Accurate thrust assessment is crucial for characterizing the performance of micro-thrusters.This paper presents a comprehensive evaluation of the thrust generated by a needle-type indium field emission electric propul...Accurate thrust assessment is crucial for characterizing the performance of micro-thrusters.This paper presents a comprehensive evaluation of the thrust generated by a needle-type indium field emission electric propulsion(In-FEEP)micro-thruster using three methods based on a pendulum:direct thrust measurement,indirect plume momentum transfer and beam current diagnostics.The experimental setup utilized capacitive displacement sensors for force detection and a voice coil motor as a feedback actuator,achieving a resolution better than 0.1μN.Key performance factors such as ionization and plume divergence of ejected charged particles were also examined.The study reveals that the high applied voltage induces significant electrostatic interference,becoming the dominant source of error in direct thrust measurements.Beam current diagnostics and indirect plume momentum measurements were conducted simultaneously,showing strong agreement within a deviation of less than 0.2N across the operational thrust range.The results from all three methods are consistent within the error margins,verifying the reliability of the indirect measurement approach and the theoretical thrust model based on the electrical parameters of In-FEEP.展开更多
Hollow cathode discharges are widely used as neutralizers for the electric propulsion systems and recently developed into micro-thrusters for the small satellites.In this work,a dualemitter hollow cathode thruster is ...Hollow cathode discharges are widely used as neutralizers for the electric propulsion systems and recently developed into micro-thrusters for the small satellites.In this work,a dualemitter hollow cathode thruster is developed,which can be operated in two different modes—the neutralizer mode and the micro-thruster mode.For characterizing this kind of new device,the Langmuir probe,Faraday probe,and retarding potential analyzer are used to determine the electron temperature,electron density,ion flux,and ion energy distribution function.The operating parameters,including the thrust,and specific impulse,are also measured.A two-dimensional self-consistent extended fluid model is employed to calculate the spatial distribution of plasma parameters and the fluid field of electrons in the region around the emitters.By comparing the diagnostic and modelling results,it is found that the change in the electric field and ionization zone is the essential reason for the different performances of the device in the neutralizer and micro-thruster modes.Variation in the electric field leads to an ion acceleration effect in the micro-thruster mode;moving of the ionization zone raises the plasma pressure in the orifice region of the hollow cathode,and thus leads to enhanced plasma throttling and gas expanding effects.By analyzing the above mechanisms,the possible methods for improving this kind of hollow cathode micro-thruster are discussed.展开更多
Digital micro-thruster arrays can be used for special missions of micro/nano-satellites with the requirements of high precision and small impulse.This paper presents a novel control allocation algorithm for the digita...Digital micro-thruster arrays can be used for special missions of micro/nano-satellites with the requirements of high precision and small impulse.This paper presents a novel control allocation algorithm for the digital micro-thruster array,namely status graph based control allocation(SGBCA)algorithm,which aims at finding the optimal micro thrusters combination scheme to realize the sequential control synthesis for micro/nano-satellite during real-time orbit control tasks.A mathematical model is set up for the control allocation of this multivariate over-actuated system.Through dividing thrusters into disjoint segments by offline calculation and combining segments dynamically online to provide a sequence of the required impulse for the micro/nano-satellite,the time complexity of the control allocation algorithm decreases significantly.All levels of impulse can be generated by the digital micro thruster arrays and the service life of the arrays can be extended using the segment converting strategy proposed in this paper.The simulation indicates that the algorithm can satisfy the requirements of real-time orbit control for micro/nano-satellites.展开更多
