JPH03141871A - Ion thruster and its control method - Google Patents

Abstract

PURPOSE:To use different sorts of propellants selectively by furnishing a propellant gas introduce system, which sets the rate of gas flow in accordance with the sort of propellant gas to be introduced to a discharge vessel. CONSTITUTION:In the case that Ar gas is used as propellant, a pressure reducing valve 11c of a gas container 10c solely opened, while the other pressure reducing valves are left closed, and Ar gas supply is carried out. Therein the rate of flow in each flow adjuster 12a-12c shall be approx. 1.8 times as large as when Xe gas is used, and the discharge voltage of a discharge power supply 22 belonging to the power supply system be approx. 2.2 times as large as when Xe gas is used. Thereby the ion current density of electric dissociated plasma in the discharge space becomes approx. 1.8 times as large as when Xe gas is used. As a result, the beam current of a beam power supply 23 and the acceleration current of an accelerating power supply 24 becomes approx. 1.8 times as large as when Xe gas is used.

Description

[Detailed Description of the Invention] [Object of the Invention] (Industrial Application Field) The present invention relates to spacecrafts and space platforms that navigate in Earth orbit, geostationary orbit, orbits that shuttle between these orbits, interplanetary orbits, etc. The present invention relates to an ion thruster used for propulsion or orbit control of space structures such as, and its control method.

(Prior Art) Conventionally, attempts have been made to use ion thrusters for the purpose of orbit control and propulsion of space structures such as those described above. This ion thruster generates propellant gas into ionized plasma and electrically accelerates the propellant ions to obtain thrust.

A conventional ion thruster is configured as shown in FIG. 5, for example. 3 in the figure is a discharge vessel, and an acceleration electrode composed of a screen grid 4 and an acceleration grid 5 is provided at the opening of the discharge vessel 3. A part of the propellant gas is introduced into the discharge vessel 3 via the gas insulator 1a and the hollow cathode 2, and the remaining propellant gas is directly introduced into the discharge vessel via the gas insulator 1b. configured to be installed. The propellant gas introduced through this hollow cathode 2 is formed into an ionized plasma by electric discharge as it passes through this hollow cathode. The electrons extracted from this ionized plasma are accelerated by the discharge vessel biased to the anode potential, collide with the propellant gas introduced into the discharge vessel via the gas insulator 1b, and enter the discharge vessel 3. Generates ionized plasma. The ionized propellant ions are accelerated by the acceleration electrode 4.5, given kinetic energy, and emitted to generate thrust.

In order to neutralize the emitted plasma, a neutralizer 6 is provided near the discharge vessel to neutralize the emitted ions. A magnet (not shown) is provided on the inner surface of the discharge vessel to form a cusp magnetic field to confine ionized plasma and prevent loss from the wall of the discharge vessel 3.

10 in the figure is a gas cylinder for the propellant gas introduction system, and the propellant gas supplied from this gas cylinder 10 has its pressure adjusted by a pressure reducing valve 11, and then passes through each flow rate regulator 12a, 12b, 12c. The gas is configured to be supplied to the hollow cathode 2, gas insulator 1b, and neutralizer 6, respectively, through the gas insulator 1b. Also, although not shown, the hollow cathode 2, the discharge vessel 3, and the accelerating electrode 4.

5. A power supply system is provided for supplying power to the neutralizer 6 and the like. In addition, the above-mentioned pressure reducing valve 11, flow rate adjustment valve, etc. are controlled by the controller 1.
3, and this controller 13 is operated by a command from the main body of the space structure or a station on the ground.

By the way, recently, such ion thrusters have been used, for example, as propulsion engines for interorbital transport vehicles that shuttle between a space station in Earth orbit and a stationary platform in geostationary orbit, and for orbit control of a stationary platform. is being considered. When used for such purposes, a large thrust is required, and it is necessary to increase the thrust of this ion thruster and to use a large number of ion thrusters in a cluster. In such a case,
Naturally, the amount of propellant consumed will also increase. By the way, current ion thrusters use Xe gas as a propellant. This Xe gas has a large atomic number and high efficiency, and since it is an inert gas, it will not damage the equipment of the ion thruster even if it is ionized, making it an excellent propellant for such ion thrusters. . However, since this Xe gas is a scarce resource and is expensive, it is disadvantageous in terms of resources and cost when used in large quantities as described above.

