CN108387260B - Electric propulsion plume test evaluation method - Google Patents
Electric propulsion plume test evaluation method Download PDFInfo
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- CN108387260B CN108387260B CN201711384983.3A CN201711384983A CN108387260B CN 108387260 B CN108387260 B CN 108387260B CN 201711384983 A CN201711384983 A CN 201711384983A CN 108387260 B CN108387260 B CN 108387260B
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Abstract
The invention discloses an electric propulsion plume test evaluation method, which comprises the following steps: measuring the ion beam current density, the electron concentration and the space plasma potential of the main beam plume region and the return plume region to obtain the evaluation results of the force effect and the thermal radiation effect of the plumes on the satellite solar wing; measuring the charging voltage and leakage current of different materials on the surface of the solar cell array in a plume plasma environment and the surface potential of the solar cell array in the plume plasma environment to obtain the evaluation result of the charging and discharging effect of the plumes on the surfaces of the satellite solar wing and the satellite body; and measuring the plasma plume pollution amount of the test sample of the satellite solar wing and the satellite body surface sensitive equipment to obtain the deposition pollution effect evaluation result of the plume on the satellite solar wing and the satellite body surface sensitive equipment. The invention realizes the test analysis and research of pollution of the plume on the satellite surface, ion sputtering corrosion and the influence of the plume on the satellite surface potential.
Description
Technical Field
The invention belongs to the technical field of electric propulsion systems, and particularly relates to an electric propulsion plume test evaluation method.
Background
The electric propulsion plume is a feather-shaped flow field formed by the expansion of jet flow to the external vacuum environment when the electric thruster works. The ion thruster plume is composed of ions, electrons, non-ionized neutral particles, charge exchange ions, sputtered pollutants and the like generated by the ionization of the propellant, moves and expands in a vacuum environment after leaving the thruster, has high movement speed and is strongly coupled with an electromagnetic field. The plasma generated by the electric thruster can be divided into two parts, one part is beam plasma directly ejected from the thruster, the beam ion energy is high, and the motion trail is almost a straight line due to small influence of an electromagnetic field. The beam ions are mainly distributed in a conical surface with a half angle of 15-60 degrees. The other part is the exchange charge plasma, which consists of exchange charge ions and electrons and returns to the satellite surface. The energy of the exchanged charge ions is low (generally lower than 20eV), and the ions move reversely under the action of an electric field around the satellite to form reflux ions, so that the surface of the satellite is polluted.
The beam high-energy ions collide with the surface of the solar wing of the spacecraft to generate interference force or torque, so that the attitude motion and control of the spacecraft are influenced, and the spacecraft is out of control and cannot work normally under severe conditions. In addition, the beam plasma expands and expands asymmetrically along the axial direction, and is unstable, and the interaction of the beam plasma and the geomagnetic field can cause the change of the magnitude and the direction of the thrust, so that interference force or moment is generated. The plume thermal effect of the ion thruster mainly refers to heat exchange generated by collision of beam high-energy ions and the surface of a spacecraft, so that the relevant characteristics of the surface of the spacecraft are changed.
The high-energy beam ions and the charge exchange ions collide with the surface of the spacecraft, and as long as the ion energy is greater than the sputtering threshold of the collided material, a sputtering corrosion effect is generated. For some electric propulsion satellites executing the north-south position protection tasks, beam high-energy ions may directly sweep through a solar cell array to generate severe sputtering corrosion, so that the performance of the cell array is reduced or the cell array is invalid. The sputter erosion of the spacecraft surface by the charge-exchanging ions is relatively weak, but the time-accumulating effect cannot be neglected if coupled with the local strong electric field distribution. The deposition contamination generated by the ion thruster comes from two aspects: firstly, materials such as a discharge chamber, a grid electrode and the like of the thruster are subjected to ion sputtering corrosion, and sputtering stripping materials flowing out of the thruster are deposited on the surface of the spacecraft in the form of ions, atoms and the like to form pollution; the other is one or more times of pollution generated by strippings on the surface of the spacecraft, which are corroded by ion sputtering. The influence of the surface pollution of the spacecraft is related to the range including a solar cell array, a thermal control surface, an optical sensor and the like, and the surface pollution also aggravates the unequal amount of electrification on the surface of the spacecraft. Although the deposition pollution rate generated by the ion thruster is very low, the long-time accumulation effect can also greatly influence the photo-thermal property of the thermal control material and the solar optical property of the solar cell glass cover plate due to the long working time. In addition, the interference influence of the plume during the electric propulsion operation on the field of view of the star sensor mainly causes the reduction of the test precision of the sensor, the increase of the signal-to-noise ratio and the reduction of the star number identification capability.
