66 results about "Spacecraft propulsion" patented technology

The invention provides a conventional and detonation dual-mode spontaneous combustion propellant space engine, and belongs to the technical field of spacecraft propulsion systems. The conventional and detonation dual-mode spontaneous combustion propellant space engine comprises a propellant control valve, an engine head, an engine body and an external excitation device. The propellant control valve, the engine head part and the engine body part are connected in sequence, and the external excitation device is embedded into the outer wall of the engine body part; according to the invention, by adopting an external excitation mode, the space engine in a detonation mode is realized; by means of a selective excitation mode, the space engine capable of selecting a conventional mode and a detonation mode is achieved. The detonation combustion is applied to the spontaneous combustion propellant space engine, the upper limit of the combustion efficiency of the space engine of a constant-pressure thermodynamic cycle is broken through, and the combustion efficiency is further improved on the original basis.

The invention discloses a high-temperature-impact-resistant electric heating sheet for a spacecraft, and relates to the technical field of electric heating elements, and the electric heating sheet is of a three-layer composite structure. The composite film (heat preservation and insulation layer) is formed by aerogel and polyimide (PI) in an in-situ chemical bonding mode, and the polyimide composite film (heat conduction and insulation layer) is formed by Fe-Cr-Al / Ni-Cr double-layer alloy and carboxyl functionalized carbon nano tubes in a vertical orientation arrangement mode. The material has the characteristics of high thermal conductivity, wide temperature range stability, rapid thermal response and high temperature impact resistance. The preparation method comprises the steps of preparation of the heat preservation insulating layer, integration of the alloy heating layer and the heat conduction insulating layer, imidization reaction strengthening and plasma auxiliary heat treatment. The electric heating sheet can be used for thermal management of a spacecraft propulsion system or precision electronic equipment and other extreme temperature environments, the problems of high-temperature layering, response delay and the like of a traditional electric heating sheet are solved, and the thermal control reliability is improved.

The invention discloses a magnetic particle grinding and finishing machining method for the inner surface of an additive manufacturing fuel manifold. The method comprises the steps that machining equipment is assembled; blocking an oil outlet of the fuel oil header pipe; filling a magnetic grinding particle medium, plugging an oil inlet of a fuel oil main pipe, and installing the fuel oil main pipe to processing equipment to finish the addition of the aqueous grinding fluid; inclining the fuel oil header pipe to enable the lowest end to be immersed in the aqueous grinding fluid; the magnetic grinding machine is started, and the magnetic grinding particle media are driven by magnetic field force to move disorderly to grind and finish the inner surface of the fuel header pipe; a fuel oil header pipe is periodically adjusted to rotate in the magnetic grinding process; closing the magnetic grinding machine after the magnetic grinding time is ended; dismounting the fuel oil main pipe, dismounting the plug and pouring out the magnetic grinding particle medium; flushing a fuel oil main pipe; a fuel oil main pipe is dried; and demagnetizing the fuel manifold. The problem of polishing the inner surface of the integrated additive manufacturing fuel header pipe is solved, the quality of the inner surface can be greatly improved after polishing, and the use requirement of the fuel header pipe in a spacecraft propulsion system is met.

The invention 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: establishing a valve needle coupling dynamic equation containing a fluid reaction force according to a hysteresis model used for representing a nonlinear relationship between piezoelectric ceramic input voltage and output displacement; the input voltage of expected displacement of the valve needle is calculated according to the valve needle coupling dynamic equation, the calculation result is applied to the piezoelectric ceramic stack, the obtained valve needle displacement is fed back to a preset active-disturbance-rejection control model, and the displacement compensation control quantity of the valve needle is output; and respectively carrying out parameter identification and gain updating processing on the valve needle coupling dynamic equation and the active-disturbance-rejection control model, and updating a displacement compensation control quantity according to a processing result and a real-time input voltage of the spacecraft so as to carry out real-time suppression on thrust noise of the spacecraft. According to the method, the problem of thrust fluctuation caused by the fact that a traditional linear model cannot accurately represent the hysteresis characteristic of the piezoelectric ceramic can be solved, and thrust noise is effectively reduced.

The invention discloses a magnetic particle grinding and finishing processing method for an inner runner of an additive manufacturing fuel spray rod. The method comprises the following steps: assembling processing equipment; a nozzle opening of the fuel oil spray rod is blocked; filling a magnetic grinding particle medium and an aqueous grinding liquid; blocking a main oil inlet of the fuel oil spray rod; mounting the fuel spray rod to processing equipment; the magnetic grinding machine is started, and the magnetic grinding particle medium is driven to move disorderly to grind and finish the inner flow channel of the fuel oil spray rod under the magnetic force action of the rotating magnetic field; the rotation angle of the fuel spray rod is periodically adjusted in the magnetic grinding process; closing the magnetic grinding machine after the magnetic grinding time is ended; dismounting the fuel spray rod, dismounting the plug, and pouring out the magnetic grinding particle medium and the aqueous grinding fluid; flushing the fuel spray rod by using high-pressure water; the fuel oil spraying rod is dried; and demagnetizing the fuel spray rod. According to the method, the surface polishing problem of the inner flow channel of the integrated additive manufacturing fuel spray rod is solved, the surface quality of the inner flow channel can be greatly improved after polishing, and the use requirement of the fuel spray rod in a spacecraft propulsion system is met.

