacecraft propulsion system.The method comprises the steps that nitrous oxide concentration, pressure, propellant temperature distribution, spacecraft task stages and propellant surplus data in the storage box are obtained, and the nitrous oxide concentration, pressure, propellant temperature distribution, spacecraft task stages and propellant surplus data are obtained; calculating the power adjustment amount of each partition of the external heating layer based on the concentration and temperature data, and generating a partition control instruction; the partition power output is adjusted by executing the instruction, and a temperature field matched with a preset condition is formed in the storage box; then cooperatively optimizing the temperature field and the pressure data to generate an optimized power distribution proportion; then, the task stage and the propellant remaining amount are processed in a rolling mode, and a temperature control parameter track is generated; and finally, fusing the power distribution result and the parameter track to form a comprehensive control strategy containing an updating instruction. According to the method, the efficiency of the nitrous oxide self-pressurization system in the propellant phase change dynamic process in an on-orbit task is improved.

The invention relates to the technical field of spacecraft propulsion systems, in particular to a propellant storage tank. The propellant storage box comprises a box body and a propellant management device, the box body is provided with an inner cavity, an air port and a liquid outlet, the air port and the liquid outlet are both communicated with the inner cavity, the air port is formed in the top end of the box body, the liquid outlet is formed in the bottom end of the box body, and the propellant management device is arranged in the inner cavity; the propellant management device comprises a liquid accumulator and a flow guide plate connected to the liquid accumulator, the liquid accumulator is arranged at the liquid outlet, the flow guide plate extends along the inner wall face of the box body, a flow guide gap is formed between the flow guide plate and the inner wall face of the box body, and the flow guide gap is gradually reduced from the top end to the bottom end of the box body. The invention provides a propellant storage tank to relieve the technical problem that in the prior art, flow guide of a flow guide plate on liquid is difficult to balance the liquid storage capacity and the transmission capacity of the liquid.

The invention relates to the technical field of spacecraft propulsion, in particular to an ion thruster grid assembly redundancy safe disposal method which comprises the following steps: judging whether grid piece lap joint and short circuit are caused by grid assembly redundancy or not; determining a safety disposal measure according to a judgment result, and executing a lap-joint short-circuit disposal measure when lap-joint and short-circuit are caused by redundant objects of the grid electrode assembly; and when the redundancy of the gate assembly does not cause lap joint and short circuit and is only in a floating state, an ion beam flicker abnormity treatment measure is executed. On the basis of not changing the ion thruster and matched single-machine software and hardware components thereof, the characteristics of changing the overlap joint state and position of the redundancy and the sputtering etching and eliminating principle of the plasma by fully utilizing the differential deformation of the grid plate in the alternate cooling and heating mode, and through the adjustment and optimization of individual working modes and processes, the working efficiency of the ion thruster is improved, and the reliability of the ion thruster is improved. Therefore, the redundant materials of the grid electrode assembly of the ion thruster can be safely disposed.

A pulsed power source (200) for a spacecraft propulsion system, the pulsed power source comprising: a voltage input (210); a plurality of capacitor banks (226, 236, 246, 256) arranged to be charged by the voltage input; pulse discharge connections for electrically coupling each capacitor bank to a load (260); a discharge switch (228, 238, 248, 258) for each capacitor bank, each discharge switch positioned on a discharge path between the capacitor bank and the pulse discharge connections; and an isolation switch (220, 230, 240, 250) for each capacitor bank, each isolation switch positioned on a charge path between the voltage input and the capacitor bank.