It can be seen that the plume problem for electric propulsion applications can be attributed to several areas:
(1) the interaction of the plume and the solar wing and the like has mechanical and thermal influence on the spacecraft;
(2) the charge deposition of the plume on the surface of the spacecraft can change the electrostatic environment of the spacecraft, and charging and discharging effects can be generated;
(3) the optical interference effect of the plume on the satellite sensitive parts;
(4) sputtering corrosion of the plume on the structural part of the thruster and other parts of the spacecraft;
(5) the pollution to the sensitive surface of the spacecraft after the particles generated by plume sputtering are redeposited;
(6) the current fluctuation of the thruster can cause various high-frequency plasma oscillations, which cause interference to the radio communication of the spacecraft.
Disclosure of Invention
The technical problem of the invention is solved: the defects of the prior art are overcome, and an electric propulsion plume test evaluation method is provided, so that the pollution of the plume on the satellite surface, the ion sputtering corrosion and the influence of the plume on the satellite surface potential are tested, analyzed and researched, and the rationality of the electric propulsion applied to the satellite platform is evaluated.
In order to solve the technical problem, the invention discloses an electric propulsion plume test evaluation method, which comprises the following steps:
step 1, measuring ion beam current density, electron concentration and space plasma potential of a main beam plume region and a return plume region to obtain force effect and thermal radiation effect evaluation results of the plumes on a satellite solar wing;
step 2, measuring the charging voltage and the leakage current of different materials on the surface of the solar cell array in the plume plasma environment and the surface potential of the solar cell array in the plume plasma environment to obtain the charge-discharge effect evaluation result of the plumes on the surfaces of the satellite solar wing and the satellite body;
and 3, measuring plasma plume pollution amount of the test sample of the satellite solar wing and the satellite body surface sensitive equipment to obtain a deposition pollution effect evaluation result of the plume on the satellite solar wing and the satellite body surface sensitive equipment.
In the above method for evaluating an electrically propelled plume test, the method further comprises:
measuring the ion beam current density of the main beam current plume region and the return current plume region through a Faraday diagnosis probe;
the space plasma potentials of the main beam plume region and the return plume region are measured by a Langmuir diagnostic probe.
In the above method for evaluating the electric propulsion plume test, the ion beam density, the electron concentration and the space plasma potential in the main beam plume region and the return plume region are measured to obtain the evaluation results of the force effect and the thermal radiation effect of the plume on the satellite solar wing, including:
calculating to obtain the electron temperature T according to the formula (1.1)e：
Wherein e represents the charge amount of electrons, V represents the probe potential, and VspRepresenting the space plasma potential, I represents the probe current, Ie0Representing the collected current corresponding to the space potential;
according to the formula (1.2), the electron concentration n is calculatede：
Wherein A represents the probe surface area and k represents the Boltzmann constant;
according to the formula (1.3), the particle velocity v is calculatedi：
Wherein j represents the ion beam current density, niRepresenting the particle number density;
according to the formula (1.4), the energy flux per unit time of the particles on the boundary area S is calculated
Wherein V represents the volume of the target surface,iwhich represents the energy of a single particle,
denotes the normal velocity, m, perpendicular to the boundary surface SiRepresenting the quality of the ith particle;
according to the formula (1.5), the energy flux density of single particle is calculated
The total energy flux density of all particles was calculated according to equation (1.6):
where N represents the actual total number of particles and N represents the number of mesh particles.
The force effect F is calculated according to equation (1.7):
wherein m represents the mass of the particles.
In the above method for evaluating an electric propulsion plume test, the charging voltage and the leakage current of different materials on the surface of a solar cell array in a plume plasma environment and the surface potential of the solar cell array in the plume plasma environment are measured to obtain the evaluation result of the charging and discharging effect of the plume on the surface of the satellite solar wing and the satellite body, and the evaluation result includes:
mounting a simulated solar cell test piece in a set plume test area;
testing the electrical property of the simulated solar cell test piece before testing, and recording the test result;
connecting a non-contact surface potentiometer probe with a surface material of a simulated solar cell test piece, and connecting the non-contact surface potentiometer probe with acquisition equipment by penetrating a cable through a cabin to the outside of a vacuum cabin; the non-contact surface potentiometer probe is used for testing the surface potential of the simulated solar cell test piece;
the vacuum degree is better than 5.0 × 10 by vacuum pumping-3Pa, electrically propelling to work and forming a stable plume plasma environment;
setting the analog power supply in a constant voltage state, and setting the limiting current to be 4.0A; opening a non-contact surface potentiometer, monitoring the surface potential of a test sample, and recording data and a curve in real time;
adjusting the inter-string working voltage of the analog power supply, monitoring the current in a loop through an ammeter, and recording data and curves in real time; the initial adjustment voltage of the inter-string working voltage is 50V, the maximum adjustment voltage is 200V, and the adjustment strategy is that the inter-string working voltage is increased by 10V every 10 min;
after repeating the measurement for two times, turning off the non-contact surface potentiometer, turning off the analog power supply, and stopping the electric propulsion;
judging and reading the monitoring results of the output current and the potentiometer voltage of the simulated solar cell test piece; if the measurement result does not exceed the current fluctuation of 1% of the constant state value, determining that the current leakage phenomenon does not occur;
and after the test is finished, electrical performance retesting is carried out on the simulated solar cell test piece, the electrical performance retest is compared with the electrical performance test result before the test, and if the output current performance is reduced by more than 1%, the influence of the plasma plume on the performance of the solar cell test piece is determined.