In addition to this Xe gas, Ar is used as a propellant gas.

It is possible to use an inert gas such as Kr gas.

However, since these gases have a smaller atomic number than Xe gas, there is a problem that the efficiency decreases.

In addition, it is expected that semiconductor manufacturing factories will operate in outer space in the future, and in such a case, a large amount of oxygen gas will be emitted from the space factory.If this oxygen gas could be used as a propellant, It is very advantageous in terms of resources and costs. However, since this oxygen gas plasma is highly chemically active, it causes problems in that it damages various components of the ion thruster.

Therefore, it is desired to overcome the above-mentioned problems and to use a substance other than Xe gas, preferably oxygen gas, as a propellant in order to cope with the increase in thrust of ion thrusters in the future.

(Problems to be Solved by the Invention) The present invention has been made based on the above circumstances, and a first object of the present invention is to enable the use of something other than Xe gas as a propellant, and to It is an object of the present invention to provide an ion thruster that allows the propulsion system as a whole to be operated as efficiently as possible, taking into account the reduction in efficiency when a substance other than the above is used as a propellant. A second object of the present invention is to provide an ion thruster that can prevent damage to equipment that would be expected when oxygen gas is used as a propellant. A third object of the present invention is to provide a method for efficiently controlling such an ion thruster.

[Structure of the Invention] (Means for Solving the Problems) In order to achieve the above-mentioned first object, the propellant gas supply system of the ion thruster of the present invention is capable of selectively introducing a plurality of types of gases. The power supply system changes the power supplied according to the type of propellant gas introduced.
It is configured so that the thrust is constant no matter what type of gas is introduced, and there is no need to change the structure of the ion thruster, such as the spacing between accelerating electrodes, even if the type of gas introduced is changed. It is something.

In addition, in order to achieve the second objective, when introducing oxygen as a propellant gas, the ion thruster of the present invention introduces an inert gas from the hollow cathode, and the oxygen gas does not pass through the hollow cathode. The electrons are directly introduced into the discharge vessel and extracted from the inert gas plasma ionized by the discharge of the hollow cathode, and are accelerated by the bias potential of the discharge vessel and collided with this oxygen gas. It is something that becomes. by this,
This oxygen gas plasma does not come into direct contact with the hollow cathode or other equipment, thereby preventing damage to the equipment due to the highly chemically active oxygen plasma.

In addition, in order to achieve the third objective, the ion thruster of the present invention uses a gas with a high atomic number as a propellant when there is insufficient power supplied from the space structure on which it is mounted, and If there is sufficient power, gases with lower atomic numbers are controlled to be used as propellants. This allows the entire propulsion system made up of this ion thruster to be operated as efficiently as possible.

(Example) Hereinafter, the present invention will be described in detail with reference to the drawings.

1 to 3 show a first embodiment of the present invention.

This first embodiment is a case where a plurality of types of inert gases are used as propellant gases.

First, the structure of the main body of this ion thruster will be explained with reference to FIG. 3 in the figure is a discharge vessel of this ion thruster, and this discharge vessel has a hemispherical shape with an open front. A hollow cathode 2 is provided in the inner part of the discharge vessel 3. This hollow cathode 2
A propellant gas is supplied to the hollow cathode from a propellant gas introduction system, which will be described later, via a gas insulator 1a for electrically insulating the discharge vessel 3 and other equipment. 2 and is introduced into this discharge vessel 3. Electric power is supplied to this hollow cathode 2 from a power supply system to be described later, and the propellant gas introduced through the hollow cathode 2 is configured to be formed into ionized plasma by electric discharge between the cathode and the keeper. There is.

