阅读说明：本技术 一种降低霍尔推力器热负荷的方法 (Method for reducing heat load of Hall thruster ) 是由 李保平 于 2018-06-08 设计创作，主要内容包括：本发明属于霍尔推进器技术领域,尤其涉及一种降低霍尔推力器热负荷的方法。它主要包括有如下步骤：建立热平衡模型；求解霍尔推力器的放电室能量损耗；确定热传导的边界条件；确定散热片内径大小；基于温度对磁场的影响,对磁场的性能进行评估。采用本发明的磁屏热优化方法,可以在磁场强度达到一定值的基础上,降低霍尔推力器的热负荷、优化磁屏温度分布与热流走向,在相关研究中为霍尔推力器的热磁相关性分析奠定了一定理论与实践基础。(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.)

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 heat conduction model of the wall surface of the discharge chamber 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 ions to the wall surface of the discharge chamber is gradually increased along with the increase of the length of the inner/outer wall surface, but when the length of the inner/outer wall surface is increased to a certain value in a range of 15-20mm, the heat flow density is gradually reduced along with the continuous increase of the length of the inner/outer 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) analyzing the temperature of the magnetic conduction component according to the size of the inner diameter of the radiating finDetermining the inner diameter of the radiating fin according to the influence rule of the cloth, 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 to finally obtain 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 P_b＝I_bV_b，I_bIs a beam current, V_bGenerally more than 90% of the discharge voltage;

wherein the energy deposition P of the plasma to the wall surface_iw＝I_iwΔV_iwWherein ion current is incident on the wall surfacen_iIs the ion density, e is the number of electrons, v_iThe 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 P_a＝I_aΔV_a，I_aFor anodic incident current, Δ V_aΔ V is the average energy loss of incident electrons at the anode_iwAverage energy loss for incident ions;

wherein the gas ionization energy is P_ion＝(I_b+I_iw)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.

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 P_b＝I_bV_b，I_bIs a beam current, V_bGenerally more than 90% of the discharge voltage;

wherein the energy deposition P of the plasma to the wall surface_iw＝I_iwΔV_iwWherein ion current is incident on the wall surfacen_iIs the ion density, e is the number of electrons, v_iThe 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 P_a＝I_aΔV_a，I_aFor anodic incident current, Δ V_aAverage energy loss at the anode for incident electrons;

wherein the gas ionization energy is P_ion＝(I_b+I_iw)U⁺，U⁺Represents the ion mean ionization voltage; and the loss of radiant energy P_radIs 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 P_b＝I_bV_b，I_bIs a beam current, V_bGenerally more than 90% of the discharge voltage;

wherein the energy deposition P of the plasma to the wall surface_iw＝I_iwΔV_iwWherein ion current is incident on the wall surfacen_iIs the ion density, e is the number of electrons, v_iThe 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 P_a＝I_aΔV_a，I_aFor anodic incident current, Δ V_aAverage energy loss at the anode for incident electrons;

wherein the gas ionization energy is P_ion＝(I_b+I_iw)U⁺，U⁺Represents the ion mean ionization voltage; and the loss of radiant energy P_radIs 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.