In the above method for evaluating an electrically propelled plume test, the method further comprises:
arranging a test sample and a quartz crystal microbalance at the rear end of the thruster; wherein, the test sample comprises: a solar cell glass cover plate test piece, an optical solar reflector test piece and a solar wing silver interconnection piece;
arranging a test sample and a quartz crystal microbalance on an upper side plate and a lower side plate of the thruster, wherein a normal line of a collecting surface of the quartz crystal microbalance is kept to be superposed with a cross point of an axial direction and an outlet surface of the thruster, arrangement angles of measuring points are sequentially arranged outside a main flow area of the thruster in a dense-first and sparse-later mode, the distance from the measuring points to the outlet surface of the thruster is not less than 1.2m, the axial direction of the thruster is zero, and a measuring angle is not less than 110 degrees;
the far end of the thruster is provided with an anti-sputtering molecular sink, wherein the anti-sputtering molecular sink is used for absorbing high-energy particles and preventing the high-speed particles from scouring the bottom of the vacuum chamber and sputtering particles to pollute a test object;
the outer surface of the quartz crystal microbalance is covered with a protective sleeve to prevent the vacuum chamber body from being polluted to influence the test error.
In the above method for evaluating the electric propulsion plume test, the plasma plume contamination amount of the test sample of the satellite solar wing and the satellite body surface sensitive device is measured to obtain the result of evaluating the deposition contamination effect of the flow on the satellite solar wing and the satellite body surface sensitive device, which includes:
according to the formula (2.1), calculating the change delta f of the resonant frequency of the quartz crystal microbalance:
wherein f is0The fundamental frequency of the quartz crystal microbalance is shown,
represents the piezoelectrically-enhanced shear modulus, ρ, of quartzqRepresents the quartz density, A represents the probe electrode area, Δ m represents the quality of the contaminant mass etch, SfThe representation represents the sensitivity of the sensor;
and determining the pollution mass of the substances on the quartz crystal microbalance and the mass lost by the energy ion etching surface according to the change of the resonance frequency of the quartz crystal microbalance.
The invention has the following advantages:
the invention discloses an electric propulsion plume test evaluation method, which can realize reasonable verification of the mechanical and thermal influence on a spacecraft caused by the interaction of a plume and a solar wing, the influence on the electrostatic environment of the spacecraft caused by the charge deposition of the plume on the surface of the spacecraft, the optical interference influence on a satellite sensitive part caused by the plume, the sputtering corrosion influence on a thruster structural member and other parts of the spacecraft caused by the plume, the pollution influence on the sensitive surface of the spacecraft caused by the redeposition of particles generated by plume sputtering, and the interference influence on the radio communication of the spacecraft caused by various high-frequency plasma oscillations caused by the current fluctuation of the thruster, and can evaluate the rationality of the electric propulsion applied to a satellite platform.
Drawings
FIG. 1 is a flow chart illustrating the steps of a method for evaluating an electrically propelled plume test in accordance with an embodiment of the present invention;
FIG. 2 is a schematic diagram of an overall plan of an evaluation method of an electrically-propelled plume test according to an embodiment of the present invention;
FIG. 3 is a schematic diagram of an electric propulsion plume plasma diagnosis test based on a three-dimensional moving mechanism in an embodiment of the present invention;
FIG. 4 is a schematic diagram of a testing method of a Faraday probe testing system in an embodiment of the invention;
figure 5 is a schematic diagram illustrating the measurement principle of a langmuir diagnostic probe according to an embodiment of the present invention;
figure 6 is an I-V plot of langmuir diagnostic probe measurements taken in accordance with an embodiment of the present invention;
FIG. 7 is a schematic diagram of a charge/discharge test circuit for a solar cell sample according to an embodiment of the present invention;
FIG. 8 is a schematic view of a quartz crystal microbalance test electrically propelled plume sputtering and contamination in accordance with an embodiment of the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, common embodiments of the present invention will be described in further detail below with reference to the accompanying drawings.
