CN111536007A - Method for reducing heat load of Hall thruster - Google Patents
Method for reducing heat load of Hall thruster Download PDFInfo
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- CN111536007A CN111536007A CN202010385673.9A CN202010385673A CN111536007A CN 111536007 A CN111536007 A CN 111536007A CN 202010385673 A CN202010385673 A CN 202010385673A CN 111536007 A CN111536007 A CN 111536007A
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- energy loss
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F03—MACHINES OR ENGINES FOR LIQUIDS; WIND, SPRING, OR WEIGHT MOTORS; PRODUCING MECHANICAL POWER OR A REACTIVE PROPULSIVE THRUST, NOT OTHERWISE PROVIDED FOR
- F03H—PRODUCING A REACTIVE PROPULSIVE THRUST, NOT OTHERWISE PROVIDED FOR
- F03H1/00—Using plasma to produce a reactive propulsive thrust
- F03H1/0006—Details applicable to different types of plasma thrusters
- F03H1/0031—Thermal management, heating or cooling parts of the thruster
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F03—MACHINES OR ENGINES FOR LIQUIDS; WIND, SPRING, OR WEIGHT MOTORS; PRODUCING MECHANICAL POWER OR A REACTIVE PROPULSIVE THRUST, NOT OTHERWISE PROVIDED FOR
- F03H—PRODUCING A REACTIVE PROPULSIVE THRUST, NOT OTHERWISE PROVIDED FOR
- F03H1/00—Using plasma to produce a reactive propulsive thrust
- F03H1/0037—Electrostatic ion thrusters
- F03H1/0062—Electrostatic ion thrusters grid-less with an applied magnetic field
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- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
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- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
- G06F30/23—Design optimisation, verification or simulation using finite element methods [FEM] or finite difference methods [FDM]
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- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
- G06F2119/08—Thermal analysis or thermal optimisation
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Abstract
The invention belongs to the technical field of Hall thrusters, and particularly relates to a method for reducing the thermal load of a Hall thruster. The method mainly comprises the following steps: establishing a thermal balance model; solving the energy loss of a discharge chamber of the Hall thruster; determining a boundary condition for thermal conduction; determining the inner diameter of the radiating fin; the performance of the magnetic field was evaluated based on the effect of temperature on the magnetic field. By adopting the magnetic shield thermal optimization method, the thermal load of the Hall thruster can be reduced, the temperature distribution and the heat flow trend of the magnetic shield can be optimized on the basis that the magnetic field intensity reaches a certain value, and a certain theoretical and practical basis is laid for the thermomagnetic correlation analysis of the Hall thruster in related research.
Description
Technical Field
The invention belongs to the technical field of Hall thrusters, and particularly relates to a method for reducing the thermal load of a Hall thruster.
Background
The application is a divisional application, and the application number of a parent application is as follows: CN 201810585807.4.
The Hall thruster is an advanced electric propulsion device, is widely applied to the field of satellite position keeping and attitude control, and becomes the preferred propulsion device of a future spacecraft by the advantages of simple structure, high specific impulse, high efficiency and the like. The hall thruster is accelerated by an electric field through a propellant in the thruster and confines electrons in a magnetic field, and the propellant is ionized by the electrons, so that the accelerated ions generate thrust, and the ions in the plume are neutralized.
However, for the high power hall thruster, if the magnetic screen has a high thermal load, the alignment of the magnetic domains will be adversely affected, even if the ferromagnetic material becomes a paramagnetic material, thereby greatly reducing the efficiency of the thruster. In the prior art, a related thermal optimization method or strategy is not provided, so that how to reduce the thermal load of the hall thruster and optimize the temperature distribution and the heat flow trend of the magnetic screen becomes a main problem to be solved by the invention.
Disclosure of Invention
In order to solve the defects in the prior art, the invention discloses a method for reducing the thermal load of a Hall thruster, which is realized by adopting the following technical scheme.