The three-dimensional numerical simulation of two-phase plume flow of solid propellant micro-thrusters was developed.Then it was used to investigate the plume interference effect by combining the direct simulation Mon...The three-dimensional numerical simulation of two-phase plume flow of solid propellant micro-thrusters was developed.Then it was used to investigate the plume interference effect by combining the direct simulation Monte Carlo(DSMC) method for multi-component gas flow with the two-way coupling model for two-phase rarefied flow.At different space between the two micro-thrusters and different wall temperature,the plume interference effect was analyzed specifically.The results show that under the plume interference effect the gas is compressed and the flow direction is changed,which resulted in the increasing of gas pressure and temperature;solid phase made no significant effect on the flow parameters of gas phase;with the rising of the space between the two micro-thrusters,the maximum pressure decreased and the maximum temperature increased in the domain under the plume interference effect;the wall temperature could influence the temperature of the gas which is extremely close to the wall,but not the gas pressure.展开更多
The Taiji program is focused on achieving space-based gravitational wave detection in the frequency range of 0.1 mHz-1 Hz.To achieve drag-free control,Taiji satellites must be equipped with micro-thrusters that satisf...The Taiji program is focused on achieving space-based gravitational wave detection in the frequency range of 0.1 mHz-1 Hz.To achieve drag-free control,Taiji satellites must be equipped with micro-thrusters that satisfy stringent requirements,including a continuously adjustable thrust,thrust resolution of 0.1μN,thrust noise of 0.1μN·Hz^(-0.5)and response time of less than 100 ms.This paper presents the progress of a variable cold gas micro-thruster being developed for the Taiji program.A series of technologies such as a bidirectional piezoelectric drive,spherical valve core,conical nozzle seal,miniature gas chamber,high thruster integration,combined digital and analogue communication and high-frequency closed-loop thrust control were employed to achieve engineering prototypes of a low-noise and fast-response micro-thruster.Ground performance tests indicate that the micro-thruster achieved a minimum thrust of close to 1.1 nN,a thrust resolution of 0.05μN and a maximum specific impulse of 69.1 s using nitrogen gas as the working fluid.The thrust noise was less than 0.1μN·Hz^(-0.5)in the frequency band of 10 mHz-1 Hz,and the thrust response time was 140 ms.The control parameters were further optimised to achieve a flow response time of 50 ms.The results indicate that the developed micro-thruster essentially met the performance requirements for drag-free control to facilitate space-based gravitational wave detection.展开更多
In this work,a force measurement system is proposed to measure the thrust of plasma microthruster with thrust magnitude ranging from sub-micro-Newtons to hundreds micro-Newtons.The thrust measurement system uses an el...In this work,a force measurement system is proposed to measure the thrust of plasma microthruster with thrust magnitude ranging from sub-micro-Newtons to hundreds micro-Newtons.The thrust measurement system uses an elastic torsional pendulum structure with a capacitance sensor to measure the displacement,which can reflect the position change caused by the applied force perpendicular to the pendulum axis.In the open-loop mode,the steady-state thrust or the impulse of the plasma micro-thruster can be obtained from the swing of the pendulum,and in the closed-loop mode the steady-state thrust can be obtained from the feedback force that keeps the pendulum at a specific position.The thrust respond of the system was calibrated using an electrostatic weak force generation device.Experimental results show that the system can measure a thrust range from 0 to 200μN in both open-loop mode and closed-loop mode with a thrust resolution of 0.1μN,and the system can response to a pulse bit at the magnitude of 0.1 m N s generated by a micro cathode arc thruster.The background noise of the closed-loop mode is lower than that of the open-loop mode,both less than 0.1 m N/Hz in the range of 10 mHz to 5 Hz.展开更多
The applications of the micro-thrust e r and the challenges of micro-thrust measuring are introduced.The developments in measuring techniques for the micro-thrust are reviewed.Micro-thrust measu rements have previousl...The applications of the micro-thrust e r and the challenges of micro-thrust measuring are introduced.The developments in measuring techniques for the micro-thrust are reviewed.Micro-thrust measu rements have previously been made either directly by mounting thrusters to the m easurement system or indirectly by mounting a target in the direct path of the e jected propellant.Several typical direct and indirect thrust-stands are presen ted and discussed in detail to illustrate the principles.Typical calibration me thods are also expounded.Finally,the resolution,uncertainty and thrust range of each thrust-stand are given,which may be helpful for the future thrust stan d design and micro-thrusters research.展开更多