Because the electric propulsion technology is not used in space in China, the research on the plume in China is still deficient and is in the starting stage. At present, some researches on plume numerical simulation and tests developed domestically are only the analysis research on near-field plumes, and the analysis result mainly plays a guiding role in the design and optimization of the thruster, particularly the research on the corrosion mechanism and the service life of an ion extraction system. In order to realize the application of electric propulsion on a spacecraft, the characteristics of a far field (2.6-6 m away from the exit of the thruster) and the characteristics of a backflow area of a thruster plume, the characteristics of mutual influence between the thruster and a star body and the like need to be tested and researched. Although the results of foreign research can provide a certain reference value, the research on the aspect has the particularity for the domestic electric propulsion satellite application and must be carried out by oneself. Corresponding test design and verification are not carried out in China before the date of research on the problems of far-field plume characteristics, the influence of the plume on the spacecraft, how to protect the spacecraft and the like.
At present, a plurality of organizations are developing numerical simulation research of far-field plumes, the numerical simulation analysis needs support and verification of experimental data, and the experimental research of a plume influence characteristic system is still a blank at home. The invention discloses an electric propulsion plume test evaluation method, which can realize the measurement of the divergence angle of an ion beam in a main beam plume region, the measurement of the density of the ion beam, the measurement of the plasma potential, the measurement of the density of the ion beam in a return plume region, the measurement of the plasma potential and the like, and the measurement result can be used as the evaluation basis of the comprehensive influence effect of the electric propulsion plume. Therefore, the plume influence characteristic tests in the invention fill the domestic blank and lay a solid foundation for the application of the electric propulsion technology.
Referring to FIG. 1, a flow chart of steps of a method for evaluating an electrically propelled plume test in an embodiment of the present invention is shown. Referring to fig. 2, a schematic diagram of an overall planning of an electric propulsion plume test evaluation method in an embodiment of the present invention is shown. In this embodiment, the method for evaluating an electrically propelled plume test includes:
In the present embodiment, the ion beam current densities of the main beam plume region and the return plume region can be measured by a Faraday diagnostic probe (Faraday), and the space plasma potentials of the main beam plume region and the return plume region can be measured by a langmuir diagnostic probe (L angmuir).
In this embodiment, after the ion beam current density of the plume plasma and the spatial plasma potential distribution rule are obtained, the calculation of the particle concentration and the particle parameter statistics of the target object can be performed, the energy and momentum loss upper limit of the plasma colliding with the surface of the target object under the influence of the spatial potential are counted, and the force effect on the target object is estimated. Preferably, the specific implementation flow of step 101 may be as follows:
substep 1011, calculating the electron temperature T according to the formula (1.1)eThe unit is electron voltage:
wherein e represents the charge amount of electrons (unit: coulomb), V represents the probe potential (unit: volt), and V representsspRepresents the space plasma potential (unit: volt), I represents the probe current (unit: milliamp), Ie0And represents the acquisition current (unit: milliampere) corresponding to the space potential.
Substep 1012, calculating the electron concentration n according to the formula (1.2)e(unit: number per cubic centimeter):
wherein A represents the probe surface area (unit: square centimeter) and k represents the Boltzmann constant.
Substep 1013, calculating the particle velocity v according to the formula (1.3)i：
Wherein j represents the ion beam current density, niRepresenting the particle number density.
Substep 1014, calculating the energy flux per unit time of the particle on the boundary area S according to the formula (1.4)
In this embodiment, assume that in a grid with a volume V and a boundary area S close to the target surface, the energy of a single particle isiThe particle velocity is viNormal velocity perpendicular to the boundary surface S is
The energy flux per unit time brought by this particle on the boundary surface S
Comprises the following steps:
wherein V represents the volume of the target surface,iwhich represents the energy of a single particle,
denotes the normal velocity, m, perpendicular to the boundary surface SiIndicating the quality of the ith particle.
Substep 1015, calculating the energy flux density of single particle according to the formula (1.5)
Substep 1016, calculating the total energy flux density of all particles according to equation (1.6):
where N represents the actual total number of particles and N represents the number of mesh particles.
Substep 1017, calculating the force effect F according to equation (1.7).
In this embodiment, assuming that the particle velocities are relatively uniform in each grid, equation (1.5) may be equivalent to:
then, assuming that in the mesh with the volume V and the boundary area S close to the target surface, the force effect F is obtained by the kinetic energy theorem:
wherein m represents the mass of the particles.
In this embodiment, the specific implementation flow of step 102 may be as follows:
and a substep 1021, installing the simulated solar cell test piece in the set plume test area.
And a substep 1022 of performing an electrical property test on the simulated solar cell test piece before the test and recording a test result.
1023, connecting a non-contact surface potentiometer probe with the surface material of the simulated solar cell test piece, penetrating the test piece to the outside of the vacuum chamber through a cable, and connecting the test piece with a collection device; the non-contact surface potentiometer probe is used for testing the surface potential of the simulated solar cell test piece.
Substep 1024 of evacuating to a vacuum level better than 5.0 × 10-3Pa, electric propulsion works and forms a stable plume plasma environment.