Furthermore, an annular gas diffuser 9 is provided on the inner surface of the discharge vessel 3.
is provided. This gas diffuser 9 has a hollow passage shape, and a large number of nozzle holes 9a are formed in its wall surface. This gas diffuser 9 is connected to a propellant gas introduction system, which will be described later, via a gas insulator 1b. The gas supplied from the propellant gas introduction system is introduced into the discharge vessel 3 through the nozzle hole 9a of the gas diffuser 9 and diffused therein. Further, this discharge vessel 3 is biased to an anode potential, and accelerates electrons emitted from the ionized plasma of the propellant gas introduced from the hollow cathode 2, and accelerates the propellant introduced from the gas diffuser 9. The propellant gas is configured to collide with a gas and form the gas into an ionized plasma to generate propellant gas ions.

Furthermore, an accelerating electrode is provided at the opening of the discharge vessel 3. This accelerating electrode is composed of a screen grid 4 and an accelerating grid 5. These grids 4,5
This is a thin metal plate with a large number of electrode holes, which are arranged with small gaps between them. Electric power is supplied to this acceleration electrode from a power supply system to be described later, and the screen grid 4 is biased to a positive potential, and the acceleration grid 5 is biased to a negative potential. And the above discharge vessel 3
When the propellant gas ions generated in the screen grid 4 and the acceleration grid 5 pass through the electrode holes, they are accelerated by the potential difference between these grids, are given kinetic energy, and are emitted, producing thrust. is configured to occur.

Note that a plurality of annular magnets 7 are provided on the inner surface of the discharge vessel 3.
are provided, and these magnets form a cusp magnetic field, which confines the generated electromagnetic plasma, prevents this plasma from coming into contact with the wall of the discharge vessel 3, and is configured to prevent losses from this wall. ing.

Further, a neutralizer 6 is provided near the opening of the discharge vessel 3, and electrons are emitted from the neutralizer to neutralize ions emitted from the opening of the discharge vessel.

Further, these devices such as the discharge vessel 3 and the neutralizer 6 are housed in the thruster case 8.

Next, the configuration of the propellant gas introduction system of this ion thruster will be explained with reference to FIG. This propellant gas introduction system includes three propellant gas cylinders 10a.

10b and 10c are provided. The first cylinder 10a is filled with Xe gas, and the second cylinder 10b is filled with Xe gas.
is filled with Kr gas, and the third cylinder 10c is filled with Ar gas.

Gases supplied from these cylinders 10a, 10b, and 10c are supplied to pressure reducing valves 11a which also serve as switching valves provided corresponding to these cylinders.

11b and llc into a common pipe, which is further branched to flow rate regulators 12a and 12b.

The gas is configured to be supplied to the hollow cathode 2, gas diffuser 9, and neutralizer 6, respectively, via the gas tube 12c. These pressure reducing valves 11g, 1lb+11c and flow Jl regulators 12a, 12b, 12c are configured to be controlled by a controller 13, respectively, and this controller 13 It is configured to operate according to commands from the main body of the structure.

When a command is given to use, for example, Xe gas as the propellant gas, only the pressure reducing valve 11a corresponding to the cylinder 10a is opened, and the other pressure reducing valves 10b, 10c are opened.
is closed, and Xe gas is supplied from this cylinder 10a. Further, even when a command is given to use other Kr gas or A' gas, only the commanded type of gas is selectively supplied in the same manner.

Further, the configuration of the power supply system of this ion thruster will be explained with reference to FIG. This power supply system is provided with a heater power supply 21, and this heater power supply 21 supplies power to the heater for igniting the hollow cathode. Further, 20 is a keeper power supply, and the keeper power supply 20 maintains the discharge within the hollow cathode 2. Further, 22 is a discharge power source, and main discharge is performed within the discharge vessel 3 by power supplied from the discharge power source 22. A beam power source 23 is connected to the screen grid 4 of the accelerating electrode, and the beam power source 23 biases the screen grid to a predetermined positive potential. Further, the acceleration grid 5 is configured to be biased to a predetermined negative potential by the acceleration power supply 24. Furthermore, the neutralizer 6
The ignition heater is heated by the neutralizer heater power source 25, and the discharge inside the neutralizer 6 is maintained by the power supplied from the neutralizer keeper power source 26. Each of these power supplies is configured to be controlled by a controller 13. As described above, this controller 13 controls the type of gas to be supplied, adjusts the flow rate according to the type of gas to be supplied, and also controls the voltage and current of each power source, This controls the thrust so that it remains constant no matter what gas is used as the propellant. Note that the heater power supply 21, keeper power supply 20, and discharge power supply 22 described above
Since the negative electrode side of is connected to the positive electrode side of the beam power source 23, an electrical insulator 27 is provided.