The invention relates to the technical field of spacecraft propulsion, in particular to a method for improving the plasma discharge stability of a discharge chamber, and provides a progressive and differentiated cathode ignition starting method, and on the basis, the plasma discharge stability of the discharge chamber is improved through plasma discharge baking of the discharge chamber and small-beam baking of a grid assembly. Rapid matching and balance of two physical processes of gas discharge and ion beam extraction in the ion thruster are realized, so that the discharge stability of plasmas in the discharge chamber is effectively improved under the condition of ensuring that the ignition starting time of the ion thruster does not exceed the standard, the sputtering etching rate of the plasmas in the discharge chamber can be reduced, and the service life of the plasmas in the discharge chamber is prolonged. And the working reliability and safety of the ion thruster can be effectively improved.

The invention provides a flexible composite film structure for a pressure stabilizing cavity of an air-breathing electric propulsion system and a preparation method, and relates to the technical field of spacecraft propulsion system materials and structures. Aiming at the requirements of low-orbit spacecraft components on stable air pressure control, light weight and space environment protection, an integrated solution that a pressure stabilizing cavity main body adopts a four-layer composite film structure of a flexible substrate, a barrier layer, a functional layer and an atomic oxygen prevention layer is provided. The flexible substrate selects polyimide to provide basic bearing and flexibility, and the barrier layer adopts a direct-current magnetron sputtering technology to deposit an aluminum film on the substrate, so that gas trapping sealing and leakage barrier are realized. The functional layer is made of carbon fiber or carbon nanotube reinforced polyimide and is formed through solution spraying and curing, the overall mechanical strength of the composite film is improved, and heat dissipation is assisted. The atomic oxygen prevention layer adopts a plasma-assisted enhanced chemical vapor deposition technology to realize preparation of a siloxane film and is used for resisting erosion of atomic oxygen in a space environment and prolonging the service life of the composite film.

The present application relates to the field of spacecraft propulsion power technology, in particular to a Danzhu mercury power method, the present application is stored in the storage room Danzhu is transmitted to the combustion chamber, and is ignited in the combustion chamber, so that the cinnabar in the combustion chamber will be converted into mercury under chemical reaction, and then the mercury will be reduced to cinnabar again, the reaction and reduction constitute a closed loop system, a large amount of gas will be generated in the combustion chamber in the whole process, and then the gas is discharged through the tail nozzle, so as to generate the thrust to drive the spacecraft flight, realize the start of the spacecraft flight, because the cinnabar in the combustion chamber will be continuously converted between cinnabar and mercury, so that when the cinnabar is used as power fuel, the power fuel is almost not lost, and then the demand of long-term space exploration of the spacecraft can be met.

The invention provides a motion time sequence control method of a spacecraft lunar surface landing process simulation test system. The spacecraft lunar surface landing process simulation test system comprises a spacecraft propulsion subsystem, a landing simulation motion subsystem, a motion feedback information acquisition and receiving device, a time sequence control subsystem, an ignition control subsystem and a plume parameter high-speed acquisition subsystem. The method comprises the following steps: acquiring a detection signal of a motion feedback information acquisition and receiving device in real time through a time sequence control subsystem; when the detection signal represents that the displacement of the landing simulation motion subsystem reaches a preset position, delay timing is started; and after the delayed timing is completed, an ignition instruction is sent to the ignition control subsystem through the time sequence control subsystem, and a time synchronization signal is sent to the plume parameter high-speed acquisition subsystem. Therefore, engine ignition and starting at the moon-touching moment of the spacecraft can be accurately realized, and the plume parameter high-speed acquisition subsystem can be controlled to synchronously acquire data.