And a substep 1025 of setting the analog power supply to a constant voltage state and setting the limiting current to 4.0A. And opening the non-contact surface potentiometer, monitoring the surface potential of the test sample, and recording data and curves in real time.
A substep 1026, adjusting the inter-string working voltage of the analog power supply, monitoring the current in the loop through an ammeter, and recording data and curves in real time; the initial adjustment voltage of the inter-string working voltage is 50V, the maximum adjustment voltage is 200V, and the adjustment strategy is that the inter-string working voltage is increased by 10V every 10 min.
And a substep 1027 of repeating the measurement twice, turning off the non-contact surface potentiometer, turning off the analog power supply, and stopping the electric propulsion.
Substep 1028, interpreting the monitoring results of the output current and the potentiometer voltage of the simulated solar cell test piece; and if the measurement result does not exceed the current fluctuation of 1% of the constant state value, determining that the current leakage phenomenon does not occur.
And a substep 1029, after the test is finished, electrical performance retesting is carried out on the simulated solar cell test piece, the retest is compared with the electrical performance test result before the test, and if the reduction amplitude of the output current performance exceeds 1%, the influence of the plasma plume on the performance of the solar cell test piece is determined.
In a preferred embodiment of the present invention, referring to fig. 3, a schematic diagram of an electric propulsion plume plasma diagnostic test based on a three-dimensional moving mechanism in an embodiment of the present invention is shown. In this embodiment, the test system mainly includes a plurality of diagnostic probe assemblies, a three-dimensional moving mechanism, a test power supply device, a processing device, a display device, a data storage device, and the like. The diagnostic probe assembly arrangement area is arranged according to the beam flow area and the reflux area, and the multi-point position coverage is realized by installing and fixing the probe diagnostic assembly on a three-dimensional moving mechanismOn the movable rod, the probe is driven by the three-dimensional moving mechanism to move so as to realize the measurement of different positions. The power supply equipment and the acquisition and processing equipment of the probe are both arranged outside the vacuum chamber and are connected with the probe through a chamber penetrating flange, the probe, the three-dimensional moving mechanism and the cable in the vacuum chamber are all required to be insulated from the vacuum chamber, and protection measures are provided to avoid the influence of a plume plasma environment on the probe. The system realizes the test of the parameter characteristics of beam plasma in the axial range of the thruster and the test of the parameter characteristics of plasma of plume backflow in the radial range of the thruster. Wherein, the envelope coverage of the beam density test is 0.1mA/cm2～50mA/cm2Space plume plasma space potential envelope is in the range of-50V to 50V, plasma electron temperature test envelope is 0.01eV to 50eV, and plasma concentration is 10 eV in beam area13Per m3～1017Per m3The magnitude is that the reflux area covers 108/m3～1013The plasma density is in the order of 3 m, and the ion energy test range is 0 eV-80 eV.
The method for measuring the ion beam current density by using the Faraday probe comprises the following steps:
the probe is made of a plane disc collector with a protective ring sleeved outside, and a test principle diagram refers to fig. 4, which shows a schematic diagram of a test method of a test system of the faraday probe in the embodiment of the invention. When measuring ion current, the Faraday probe is parallel to the axial direction (x direction) of the thruster, the collecting surface of the Faraday probe is perpendicular to the extracted beam, is connected with the collector signal collecting device, and applies negative bias voltage to the collector to enable the collector to collect ions. Meanwhile, negative bias voltage is applied to the protective ring on the outer layer of the collector, the protective ring plays a role in shielding non-axial ions, and the probability of collecting the non-axial ions by the collector is reduced.
The testing method is that the collector bias voltage (generally-12V to-30V) is set, the probe collects ions through repelling electrons, the current value of the ions is acquired, and the ion current density can be obtained through calculation. The ion current density is calculated according to the following formula:
in the formula: j represents receivingThe integrated ion beam current density (unit: milliampere/square centimeter), V represents the measured voltage (unit: volt), R represents the shunt resistance value (unit: ohm), ApRepresenting the area of the collection disk (in square centimeters).