Next, the operation of the propellant gas introduction system and power supply system when the type of gas used as the propellant is changed will be explained. First, when using Xe gas, use a cylinder of 10
Only the pressure reducing valve 11a of a is opened, and the other pressure reducing valves are closed. In this case, the gas flow rate and the electric power supplied from each power source of the power supply system are set so as to achieve the highest efficiency corresponding to this Xe gas, as in the conventional case.

Next, when using Kr gas, only the pressure reducing valve 11b of the cylinder 10b is opened, and the other pressure reducing valves are closed to supply the gas. 12
a, 12b, and 12c, and adjust these gas flow rates to
It should be about 1.3 times that of gas and gas. Further, the discharge voltage of the discharge power supply 22 of the power supply system is set to be approximately 1.4 times that in the case of Xe gas. As a result, the ion current density of the ionized plasma in the discharge vessel is approximately 183 times that of Xe gas.

By increasing the ion current density in the discharge vessel by approximately 1.3 times, the beam current of the beam power source 23 and the acceleration current of the acceleration power source 24 are each approximately 1.3 times that of Xe.
It will be doubled.

Further, when using Ar gas as a propellant, only the pressure reducing valve 11C of the cylinder 10c is opened, the other pressure reducing valves are closed, and this Ar gas is supplied.

In this case, each of the above flow rate regulators 12a.

Each flow rate of 12b and 12c is about 1.8 times that of Xe gas. Further, the discharge voltage of the discharge power supply 22 of the power supply system is set to be approximately 2.2 times that in the case of Xe gas. As a result, the ion current density of the ionized plasma in the discharge space is 1
．． It will be about 8 times as much.

As a result, the beam current of the beam power source 23 and the acceleration current of the acceleration power source 24 are approximately 11.8 times as large as those in the case of Xe gas.

As described above, by adjusting the flow rate, current, etc. in accordance with the type of propellant gas, the thrust can be made constant no matter what type of gas is used. Furthermore, in the above case, when using Kr gas or Ar gas, the gas flow rate and ion current density are set corresponding to the square root of the ratio of the atomic weight of Xe to the atomic weights of Kr and Ar. Therefore, in the case of Kr and Ar, the state of the ion flow is almost the same as in the case of Xe. therefore,
There is no need to change the dimensions of each part of the ion thruster depending on the type of gas, and different types of gases can be used as propellants in the same ion thruster. This feature is particularly effective in that it is not necessary to change the spacing of the accelerating electrodes depending on the type of propellant gas. That is, in order to improve efficiency, it is necessary to reduce the divergence angle of the ion beam passing through the screen grid 4 of the accelerating electrode and the electrode holes of the accelerating grid 5, and the type of propellant gas is changed to If the flow conditions change, these screen grids 4
It is necessary to change the spacing between the acceleration grid 5 and the acceleration grid 5. Generally, the spacing between the screen grid 4 and the acceleration grid 5 is as small as about 1 mm, so it is not easy to mechanically change the spacing between them accurately.

However, in this example, the state of ion flow is the same regardless of the type of gas, as described above.
This is extremely advantageous in that the spacing between the screen grid 4 and the acceleration grid 5 does not have to be changed.

Furthermore, the ion thruster of the present invention can utilize not only inert gas but also oxygen gas as a propellant. This oxygen gas will be emitted as a by-product from space semiconductor factories that are expected to be built in the future, so being able to use this oxygen gas as a propellant is extremely advantageous in terms of cost. However, since the ions of this oxygen gas are highly chemically active, they may damage various parts of the ion thruster, so this point needs to be taken into account.