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.

A spacecraft 8 for propulsion utilising space debris as a fuel source. Space debris 1 can be harnessed by capturing with a manipulator arm 2. Laser 3 (e.g. robotically driven pulsed laser) can ablate the debris to produce thrust. Space junk may include defunct satellites, spent rocket stages, collision fragments, other objects posing a threat to operational satellites and spacecraft orbiting Earth. The manipulator arm may precisely position space debris towards the laser for ablation, generating thrust as an ejecta plume, in order for the spacecraft to then reach subsequent debris. If debris mass is insufficient to reach the next debris, spacecraft propulsion system 6 may be employed. Excess debris mass may be stored within the spacecraft for future use. Rotational moment may be nullified by creating a counterthrust on the debris surface.

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 present invention relates to a spacecraft propulsion system and a high-operating-condition starting method thereof in the field of spacecraft propulsion technology, comprising: a pressurized gas management module is respectively connected to the inlet of a fuel tank and an oxidant tank through a gas decompression device, a track control engine is respectively connected to the outlet of the fuel tank and the oxidant tank, a fuel gas path unidirectional flow management device and a fuel gas path valve module are sequentially arranged between the gas decompression device and the fuel tank, an oxidant gas path unidirectional flow management device and an oxidant gas path valve module are sequentially arranged in the pipeline between the gas decompression device and the oxidant tank, and pressure sensors are respectively arranged at the outlet of the gas decompression device, the outlet of the oxidant tank, the outlet of the fuel tank, the oxidant inlet of the track control engine and the fuel inlet. The present invention solves the reverse pressure problem of the gas path unidirectional management device in the process of starting the track control engine with conventional constant pressure under the high-temperature and high-pressure working conditions of the tank.

The invention relates to the technical field of spacecraft propulsion, in particular to an ion thruster high-voltage cable withstand voltage performance test system and method, and the system comprises a vacuum device, an ion thruster, a high-voltage cable, a DC power supply, a stepping motor, and a discharge monitoring system. Wherein the vacuum equipment adopts a flange binding post to realize power supply connection inside and outside the equipment; the ion thruster is mounted in the vacuum equipment in an insulated mounting manner; the direct-current power supply is a variable voltage-stabilizing direct-current power supply and is mounted outside the vacuum equipment; the high-voltage cable comprises a first cable and a second cable; the middle area of the first cable is fixed on the stepping motor in an insulated mounting manner; and the discharge monitoring system is arranged in the vacuum equipment. According to the invention, safe and reliable loading of high-voltage cable power supply is realized in a manner of loading the screen grid voltage in a stepped manner, and the risk of misjudgment of the power supply performance of the high-voltage cable caused by ion beam current flickering in the high-voltage loading process is solved.

The invention discloses an acting efficiency calculation method in a storage tank inflation process and a computer program product, which are used for calculating the acting efficiency in the storage tank inflation process, and relate to the field of propellant filling of a spacecraft propulsion system, and the method comprises the following steps: calculating a gas acting condition in the storage tank inflation process to obtain acting Q work of gas on a storage tank; a storage tank shell and internal gas are regarded as a closed system, storage tank heat energy change and gas heat energy change in the storage tank inflation process are calculated, and the sum of the storage tank heat energy change and the gas heat energy change is energy change Q heat of the whole closed system; and according to the energy change Q heat of the whole closed system and the work Q work of the gas on the storage tank, the work efficiency X of the gas is obtained. The work efficiency obtained through calculation is used for a gas supplementing link or a gas cylinder gas adding link of spacecraft propellant filling, and the actual needed filling pressure accurate value in the filling process is obtained.

The invention relates to the technical field of spacecraft propulsion, in particular to a gas collection efficiency measuring system which comprises an ion source, a Faraday probe array, a gas collection supercharging device, an electric thruster, a base and an upper computer system which are arranged in a vacuum cabin. The generated beam ions are collected and pressurized by the gas collecting and pressurizing device and then enter the electric thruster, and the electric thruster ignites to lead out the ions to generate a beam; the Faraday probe array is arranged at an inlet of the gas collecting and pressurizing device and is fixed on the base through the support frame; the gas collecting and pressurizing device and the electric thruster are fixed on the base through a support frame; and the upper computer system is electrically connected with the Faraday probe array. The mass flow rate of the gas entering the gas collecting and pressurizing device and the mass flow rate of the output gas of the gas collecting and pressurizing device can be accurately measured, so that the gas collecting efficiency of the gas collecting and pressurizing device is accurately obtained.