Method for propelling a spacecraft in which at least a first propulsion device (38, 46) of a propulsion system (1) of the spacecraft is operated by combustion of a premixed flammable fluid (14) containing a hydrocarbon as fuel and N2O as oxidizer, wherein the flammable fluid (14) is stored in a storage device (12) in a premixed state and wherein the first drive device (18, 38, 46) is supplied with the flammable fluid (14) by means of a first supply line (31) arranged between the storage device (12) and the drive device (38, 46), wherein the storage device (12) and the drive device (38, 46) are in flow communication with each other by means of a piping system (28) of the first supply line (31) arranged between them, characterized by that the flammable fluid (14) is essentially completely liquid in the storage device (12), wherein the temperatures within the storage device (12) are below the boiling points of the individual components of the flammable fluid at a certain pressure, and is drawn from it in liquid form and conveyed in gaseous form to the drive device (18, 38, 46), wherein the flammable fluid (14) is vaporized within the first supply line (31) by means of at least one vaporization arrangement, the vaporization arrangement comprising a throttling device (30) which is arranged in the piping system (28) downstream of the storage device (12) and comprising at least one heat supply device (50) in which the flammable fluid is, if necessary, further vaporized and the pressure in the supply line (31), between the storage device (12) and the drive device (38, 46), is increased to a delivery pressure, and wherein the flammable fluid (14) can be stored in an intermediate volume (34) in the piping system (28) downstream of the throttling device (30) and fluctuations in the composition are homogenized by the intermediate volume (34), wherein the heat supply device (32, 50) is arranged on and / or in the intermediate volume (34) in thermal contact with the flammable fluid (14) and / or on and / or in the piping system (28) downstream of the throttling device (30).

The application discloses a spacecraft propulsion thrust noise suppression method and device, and belongs to the technical field of spacecraft propulsion system regulation and control. The method comprises the following steps: according to a hysteresis model used for representing the nonlinear relationship between the input voltage of a piezoelectric ceramic and the output displacement, a valve needle coupling dynamic equation containing fluid reaction force is established; the input voltage of the expected displacement of the valve needle is calculated according to the valve needle coupling dynamic equation, and the calculation result is applied to the piezoelectric ceramic stack, the valve needle displacement is fed back to a preset active disturbance rejection control model, and a displacement compensation control amount of the valve needle is output; the valve needle coupling dynamic equation and the active disturbance rejection control model are subjected to parameter identification and gain updating processing respectively, and the displacement compensation control amount is updated according to the processing result and the real-time input voltage of the spacecraft, so that the thrust noise of the spacecraft is suppressed in real time. The application can solve the problem of thrust fluctuation caused by the fact that the traditional linear model cannot accurately represent the hysteresis characteristic of the piezoelectric ceramic, and effectively reduces the thrust noise.

The current invention relates to a spacecraft propulsion system designed to enable travel through interplanetary and interstellar spaces. The propulsion system functions by modifying the momentum of electrons ejected from the spacecraft as they interact with external electric fields found in both interplanetary and interstellar settings. Additionally, it incorporates a mechanism to maintain the spacecraft's mass and electric charge neutrality. Such an approach enables the spacecraft to accelerate or decelerate without the traditional requirement of expelling propellant mass.

The invention relates to the technical field of spacecraft propulsion systems, in particular to a micro-Newton-level mixed mode propulsion system and method based on a gas working medium. The system comprises a gas storage unit, a gas supply unit, a thruster unit and a propelling system driving and control unit. The gas storage unit adopts high-pressure gas cylinders which are symmetrically arranged and are connected through a high-pressure isolating valve, and the problem of on-orbit mass center drifting is effectively solved in combination with an automatic balancing strategy of the control unit. The gas supply unit comprises a cold air propulsion branch and an electric propulsion branch which are connected in parallel, a mechanical pressure reducer is adopted in each branch, vibration and noise caused by action of an electromagnetic valve are eliminated from the source, and ultra-static operation of the system is achieved. And the propulsion system driving and control unit controls the micro-Newton-level cold air thruster and the micro-Newton-level electric thruster to work independently or cooperatively according to a distribution strategy based on a thrust critical point, so that the optimal comprehensive efficiency of considering high precision and high specific impulse within a full thrust range is realized, and the method is suitable for an ultrahigh-precision space scientific exploration task.