Referring to fig. 5, a schematic diagram of the measurement principle of a langmuir diagnostic probe according to an embodiment of the present invention is shown. Referring to figure 6, an I-V plot of measurements from a langmuir diagnostic probe in an embodiment of the present invention is shown. In this example, a Langmuir diagnostic probe was placed in plasma, and an external circuit was added to the plasma to constitute a Langmuir diagnostic probe measurement circuit. In this example, the method of measuring the spatial plasma potentials in the plume flow region and the return flow region using the langmuir diagnostic probe is as follows:
the scanning power supply was adjusted to cause the potential of the probe to change from negative to positive and the current value in the measurement loop for each voltage was plotted to form a typical voltammogram for an L angmuir probe in FIG. 5, V represents the probe potential and V represents the probe potentialfIndicating the floating potential, V, of the plasmaspRepresenting the space plasma potential. When V is equal to VfWhen the plasma is detected, the electrons collected by the probe are equal to the ion current and opposite to the ion current, the total current of the probe is zero, and the voltage value of the probe current on the I-V curve is zero, namely the suspension potential V of the plasmaf。
Preferably, the method for acquiring the plasma parameters from the voltammetry curve of the L angmuir probe is as follows:
an inflection point appears on the I-V curve of the probe current at the joint of the electron saturation flow region and the transition region, and the voltage value corresponding to the inflection point is the space plasma potential Vsp. The inflection point is not obvious, particularly, the position of the inflection point is not easy to obtain by direct observation of a planar probe, a first derivative is required to be obtained for an I-V curve, and the point corresponding to the maximum value of the derivative is the inflection point. I.e. when the slope of the curve
The current at which the maximum is reached is I ═ Ieo-Iio. Therefore, the slope on the test curve
The voltage corresponding to the maximum value is space plasma potential Vsp。
The quartz crystal microbalance testing technology can be used for monitoring the outgassing of non-metallic materials and the deposition of molecular contamination on the surface of a spacecraft sensitive system. The sensor selects piezoelectric quartz crystal as a detecting element of micro-molecular pollutants.
The invention achieves the purpose of measuring and analyzing the current leakage event by monitoring the output current of the solar cell small plate at the characteristic position in real time. Referring to fig. 7, a schematic connection diagram of a charge and discharge test circuit of a solar cell sample in an embodiment of the present invention is shown. As shown in fig. 7, R is the loop load resistance; the meter A is an ammeter and monitors the current in a loop; c is the solar cell array to substrate analog capacitance; r is a loop load resistor; and P is an analog power supply and provides working voltage for the solar cell strings. The test isolates the high voltage solar cell array sample structure from the vacuum chamber. Charging the surface of the sample in a plume environment, and isolating the potential array sample structure of different materials on the surface of the high-voltage solar cell array sample from the vacuum chamber by using a non-contact potentiometer. And charging the surface of the sample in a plume environment, and measuring the potential of different materials on the surface of the high-voltage solar cell array sample by using a non-contact potentiometer.
103, measuring plasma plume pollution amount of the test sample of the satellite solar wing and the satellite body surface sensitive equipment to obtain a deposition pollution effect evaluation result of the plume on the satellite solar wing and the satellite body surface sensitive equipment.
Referring to fig. 8, a schematic diagram of a quartz crystal microbalance test electrically-propelled plume sputtering and contamination prevention in the embodiment of the invention is shown, in the embodiment, a backflow test area is shown in fig. 8, a distance L relative to a nozzle surface of a thruster is determined according to test chamber conditions and generally should not be less than 1m, a test sample and a Quartz Crystal Microbalance (QCM) can be arranged at the rear end of the thruster, wherein the test sample comprises a solar cell glass cover test piece, an optical solar reflector test piece and a solar wing silver interconnection piece, then the test sample and the quartz crystal microbalance are arranged on an upper side plate and a lower side plate of the thruster, a normal line of a collection surface of the quartz crystal microbalance is kept to be coincident with an intersection point of an axial direction and an outlet surface of the thruster, a measuring point arrangement angle is sequentially arranged outside a main beam area of the thruster in a dense-sparse manner, a distance R from the measuring point to the outlet surface of the thruster is not less than 1.2m, the axial direction of the thruster is zero degree, a measuring angle is not less than 110, a sputtering molecule sink is arranged at the far end of the thruster, a vacuum protection sleeve is used for absorbing high-energy particles, preventing the vacuum particles from being flushed out and preventing the pollution of a test object from being discharged from the pollution of the vacuum capsule.
Preferably, the specific implementation flow of step 103 may be as follows:
and a substep 1031, calculating to obtain the resonance frequency change of the quartz crystal microbalance according to the formula (2.1).
In this example, the change in resonant frequency of the quartz crystal microbalance is proportional to the applied mass, and the change in resonant frequency Δ f (in Hz) of the quartz crystal microbalance can be calculated according to the following equation (2.1):
wherein f is0Representing the fundamental frequency (in hz) of a quartz crystal microbalance,
showing the piezoelectric enhanced shear modulus (unit: Pa), ρ of quartzqThe quartz density (unit: grams per cubic centimeter), A the probe electrode area (unit: square centimeter), Δ m the mass of the contaminant mass etch (unit: nanograms per square centimeter), SfThe representation represents the sensitivity of the sensor (in volts per gram).
Sub-step 1032 determines the contaminating mass of the substance on the quartz crystal microbalance and the mass lost to the energetic ion etching surface based on the change in the resonant frequency of the quartz crystal microbalance.