A method for reducing the thermal load of a Hall thruster is characterized by comprising the following steps: the method comprises the following steps:
1) establishing a thermal balance model; the energy loss of the Hall thruster mainly comprises energy loss of a discharge chamber, energy loss of a cathode and energy loss of an excitation coil, wherein the energy loss of the discharge chamber mainly comes from ion energy, anode energy, plasma wall surface energy deposition, gas ionization energy, radiation energy loss and the like;
2) according to the thermal balance model, solving the energy loss of the discharge chamber of the Hall thruster and the ratio of each energy loss to the total energy loss, and determining that the energy loss ratio of the wall surface of the discharge chamber is the highest;
3) establishing a discharge chamber wall surface heat conduction model by adopting a finite element analysis method, establishing a heat flow balance equation for each discrete node in the heat conduction model by combining a Fourier heat conduction law and an energy conservation law, simultaneously determining a boundary condition of the heat conduction, and verifying an error range of the heat conduction obtained based on the boundary condition through experiments; if the error range is within the allowed error range threshold value interval, judging that the simulation model has the capability of describing the temperature distribution of the thruster, and if not, re-determining the boundary condition until the error range is within the allowed error range threshold value interval;
4) analyzing the influence rule of the inner diameter of the radiating fin on the temperature distribution of the magnetic conduction assembly, determining the inner diameter of the radiating fin, and if the amplitude of temperature flattening and the temperature drop degree are considered, properly increasing the inner diameter of the radiating fin on the basis, wherein R0The maximum distance between the radiating fin and the rotary central shaft;
5) the magnetic screen material is pure iron, and the performance of the magnetic field is evaluated based on the influence of temperature on the magnetic field; selecting discrete temperature point values in a certain temperature interval respectively to measure the magnetic field intensity of the magnetic screen material at the same position; wherein a temperature setting of around 200K has the largest contribution to the magnetic field enhancement of the magnetic shield material.
As a further improvement of the technology, in the step 1), the ion energy is lost by Pb＝IbVb，IbIs a beam current, VbGenerally more than 90% of the discharge voltage;
wherein the energy deposition P of the plasma to the wall surfaceiw＝IiwΔViwWherein ion current is incident on the wall surface
niIs the ion density, e is the number of electrons, viThe ion Bohm velocity is shown, A is the effective area of the wall surface of a discharge chamber, k is a Boltzmann constant, Te is the electron temperature, and M is the electron mass;
wherein the anode energy loss is Pa＝IaΔVa，IaFor anodic incident current, Δ VaAverage energy loss at the anode for incident electrons;
wherein the gas ionization energy is Pion＝(Ib+Iiw)U+，U+Represents the ion mean ionization voltage; and the loss of radiant energy PradIs a constant.
As a further improvement of the present technology, the error range threshold interval in step 3) is set to (9%, 9.5%);
as a further improvement of the technology, the size of the inner diameter of the radiating fin in the step 4) is preferably 0.45R0-0.55R0。
As a further improvement of the present technology, each discrete temperature point value within a certain temperature interval in the step 5) is specifically a temperature of 100K, 200K, 300K, …, 800K.
By adopting the magnetic shield thermal optimization method, the thermal load of the Hall thruster can be reduced, the temperature distribution and the heat flow trend of the magnetic shield can be optimized on the basis that the magnetic field intensity reaches a certain value, and a theoretical and practical basis is laid for the thermomagnetic correlation analysis of the Hall thruster in related researches.
Drawings
Fig. 1 is a schematic diagram of the total loss of energy distribution of a hall thruster.
FIG. 2 is a graph of the energy loss profile of the discharge chamber.
FIG. 3 is a calculated boundary-change condition.
FIG. 4 is a graph of heat flux density of the inner and outer walls of the discharge chamber.
Fig. 5 shows the heat flux density of the anode wall.
Detailed Description
As shown in fig. 1 and 2, the energy loss of the hall thruster mainly includes energy loss of a discharge chamber, energy loss of a cathode and energy loss of an excitation coil, wherein the energy loss of the discharge chamber mainly comes from ion energy, anode energy, energy deposition of plasma on a wall surface, gas ionization energy and radiation energy loss, and the energy loss accounts for 95% of the energy loss of the whole discharge chamber, so that the analysis of the energy loss has great significance for the analysis of the energy loss ratio of the hall thruster, and a foundation can be provided for the establishment of a subsequent thermal balance equation and the acquisition of a variable boundary condition.
As shown in fig. 3, the heat flux density of the ions to the wall surface of the discharge chamber and the heat flux density of the electrons to the anode wall surface are obtained by PIC/MCC calculation results, and the particle simulation algorithm related to the hall thruster in the field can be referred to for the PIC/MCC algorithm, the wall surface shell model and the anode electron deposition model. Fig. 4 and 5 show the heat flow density of the inner and outer walls of the discharge vessel and the heat flow density of the anode wall, respectively, which are dependent on the longitudinal length of the inner/outer wall or anode wall in the relevant direction.