The present disclosure relates to systems, devices, and methods for spacecraft propulsion. In an embodiment, the present disclosure relates to an apparatus comprising a heat exchanger body defining a plurality of propellant channels configured to contain a propellant, a central cavity configured to contain a working fluid and fluidically connected to a plurality of working fluid channels that extend along a radial dimension of the apparatus, and a nozzle fluidically connected to the plurality of propellant channels and configured to expel the propellant.

A pulsed power source for a plasma thruster of a spacecraft propulsion system, the pulsed power source comprising a voltage input 110, a plurality of capacitor banks 126 arranged to be charged by the

The invention relates to the technical field of spacecraft propulsion systems, and provides a rocket engine injector rectifying device and method, a propulsion system and a rocket. The rocket engine injector rectifying device comprises a tangential inlet, a divergent flow channel, an arc-shaped dome liquid collecting cavity and an injection hole; the two ends of the divergent flow channel are connected with the tangential inlet and one end of the arc-shaped dome liquid collecting cavity respectively; a micro-vortex generator array is arranged on the inner wall face of the divergent flow channel, and the injection hole is connected with the other end of the arc-shaped dome liquid collecting cavity to form a propellant injection unit. Bubbles are pushed to the center of a flow channel through a centrifugal force field, and bubble retention at the near wall is avoided; the inner wall face of the divergent flow channel induces turbulence pulsation through a fin array, enhances boundary layer energy exchange, inhibits local pressure sudden drop, reduces vaporization phase change of a downstream propellant, ensures normal propellant conveying flow, improves the stability of injection flow, atomization and combustion processes, and maintains normal and reliable output thrust of an engine.

The invention relates to the technical field of spacecraft propulsion, in particular to a hollow cathode orbit environment gas atmosphere life assessment experiment system which comprises a vacuum chamber, an experiment box, a hollow cathode, an oxygen cylinder, an atomic oxygen generation device, a mixed gas cylinder and a xenon cylinder, and the experiment box is fixed in the vacuum chamber; the hollow cathode is fixed in the experiment box; the atomic oxygen generating device is arranged at the top in the vacuum cabin; the oxygen cylinder is connected with the atomic oxygen generating device through a gas pipeline; the mixed gas cylinder is arranged on one side outside the vacuum cabin and is communicated with the experiment box; and the xenon gas cylinder is arranged at the bottom end outside the vacuum cabin and is communicated with the experiment box. By arranging the vacuum gauge near the hollow cathode, the partial pressure of each component of the gas atmosphere where the hollow cathode is located can be accurately measured, and the atomic oxygen partial pressure, the nitrogen partial pressure and the oxygen partial pressure of the gas atmosphere where the hollow cathode is located can be adjusted.

The application relates to the field of spacecraft propulsion technology, in particular to a continuous adjustable electric propulsion testing method and device, the method comprising setting a test instruction change period synchronized with a satellite orbit period; constructing an atmospheric resistance interference model changing with an orbit phase and converting the atmospheric resistance interference model into a dynamic thrust demand benchmark; generating at least one dynamic thrust instruction in a sine wave, a triangular wave, a step signal or a composite waveform; using a three-dimensional interpolation matrix to solve power and gas supply parameters in real time and driving a thruster to perform wide-range dynamic adjustment; and evaluating dynamic tracking performance based on a root mean square error and a frequency response characteristic index. The application realizes high-fidelity ground reproduction of a complex interference environment of a satellite in orbit, overcomes the limitations of existing static tests and single working conditions, and provides a scientific and effective testing scheme for dynamic performance verification of a continuous adjustable electric propulsion system.