Pulsed plasma thrusters （PPT） are micro-propulsion devices used in satellites for station keeping. Conventionally the plasma discharge in a PPT is initiated by a spark plug. The primary objective of the present work...Pulsed plasma thrusters （PPT） are micro-propulsion devices used in satellites for station keeping. Conventionally the plasma discharge in a PPT is initiated by a spark plug. The primary objective of the present work was to develop and characterize a PPT that does not need a spark plug to initiate the plasma discharge. If the spark plug is eliminated, the size of the thrusters can be reduced and arrays of such thrusters can be manufactured using micro electro mechanical systems （MEMS） techniques, which can provide tremendous control authority over the satellite positioning. A parallel rail thruster was built and its performances were characterized inside a vacuum chamber to elucidate the effect of vacuum level on the performance. The electrical performance of the thruster was quantified by measuring the voltage output from a Rogowski coil, and the thrust produced by the developed thruster was estimated by measuring the force exerted by the plume on a light weight pendulum, whose deflection was measured using a laser displacement sensor. It was observed that the thruster can operate without a spark plug. In general, the performance parameters such as thrust, mass ablation, impulse bit, and specific impulse per discharge, would increase with the increase in pressure up to an optimum level due to the increase in discharge energy as well as the decrease in the total impedance of the plasma discharge. The thrust efficiency is found to be affected by the discharge energy.展开更多
In this paper, the technological progress on Chinese gravity exploration satellites is presented. Novel features such as ultra-stable structure, high accurate thermal control, drag-free and attitude control, micro-thr...In this paper, the technological progress on Chinese gravity exploration satellites is presented. Novel features such as ultra-stable structure, high accurate thermal control, drag-free and attitude control, micro-thrusters, aerodynamic configuration, the ability to perform micro-vibration analyses, microwave ranging system and mass center trimmer are described.展开更多
An ammonia self-managed vaporization propulsion (ASVP) system for micro-nano satellites is presented. Compared with a normal cold gas or liquefied gas propulsion system, a multiplex parallel sieve type vaporizer and r...An ammonia self-managed vaporization propulsion (ASVP) system for micro-nano satellites is presented. Compared with a normal cold gas or liquefied gas propulsion system, a multiplex parallel sieve type vaporizer and related vaporization control methods are put forward to achieve self-managed vaporization of liquefied propellant. The problems of high vaporization latent heat and incomplete vaporization of liquefied ammonia are solved, so that the ASVP system takes great advantage of high theoretical specific impulse and high propellant storage density. Furthermore, the ASVP operation procedure and its physical chemistry theories and mathematical models are thoroughly analyzed. An optimal strategy of thrust control is proposed with consideration of thrust performance and energy efficiency. The ground tests indicate that the ASVP system weighs 1.8 kg (with 0.34-kg liquefied ammonia propellant) and reaches a specific impulse of more than 100 s, while the power consumption is less than 10 W. The ASVP system meets multiple requirements including high specific impulse, low power consumption, easy fabrication, and uniform adjustable thrust output, and thus is suitable for micro-nano satellites.展开更多
基金Project supported by the National Key Research and Development Program of China(Grant No.2020YFC2201001)the Guangdong Major Project of Basic and Applied Basic Research(Grant No.2019B030302001)+1 种基金the National Natural Science Foundation of China(Grant Nos.12105373,12105374,and 11927812)the Science and Technology Research Project of Jiangxi Provincial Department of Education(Grant No.GJJ2402105).
文摘Accurate thrust assessment is crucial for characterizing the performance of micro-thrusters.This paper presents a comprehensive evaluation of the thrust generated by a needle-type indium field emission electric propulsion(In-FEEP)micro-thruster using three methods based on a pendulum:direct thrust measurement,indirect plume momentum transfer and beam current diagnostics.The experimental setup utilized capacitive displacement sensors for force detection and a voice coil motor as a feedback actuator,achieving a resolution better than 0.1μN.Key performance factors such as ionization and plume divergence of ejected charged particles were also examined.The study reveals that the high applied voltage induces significant electrostatic interference,becoming the dominant source of error in direct thrust measurements.Beam current diagnostics and indirect plume momentum measurements were conducted simultaneously,showing strong agreement within a deviation of less than 0.2N across the operational thrust range.The results from all three methods are consistent within the error margins,verifying the reliability of the indirect measurement approach and the theoretical thrust model based on the electrical parameters of In-FEEP.