The specific embodiments of the present application are as follows:
a method for reducing the thermal load of a Hall thruster is characterized by comprising the following steps: the method comprises the following steps:
1) establishing a thermal balance model; the energy loss of the Hall thruster mainly comprises energy loss of a discharge chamber, energy loss of a cathode and energy loss of an excitation coil, wherein the energy loss of the discharge chamber mainly comes from ion energy, anode energy, plasma wall surface energy deposition, gas ionization energy, radiation energy loss and the like;
2) according to the thermal balance model, solving the energy loss of the discharge chamber of the Hall thruster and the ratio of each energy loss to the total energy loss, and determining that the energy loss ratio of the wall surface of the discharge chamber is the highest;
3) establishing a discharge chamber wall surface heat conduction model by adopting a finite element analysis method, establishing a heat flow balance equation for each discrete node in the heat conduction model by combining a Fourier heat conduction law and an energy conservation law, simultaneously determining a boundary condition of the heat conduction, and verifying an error range of the heat conduction obtained based on the boundary condition through experiments; if the error range is within the allowed error range threshold value interval, judging that the simulation model has the capability of describing the temperature distribution of the thruster, and if not, re-determining the boundary condition until the error range is within the allowed error range threshold value interval;
4) analyzing the influence rule of the inner diameter of the radiating fin on the temperature distribution of the magnetic conduction assembly, determining the inner diameter of the radiating fin, and if the amplitude of temperature flattening and the temperature drop degree are considered, properly increasing the inner diameter of the radiating fin on the basis, wherein R0The maximum distance between the radiating fin and the rotary central shaft;
5) the magnetic screen material is pure iron, and the performance of the magnetic field is evaluated based on the influence of temperature on the magnetic field; selecting discrete temperature point values in a certain temperature interval respectively to measure the magnetic field intensity of the magnetic screen material at the same position; wherein a temperature setting of around 200K has the largest contribution to the magnetic field enhancement of the magnetic shield material.
As a further improvement of the technology, in the step 1), the ion energy is lost by Pb＝IbVb，IbIs a beam current, VbGenerally more than 90% of the discharge voltage;
wherein the energy deposition P of the plasma to the wall surfaceiw＝IiwΔViwWherein ion current is incident on the wall surface
niIs the ion density, e is the number of electrons, viThe ion Bohm velocity is shown, A is the effective area of the wall surface of a discharge chamber, k is a Boltzmann constant, Te is the electron temperature, and M is the electron mass;
wherein the anode energy loss is Pa＝IaΔVa，IaFor anodic incident current, Δ VaAverage energy loss at the anode for incident electrons;
wherein the gas ionization energy is Pion＝(Ib+Iiw)U+，U+Represents the ion mean ionization voltage; and the loss of radiant energy PradIs a constant.
As a further improvement of the present technology, the error range threshold interval in step 3) is set to (9%, 9.5%);
as a further improvement of the technology, the size of the inner diameter of the radiating fin in the step 4) is preferably 0.45R0-0.55R0。
As a further improvement of the present technology, each discrete temperature point value within a certain temperature interval in the step 5) is specifically a temperature of 100K, 200K, 300K, …, 800K.
The above examples are merely representative of preferred embodiments of the present invention, and the description thereof is more specific and detailed, but not to be construed as limiting the scope of the present invention. It should be noted that, for those skilled in the art, various changes, modifications and substitutions can be made without departing from the spirit of the present invention, and these are all within the scope of the present invention. Therefore, the protection scope of the present patent shall be subject to the appended claims.
Claims (3)
1. A method for reducing the thermal load of a Hall thruster is characterized by comprising the following steps: the method comprises the following steps:
1) establishing a thermal balance model; the energy loss of the Hall thruster mainly comprises energy loss of a discharge chamber, energy loss of a cathode and energy loss of an excitation coil, wherein the energy loss of the discharge chamber mainly comes from ion energy, anode energy, plasma wall surface energy deposition, gas ionization energy, radiation energy loss and the like;
2) according to the thermal balance model, solving the energy loss of the discharge chamber of the Hall thruster and the ratio of each energy loss to the total energy loss, and determining that the energy loss ratio of the wall surface of the discharge chamber is the highest;
3) establishing a discharge chamber wall surface heat conduction model by adopting a finite element analysis method, and establishing a heat flow balance equation for each discrete node in the heat conduction model by combining a Fourier heat conduction law and an energy conservation law, wherein the heat flow density of electrons to an anode is gradually increased along with the increase of the length of an anode wall surface, but when the length of the anode wall surface is increased to a certain value in an interval of 4-6mm, the heat flow density is gradually reduced along with the continuous increase of the length of the anode wall surface; determining boundary conditions of the heat conduction, and verifying error range of the heat conduction obtained based on the boundary conditions through experiments; if the error range is within the allowed error range threshold value interval, judging that the simulation model has the capability of describing the temperature distribution of the thruster, if not, re-determining the boundary condition until the error range is within the allowed error range threshold value interval, wherein the error range threshold value interval is (9%, 9.5%);
4) gauge for analyzing influence of inner diameter of radiating fin on temperature distribution of magnetic conduction assemblyDetermining the inner diameter of the radiating fin, if the amplitude of temperature flattening and the temperature drop degree are considered, properly increasing the inner diameter of the radiating fin on the basis, and finally obtaining the inner diameter of the radiating fin which is specifically 0.45R0-0.55R0Wherein R is0The maximum distance between the radiating fin and the rotary central shaft;
5) the magnetic screen material is pure iron, and the performance of the magnetic field is evaluated based on the influence of temperature on the magnetic field; selecting discrete temperature point values in a certain temperature interval respectively to measure the magnetic field intensity of the magnetic screen material at the same position; wherein a temperature setting of around 200K has the largest contribution to the magnetic field enhancement of the magnetic shield material.