The invention relates to the technical field of spacecraft propulsion systems, in particular to a variable thrust cold air propulsion system and a control method and a valve element position calibration method thereof. The system comprises a gas cylinder, a high-pressure pressure sensor, a micro pressure module, a low-pressure pressure sensor, a buffer gas capacitor, a micro-Newton-level cold air variable thrust module and a variable thrust control circuit box. And the micro-Newton-level cold air variable thrust module is integrated with a piezoelectric driving valve, a displacement sensor and a spray pipe. According to the method, instruction thrust and pressure signals are received, and a thrust-valve element opening degree conversion model is used for calculating a valve element opening degree target value; according to the actual displacement, measured by the displacement sensor, of the valve element and the pre-calibrated valve element position parameters, the current valve element opening degree is calculated; closed-loop control is carried out with the target value as the set value and the current opening degree as the feedback value, the piezoelectric driving valve is driven to generate accurate thrust, and the method has the advantages of being short in response time, high in adjustment resolution, wide in output range and simpler in structure and control logic.

The application provides a self-learning feedforward control method and system of a storage and supply system, and belongs to the technical field of spacecraft propulsion. The method solves the problem that the compensation of the existing method deviates from the actual demand of the system and cannot guarantee stable supply of the flow in the whole life cycle. The method is as follows: a feedforward compensation branch is connected in parallel on the closed-loop feedback control branch of the proportional valve in the flow control module; during the system operation, after each depressurization module completes the inflation of the buffer tank, the current buffer tank pressure and the proportional valve opening degree output by the closed-loop feedback control branch at the corresponding moment are collected as a group of sample data during the subsequent deflation stage; based on the multiple groups of sample data, the parameters of the feedforward model used in the feedforward compensation branch are identified and updated online; and the feedforward model with the updated parameters is used for feedforward compensation of the opening degree of the proportional valve in the next inflation stage. The method is used in the field of engines.

The invention relates to a spacecraft (200) having a propulsion device (100) which comprises at least one propulsion module (10) having at least one laser system (12) for emitting a laser beam (50), and an apparatus (40), in particular a scanning apparatus, for receiving and guiding the laser beam (50) onto a region (60), in particular on an outer side (202), of the spacecraft (200), which region is designed or provided as a material source for removing material on the spacecraft (200) by means of the laser beam (50). The invention also relates to a propulsion device (100) for a spacecraft (200), and to a method for propelling a spacecraft (200).

The invention relates to the technical field of spacecraft propulsion, in particular to a multi-gas-supply system based on a Kaufman electric thruster, which comprises a passive gas inlet device, a turbo molecular pump, an electric thruster, a microwave power supply, a microwave neutralizer, a microwave cathode assembly, a gas cylinder and a controller, and is characterized in that the output end of the passive gas inlet device is connected with the turbo molecular pump; the turbo molecular pump is connected with the electric thruster through a main air inlet pipeline; the electric thruster is an electric thruster of a Kaufman structure and comprises a discharge chamber and a microwave cathode assembly. The microwave power supply is respectively connected with the microwave cathode assembly and the microwave neutralizer; the gas cylinder is connected with the main gas inlet pipeline through a flow meter and a second pipeline; and the controller is respectively connected with the turbo molecular pump and the flow meter. According to the invention, based on the Kaufman type electric thruster, a multi-gas supply mode is adopted for gas supply, so that the spacecraft can reside in the ultralow earth orbit for a long time under the condition that the environment gas condition has certain amplitude fluctuation.

The invention relates to the technical field of spacecraft propulsion, in particular to an electrodeless Lorentz force thruster pulse gas supply control system and method.The control system comprises a gas supply module, a detection module and a control module, and the gas supply module comprises a gas storage tank, a pulse deflation valve and a gas transmission pipeline; the detection module comprises a Hall sensor, a pressure sensor and a temperature sensor; the control module comprises an integrated FPGA controller and a DSP coprocessor. The invention provides a three-in-one control method of time sequence prediction, dynamic compensation and closed-loop optimization, and breaks through the core bottlenecks of a traditional propulsion system in the aspects of working medium utilization rate, synchronization precision, thrust stability, extreme working condition adaptability and the like; through strict synchronization of pulse air supply and magnetic field driving, the working medium utilization rate is greatly improved, the working medium consumption cost is reduced, and the device is suitable for long-term deep space tasks.