In summary, the present invention provides an electric propulsion plume test evaluation method, which includes a characteristic parameter measurement test for a thruster-derived beam, a characteristic parameter measurement test for a backflow plasma, a plume charge-discharge effect verification test, a plume pollution effect verification test, a plume light interference verification test, and the like: testing the ion beam density of the main beam area by using a Faraday diagnosis probe; measuring the space plasma potential in a beam flow area of the electric thruster by using a Langmuir diagnostic probe so as to calculate the concentration of particles, and evaluating the basic distribution rule and influence rule of plasmas in a main beam flow area and a return flow area of the electric thruster and the influence on the force effect of the satellites and the solar wings; the method comprises the following steps of performing test testing on a charge-discharge phenomenon of a simulated solar cell test piece under the action of plume, and evaluating the charge-discharge effect of the plume on the test piece; experimental testing of the contamination effect of the plume on the optically sensitive sample and QCM was performed to assess the plume contamination effect. The test results can provide data support for engineering to master an electric propulsion plume physical mechanism, provide data support for optimization layout of the spacecraft, provide data support for analysis of interference torque on the outer surface of the spacecraft, and provide data support for self-compatibility of the spacecraft and electric propulsion.
The embodiments in the present description are all described in a progressive manner, each embodiment focuses on differences from other embodiments, and the same and similar parts among the embodiments are referred to each other.
The above description is only for the best mode of the present invention, but the scope of the present invention is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present invention are included in the scope of the present invention.
Those skilled in the art will appreciate that the invention may be practiced without these specific details.
Claims (3)
1. An electrically propelled plume test evaluation method, comprising:
step 1, measuring ion beam current density, electron concentration and space plasma potential of a main beam plume region and a return plume region to obtain force effect and thermal radiation effect evaluation results of the plumes on a satellite solar wing; the method comprises the following steps:
calculating to obtain the electron temperature T according to the formula (1.1)e：
Wherein e represents the charge amount of electrons, V represents the probe potential, and VspRepresenting the space plasma potential, I represents the probe current, Ie0Representing the collected current corresponding to the space potential;
according to the formula (1.2), the electron concentration n is calculatede：
Wherein A represents the probe surface area and k represents the Boltzmann constant;
according to the formula (1.3), the particle velocity v is calculatedi：
Wherein j represents the ion beam current density, niRepresenting the particle number density;
according to the formula (1.4), the energy flux per unit time of the particles on the boundary area S is calculated
Wherein V represents the volume of the target surface,iwhich represents the energy of a single particle,
denotes the normal velocity, m, perpendicular to the boundary surface SiRepresenting the quality of the ith particle;
according to the formula (1.5), the energy flux density of single particle is calculated
The total energy flux density of all particles was calculated according to equation (1.6):
wherein N represents the total number of actual particles, and N represents the number of grid particles;
the force effect F is calculated according to equation (1.7):
wherein m represents the mass of the particles;
step 2, measuring the charging voltage and the leakage current of different materials on the surface of the solar cell array in the plume plasma environment and the surface potential of the solar cell array in the plume plasma environment to obtain the charge-discharge effect evaluation result of the plumes on the surfaces of the satellite solar wing and the satellite body; the method comprises the following steps:
mounting a simulated solar cell test piece in a set plume test area;
testing the electrical property of the simulated solar cell test piece before testing, and recording the test result;
connecting a non-contact surface potentiometer probe with a surface material of a simulated solar cell test piece, and connecting the non-contact surface potentiometer probe with acquisition equipment by penetrating a cable through a cabin to the outside of a vacuum cabin; the non-contact surface potentiometer probe is used for testing the surface potential of the simulated solar cell test piece;
the vacuum degree is better than 5.0 × 10 by vacuum pumping-3Pa, electrically propelling to work and forming a stable plume plasma environment;
setting the analog power supply in a constant voltage state, and setting the limiting current to be 4.0A; opening a non-contact surface potentiometer, monitoring the surface potential of a test sample, and recording data and a curve in real time;
adjusting the inter-string working voltage of the analog power supply, monitoring the current in a loop through an ammeter, and recording data and curves in real time; the initial adjustment voltage of the inter-string working voltage is 50V, the maximum adjustment voltage is 200V, and the adjustment strategy is that the inter-string working voltage is increased by 10V every 10 min;
after repeating the measurement for two times, turning off the non-contact surface potentiometer, turning off the analog power supply, and stopping the electric propulsion;
judging and reading the monitoring results of the output current and the potentiometer voltage of the simulated solar cell test piece; if the measurement result does not exceed the current fluctuation of 1% of the constant state value, determining that the current leakage phenomenon does not occur;
after the test is finished, electrical performance retesting is carried out on the simulated solar cell test piece, the retest is compared with the electrical performance test result before the test, and if the output current performance is reduced by more than 1%, the influence of the plasma plume on the performance of the solar cell test piece is determined;
step 3, measuring plasma plume pollution amount of the test sample of the satellite solar wing and the satellite body surface sensitive equipment to obtain a deposition pollution effect evaluation result of the plume on the satellite solar wing and the satellite body surface sensitive equipment; the method comprises the following steps:
according to the formula (2.1), calculating the change delta f of the resonant frequency of the quartz crystal microbalance:
wherein f is0The fundamental frequency of the quartz crystal microbalance is shown,
represents the piezoelectrically-enhanced shear modulus, ρ, of quartzqRepresents the quartz density, A represents the probe electrode area, Δ m represents the quality of the contaminant mass etch, SfThe representation represents the sensitivity of the sensor;
and determining the pollution mass of the substances on the quartz crystal microbalance and the mass lost by the energy ion etching surface according to the change of the resonance frequency of the quartz crystal microbalance.