2. The method for reducing the thermal load of the Hall thruster according to claim 1, wherein the method comprises the following steps: in the step 1), ion energy loss Pb＝IbVb，IbIs a beam current, VbGenerally more than 90% of the discharge voltage;
wherein the energy deposition P of the plasma to the wall surfaceiw＝IiwΔViwWherein ion current is incident on the wall surface
niIs the ion density, e is the number of electrons, viThe ion Bohm velocity is shown, A is the effective area of the wall surface of a discharge chamber, k is a Boltzmann constant, Te is the electron temperature, and M is the electron mass;
wherein the anode energy loss is Pa＝IaΔVa，IaFor anodic incident current, Δ VaAverage energy loss at the anode for incident electrons;
wherein the gas ionization energy is Pion＝(Ib+Iiw)U+，U+Represents the ion mean ionization voltage.
3. The method for reducing the thermal load of the Hall thruster according to claim 1, wherein the method comprises the following steps: the discrete temperature point value in the certain temperature interval in the step 5) is specifically the temperature of 100K, 200K, 300K, or 800K.
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| CN201810585807.4A CN108799033B (en) | 2018-06-08 | 2018-06-08 | Method for reducing heat load of Hall thruster |
| CN202010385673.9A CN111536007B (en) | 2018-06-08 | 2018-06-08 | Method for reducing heat load of Hall thruster |
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| KR20010088741A (en) * | 2001-08-28 | 2001-09-28 | 윤정웅 | Organic el display apparatus |
| EP2461643A1 (en) * | 2010-12-02 | 2012-06-06 | Alternative Heating Systems Inc. | Electrical safety grounding system |
| CN104632565A (en) * | 2014-12-22 | 2015-05-20 | 兰州空间技术物理研究所 | Hall thruster magnetic circuit structure |
| CN105889006A (en) * | 2016-05-03 | 2016-08-24 | 哈尔滨工业大学 | Hall thruster ceramic cooling support |
| US20170367168A1 (en) * | 2016-06-16 | 2017-12-21 | The Government Of The United States Of America, As Represented By The Secretary Of The Navy | Thermally isolated thermionic hollow cathodes |
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| CN2930182Y (en) * | 2006-07-05 | 2007-08-01 | 苏朝明 | Electric water heating bag |
| CN208441979U (en) * | 2018-06-19 | 2019-01-29 | 河南理工大学 | A kind of electromagnetism accelerating structure for Plasma propulsion device |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| KR20010088741A (en) * | 2001-08-28 | 2001-09-28 | 윤정웅 | Organic el display apparatus |
| EP2461643A1 (en) * | 2010-12-02 | 2012-06-06 | Alternative Heating Systems Inc. | Electrical safety grounding system |
| CN104632565A (en) * | 2014-12-22 | 2015-05-20 | 兰州空间技术物理研究所 | Hall thruster magnetic circuit structure |
| CN105889006A (en) * | 2016-05-03 | 2016-08-24 | 哈尔滨工业大学 | Hall thruster ceramic cooling support |
| US20170367168A1 (en) * | 2016-06-16 | 2017-12-21 | The Government Of The United States Of America, As Represented By The Secretary Of The Navy | Thermally isolated thermionic hollow cathodes |
Non-Patent Citations (1)
| Title |
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| 龙建飞等: "霍尔推力器热模型研究", 《强激光与粒子束》 * |
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| CN111536006B (en) | 2021-02-12 |
| CN111536006A (en) | 2020-08-14 |
| CN108799033B (en) | 2020-07-24 |
| CN111536007B (en) | 2021-02-12 |
| CN108799033A (en) | 2018-11-13 |
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Effective date of registration: 20221110 Address after: No. 99, Gangcheng Road, Dongying Port Economic Development Zone, Dongying City, Shandong Province 257237 Patentee after: Donggang Zhike Industrial Park Co.,Ltd. Address before: 310012 Room 401, 2 unit 155, Ma Shi street, Shangcheng District, Hangzhou, Zhejiang. Patentee before: HANGZHOU QICHENG SCIENCE & TECHNOLOGY Co.,Ltd. |