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 invention provides a spontaneous combustion propellant space detonation engine injector, which belongs to the technical field of spacecraft propulsion systems and comprises an injection surface and an injection structure. The injection surface bears an injection structure, and the injection structure is formed by three annularly arranged injection rings with radiuses gradually increased from inside to outside; aiming at the urgent demand of a space engine with more advanced performance indexes, a novel space detonation engine design configuration is provided, and detonation wave stable propagation of a spontaneous combustion propellant for a long time can be realized; through cooperative cooperation of multiple circles of injection holes and matching with a damping structure, the anti-quenching capacity is greatly improved compared with that of an annular detonation combustion chamber, stable propagation of detonation waves is achieved, and the duration time is larger than or equal to 5 s; the combustion efficiency in the detonation mode can be larger than or equal to 99.5% (the combustion efficiency of a conventional injector is larger than or equal to 95%), and the combustion efficiency of a space engine can be further improved compared with that of an existing space engine.

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 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 discloses a PPT body carbon deposition prevention structure and method. The PPT body carbon deposition prevention structure comprises a shell and at least one thrust chamber arranged on the shell. The thrust chamber comprises a thrust chamber frame shell, a cathode plate, an anode plate, a spark plug, a solid working medium and a carbon deposition prevention assembly, and a thrust outlet is formed in one end of the thrust chamber; the cathode plate and the anode plate are arranged on the upper inner wall and the lower inner wall of the thrust chamber frame shell; one end of the solid working medium is arranged in the thrust chamber and is in contact with the cathode plate and the anode plate; the anti-carbon assembly comprises two anti-carbon insulation sheets, the two anti-carbon insulation sheets are located on the left inner wall and the right inner wall, close to the outlet, of the thrust chamber frame shell, and a notch extending towards the interior of the thrust chamber is formed in the middle area, corresponding to one end of the thrust outlet, of each anti-carbon insulation sheet. The plasma thruster is applied to the field of spacecraft propulsion, a discharge loop can be prevented from being formed between the cathode plate and the anode plate in the long-time operation process of the pulse plasma thruster, and long-time stable and reliable operation of the thruster is ensured.

The invention discloses a high-integration cold air propeller and a satellite, and relates to the technical field of spacecraft propelling systems.According to the high-integration cold air propeller and the satellite, an independent heating cavity is arranged, and liquid ammonia must flow through the heating cavity, so that the liquid ammonia can be locally heated in the heating cavity, the situation that the heating range is too large is avoided, and the response speed can be increased; and meanwhile, the heating cavity penetrating through the storage cavity can dissipate heat into the storage cavity during heat dissipation, and therefore energy waste is reduced.

The application provides a common bottom tank propellant management device, and belongs to the field of spacecraft propellant storage and delivery. The device solves the problem of unstable propellant discharge of a small common bottom tank of a satellite. The device comprises a common bottom tank, which is internally separated into a first cabin and a second cabin by a cabin separation part. The cabin separation part is in a circular arc protruding shape in the first cabin to form an edge liquid collection area in the first cabin. A first cabin PMD is connected to the same end of all first liquid delivery parts and second liquid delivery parts. A second cabin PMD is arranged in the second cabin and is provided with a plurality of third liquid delivery parts which are uniformly arranged in a circle and are co-shaped with the inner wall of the second cabin. The first liquid delivery part, the second liquid delivery part and the third liquid delivery part are kept at a preset distance from the corresponding cabin wall to form a capillary flow space. A capillary porous plate is arranged in each window to connect the inside of the corresponding delivery part with the capillary flow space. The device is mainly used for common bottom tank propellant management.

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.

The invention discloses a spacecraft propulsion system utilizing magnetic repulsion forces for efficient thrust generation. The system comprises a rotating electromagnet configured to generate a repulsive magnetic field upon activation. A stationary solenoid positioned within the housing interacts with the rotating electromagnet, generating an opposing magnetic force. A tether mechanism connects the rotating electromagnet to a rotating member, with an adjustable angle relative to the stationary solenoid. A magnetic bearing supports the rotation of the member for minimal friction. The core functionality involves generating a non-uniform magnetic field through the rotating electromagnet. The electromagnet is activated as it approaches the stationary solenoid, creating a repulsive force. This force is deactivated before full engagement and channeled through the angled tether mechanism to produce a net forward thrust for the spacecraft. The control system manages the activation sequence, timing, and tether angle adjustments to optimize propulsion efficiency. This invention offers a novel approach to spacecraft propulsion using magnetic forces, potentially lower to more efficient and controllable spacecraft maneuvering compared to traditional methods.

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 an embedded distributed array Hall electric thruster and a method for generating thrust. The Hall electric thruster comprises a base, at least one anode and at least one cathode, and the at least one anode and the at least one cathode are of an embedded distribution structure on the base. By applying the array type Hall electric thruster, through the array type design scheme, the Hall electric thruster is changed into a modular combined structure comprising a plurality of cathodes and anodes, the anodes and the cathodes are of the same model, and batch production is easy to achieve. Through the embedded distribution design, the one-to-one correspondence relationship between the cathode and the anode is canceled, the same cathode can ignite a plurality of adjacent anodes, and the same anode can ignite a plurality of adjacent cathodes when working. And the cathodes and the anodes are backups for each other, so that when individual modules are damaged, the modules can still work normally to generate thrust, and the diversified requirements of the market on the performance of Hall electric propulsion products are met.

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.