2. The electrically propelled plume test evaluation method of claim 1, further comprising:
measuring the ion beam current density of the main beam current plume region and the return current plume region through a Faraday diagnosis probe;
the space plasma potentials of the main beam plume region and the return plume region are measured by a Langmuir diagnostic probe.
3. The electrically propelled plume test evaluation method of claim 1, further comprising:
arranging a test sample and a quartz crystal microbalance at the rear end of the thruster; wherein, the test sample comprises: a solar cell glass cover plate test piece, an optical solar reflector test piece and a solar wing silver interconnection piece;
arranging a test sample and a quartz crystal microbalance on an upper side plate and a lower side plate of the thruster, wherein a normal line of a collecting surface of the quartz crystal microbalance is kept to be superposed with a cross point of an axial direction and an outlet surface of the thruster, arrangement angles of measuring points are sequentially arranged outside a main flow area of the thruster in a dense-first and sparse-later mode, the distance from the measuring points to the outlet surface of the thruster is not less than 1.2m, the axial direction of the thruster is zero, and a measuring angle is not less than 110 degrees;
the far end of the thruster is provided with an anti-sputtering molecular sink, wherein the anti-sputtering molecular sink is used for absorbing high-energy particles and preventing the high-speed particles from scouring the bottom of the vacuum chamber and sputtering particles to pollute a test object;
the outer surface of the quartz crystal microbalance is covered with a protective sleeve to prevent the vacuum chamber body from being polluted to influence the test error.
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| CN109507489A (en) * | 2018-10-18 | 2019-03-22 | 北京理工大学 | A kind of probe system for low temperature plasma potential diagnostic |
| CN110132606B (en) * | 2019-05-28 | 2020-06-19 | 北京航空航天大学 | System, method and device for measuring plume parameters of thruster |
| CN110677967B (en) * | 2019-09-18 | 2021-09-07 | 北京空间飞行器总体设计部 | Satellite surface plasma charging current monitoring method based on Faraday cup |
| CN115696708A (en) * | 2020-03-02 | 2023-02-03 | 辽宁工业大学 | Double-probe detection device and detection method for discharge state in welding arc |
| CN112329201B (en) * | 2020-09-30 | 2022-11-04 | 兰州空间技术物理研究所 | Integrated ionic electric propulsion simulation model and method |
| CN113466920A (en) * | 2021-07-02 | 2021-10-01 | 兰州空间技术物理研究所 | Electric propulsion beam characteristic parameter diagnosis equipment |
| CN113553752B (en) * | 2021-07-23 | 2022-08-02 | 中国电子科技集团公司第十八研究所 | Research method for influence of electric propulsion plume on solar cell array |
| CN114173463B (en) * | 2021-11-24 | 2024-03-22 | 广州市微生物研究所集团股份有限公司 | Method and device for testing plasma density of air purifier |
| CN114113717B (en) * | 2021-11-24 | 2022-08-30 | 北京航空航天大学 | Plug-in type totally-enclosed Faraday probe |
| CN114245554A (en) * | 2021-12-03 | 2022-03-25 | 北京东方计量测试研究所 | Plasma thruster plume parameter multipoint measuring device and measuring method |
| CN114264446B (en) * | 2021-12-23 | 2023-08-29 | 中电科蓝天科技股份有限公司 | Multi-degree-of-freedom pose-adjustable electric propulsion plume test tool |
| CN117169606A (en) * | 2023-11-01 | 2023-12-05 | 国科大杭州高等研究院 | Multifunctional ion characteristic detection probe and ion characteristic detection method |
| CN117740385B (en) * | 2024-02-20 | 2024-05-28 | 国科大杭州高等研究院 | Micro thruster diagnosis measuring device and operation method thereof |
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| CN101876615B (en) * | 2009-12-17 | 2012-07-11 | 中国航天科技集团公司第五研究院第五一○研究所 | Method for monitoring plume contamination of attitude control engine of spacecraft |
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