基金supported by the National Natural Science Foundation of China(Nos.61571166,11775063 and 51736003)。
文摘Hollow cathode discharges are widely used as neutralizers for the electric propulsion systems and recently developed into micro-thrusters for the small satellites.In this work,a dualemitter hollow cathode thruster is developed,which can be operated in two different modes—the neutralizer mode and the micro-thruster mode.For characterizing this kind of new device,the Langmuir probe,Faraday probe,and retarding potential analyzer are used to determine the electron temperature,electron density,ion flux,and ion energy distribution function.The operating parameters,including the thrust,and specific impulse,are also measured.A two-dimensional self-consistent extended fluid model is employed to calculate the spatial distribution of plasma parameters and the fluid field of electrons in the region around the emitters.By comparing the diagnostic and modelling results,it is found that the change in the electric field and ionization zone is the essential reason for the different performances of the device in the neutralizer and micro-thruster modes.Variation in the electric field leads to an ion acceleration effect in the micro-thruster mode;moving of the ionization zone raises the plasma pressure in the orifice region of the hollow cathode,and thus leads to enhanced plasma throttling and gas expanding effects.By analyzing the above mechanisms,the possible methods for improving this kind of hollow cathode micro-thruster are discussed.
文摘Digital micro-thruster arrays can be used for special missions of micro/nano-satellites with the requirements of high precision and small impulse.This paper presents a novel control allocation algorithm for the digital micro-thruster array,namely status graph based control allocation(SGBCA)algorithm,which aims at finding the optimal micro thrusters combination scheme to realize the sequential control synthesis for micro/nano-satellite during real-time orbit control tasks.A mathematical model is set up for the control allocation of this multivariate over-actuated system.Through dividing thrusters into disjoint segments by offline calculation and combining segments dynamically online to provide a sequence of the required impulse for the micro/nano-satellite,the time complexity of the control allocation algorithm decreases significantly.All levels of impulse can be generated by the digital micro thruster arrays and the service life of the arrays can be extended using the segment converting strategy proposed in this paper.The simulation indicates that the algorithm can satisfy the requirements of real-time orbit control for micro/nano-satellites.
文摘The three-dimensional numerical simulation of two-phase plume flow of solid propellant micro-thrusters was developed.Then it was used to investigate the plume interference effect by combining the direct simulation Monte Carlo(DSMC) method for multi-component gas flow with the two-way coupling model for two-phase rarefied flow.At different space between the two micro-thrusters and different wall temperature,the plume interference effect was analyzed specifically.The results show that under the plume interference effect the gas is compressed and the flow direction is changed,which resulted in the increasing of gas pressure and temperature;solid phase made no significant effect on the flow parameters of gas phase;with the rising of the space between the two micro-thrusters,the maximum pressure decreased and the maximum temperature increased in the domain under the plume interference effect;the wall temperature could influence the temperature of the gas which is extremely close to the wall,but not the gas pressure.
基金the National Key Research and Development Program of China(Nos.2021YFC2202800 and 2021YFC2202604)the Opening Project of the Key Laboratory of Microgravity of the Chinese Academy of Sciences(No.NML202401)+1 种基金the Youth Innovation Promotion Association of the Chinese Academy of Sciences(No.2023022)the Strategic Priority Research Program of the Chinese Academy of Sciences(No.XDA1502110105)。
文摘The Taiji program is focused on achieving space-based gravitational wave detection in the frequency range of 0.1 mHz-1 Hz.To achieve drag-free control,Taiji satellites must be equipped with micro-thrusters that satisfy stringent requirements,including a continuously adjustable thrust,thrust resolution of 0.1μN,thrust noise of 0.1μN·Hz^(-0.5)and response time of less than 100 ms.This paper presents the progress of a variable cold gas micro-thruster being developed for the Taiji program.A series of technologies such as a bidirectional piezoelectric drive,spherical valve core,conical nozzle seal,miniature gas chamber,high thruster integration,combined digital and analogue communication and high-frequency closed-loop thrust control were employed to achieve engineering prototypes of a low-noise and fast-response micro-thruster.Ground performance tests indicate that the micro-thruster achieved a minimum thrust of close to 1.1 nN,a thrust resolution of 0.05μN and a maximum specific impulse of 69.1 s using nitrogen gas as the working fluid.The thrust noise was less than 0.1μN·Hz^(-0.5)in the frequency band of 10 mHz-1 Hz,and the thrust response time was 140 ms.The control parameters were further optimised to achieve a flow response time of 50 ms.The results indicate that the developed micro-thruster essentially met the performance requirements for drag-free control to facilitate space-based gravitational wave detection.
基金supported by the Shanghai Engineering Research Center of Space Engine(No.17DZ2280800)。
文摘In this work,a force measurement system is proposed to measure the thrust of plasma microthruster with thrust magnitude ranging from sub-micro-Newtons to hundreds micro-Newtons.The thrust measurement system uses an elastic torsional pendulum structure with a capacitance sensor to measure the displacement,which can reflect the position change caused by the applied force perpendicular to the pendulum axis.In the open-loop mode,the steady-state thrust or the impulse of the plasma micro-thruster can be obtained from the swing of the pendulum,and in the closed-loop mode the steady-state thrust can be obtained from the feedback force that keeps the pendulum at a specific position.The thrust respond of the system was calibrated using an electrostatic weak force generation device.Experimental results show that the system can measure a thrust range from 0 to 200μN in both open-loop mode and closed-loop mode with a thrust resolution of 0.1μN,and the system can response to a pulse bit at the magnitude of 0.1 m N s generated by a micro cathode arc thruster.The background noise of the closed-loop mode is lower than that of the open-loop mode,both less than 0.1 m N/Hz in the range of 10 mHz to 5 Hz.
基金National Natural Science Foundation of China (No.91216201, No.51205403)
文摘The applications of the micro-thrust e r and the challenges of micro-thrust measuring are introduced.The developments in measuring techniques for the micro-thrust are reviewed.Micro-thrust measu rements have previously been made either directly by mounting thrusters to the m easurement system or indirectly by mounting a target in the direct path of the e jected propellant.Several typical direct and indirect thrust-stands are presen ted and discussed in detail to illustrate the principles.Typical calibration me thods are also expounded.Finally,the resolution,uncertainty and thrust range of each thrust-stand are given,which may be helpful for the future thrust stan d design and micro-thrusters research.
文摘Pulsed plasma thrusters （PPT） are micro-propulsion devices used in satellites for station keeping. Conventionally the plasma discharge in a PPT is initiated by a spark plug. The primary objective of the present work was to develop and characterize a PPT that does not need a spark plug to initiate the plasma discharge. If the spark plug is eliminated, the size of the thrusters can be reduced and arrays of such thrusters can be manufactured using micro electro mechanical systems （MEMS） techniques, which can provide tremendous control authority over the satellite positioning. A parallel rail thruster was built and its performances were characterized inside a vacuum chamber to elucidate the effect of vacuum level on the performance. The electrical performance of the thruster was quantified by measuring the voltage output from a Rogowski coil, and the thrust produced by the developed thruster was estimated by measuring the force exerted by the plume on a light weight pendulum, whose deflection was measured using a laser displacement sensor. It was observed that the thruster can operate without a spark plug. In general, the performance parameters such as thrust, mass ablation, impulse bit, and specific impulse per discharge, would increase with the increase in pressure up to an optimum level due to the increase in discharge energy as well as the decrease in the total impedance of the plasma discharge. The thrust efficiency is found to be affected by the discharge energy.
基金support provided by the Foundation for Innovative Research Groups of the National Natural Science Foundation of China(51121004,41274041,51207033,and 11275055)
文摘In this paper, the technological progress on Chinese gravity exploration satellites is presented. Novel features such as ultra-stable structure, high accurate thermal control, drag-free and attitude control, micro-thrusters, aerodynamic configuration, the ability to perform micro-vibration analyses, microwave ranging system and mass center trimmer are described.
基金Project supported by the National Natural Science Foundation of China(No.61503334)the National Science Fund for Distinguished Young Scholars(No.61525403)。
文摘An ammonia self-managed vaporization propulsion (ASVP) system for micro-nano satellites is presented. Compared with a normal cold gas or liquefied gas propulsion system, a multiplex parallel sieve type vaporizer and related vaporization control methods are put forward to achieve self-managed vaporization of liquefied propellant. The problems of high vaporization latent heat and incomplete vaporization of liquefied ammonia are solved, so that the ASVP system takes great advantage of high theoretical specific impulse and high propellant storage density. Furthermore, the ASVP operation procedure and its physical chemistry theories and mathematical models are thoroughly analyzed. An optimal strategy of thrust control is proposed with consideration of thrust performance and energy efficiency. The ground tests indicate that the ASVP system weighs 1.8 kg (with 0.34-kg liquefied ammonia propellant) and reaches a specific impulse of more than 100 s, while the power consumption is less than 10 W. The ASVP system meets multiple requirements including high specific impulse, low power consumption, easy fabrication, and uniform adjustable thrust output, and thus is suitable for micro-nano satellites.