FIG. 4 shows a propellant gas introduction system for an ion thruster according to a second embodiment that uses oxygen gas as a propellant. This propellant gas introduction system is provided with an oxygen gas cylinder 10d and an inert gas cylinder 10e.
is filled with oxygen gas, and an inert gas cylinder 10e
Ar.

It is filled with an inert gas such as Kr or Xe. and,
The gas supplied from the inert gas cylinder 10e is configured to be supplied to the hollow cathode 2 and the neutralizer 6 via the pressure reducing valve 11e1 and the flow rate regulators 12a and 12c. Further, the oxygen gas supplied from the oxygen gas cylinder 10d is configured to be directly introduced into the discharge vessel 3 via the gas diffuser 9. Note that this embodiment has the same configuration as the previous embodiment except for the above points. In addition, although only the oxygen gas introduction system and related systems are shown in FIG. 4, if an inert gas other than oxygen gas is selectively used as a propellant, the above-mentioned embodiment may be used. Of course, a similar propellant gas introduction system is also added.

In this embodiment, an ionized plasma of an inert gas is generated in the hollow cathode 2, and electrons extracted from this plasma are accelerated by the bias potential of the discharge vessel and collide with oxygen gas, resulting in an ionized plasma of the oxygen gas. and generate oxygen ions. The pressure inside the hollow cathode 2 is approximately 1 tOrr, whereas the pressure inside the discharge vessel 3 is approximately 1 tOrr.
The pressure inside is lXl0-'torr, and the pressure in space is 5
The pressure is approximately XIQ-6 torr, and oxygen ions do not flow back into the hollow cathode 2, and only inert gas ions come into contact with the hollow cathode. Therefore, this hollow cathode will not be damaged by oxygen ions.

In addition, the type of propellant gas used in such an ion thruster is controlled according to the power that can be supplied from the space structure on which the ion thruster is mounted.
This propulsion system as a whole is operated to achieve the highest efficiency. That is, when there is little margin in the power that can be supplied from the power source of the space structure on which the ion thruster is mounted, such as a solar cell, a propellant with a higher atomic number, such as Xe gas, is used. Further, if there is sufficient power to be supplied, a propellant with a smaller atomic number, such as Kr or Ar gas, is used. For example, when using Ar gas, approximately 1.9 times more power is required to obtain the same thrust than when using Xe gas, but this is only when there is sufficient power. , maximum thrust can be obtained. Therefore, by performing such control, the maximum thrust can be obtained at all times, and the entire propulsion system including the ion thruster can be operated with maximum efficiency. In addition, when multiple ion thrusters are used in a cluster, different types of propellants can be used for some of the ion thrusters and others, depending on the power margin. It is also possible to achieve more flexible operation control.

Note that the present invention is not limited to the above-described embodiments, and various modifications can be made, and the number of electrode plates of the accelerating electrode, the configuration of the propellant gas introduction system, the presence or absence of plasma containment by a cusp magnetic field, etc. are subject to specifications. It can be changed as appropriate depending on the situation. for example,
If it is not necessary to change the type of propellant frequently, the propellant gas introduction system may be provided with only one cylinder and the type of propellant gas filled may be changed. Furthermore, the means for generating ionized plasma is not limited to those of the above embodiments, and is not limited to the cusp magnetic field confinement electron impact type as described above, but may be of a high frequency type provided with a high frequency power source.

[Effects of the Invention] As described above, according to the present invention, a plurality of types of propellants can be selectively used as appropriate, and these can be selected and used according to the margin of power that can be supplied. Resource constraints and cost limitations on propellants can be resolved. Also,
There is no need to change the ion thruster itself when changing the propellant, and the structure is simple. Furthermore, according to the control method of the present invention, the type of propellant is selected according to the atomic number of the propellant in accordance with the margin of power that can be supplied.
This has great effects, such as being able to maintain the efficiency of the entire propulsion system including the ion thruster at maximum efficiency.

[Brief explanation of the drawing]

Figures 1 to 3 show a first embodiment of the present invention, with Figure 1 being a schematic diagram of the propellant gas introduction system, Figure 2 being a schematic diagram of the power supply system, and Figure 3 being the main body of the ion thruster. It is a perspective view which shows a part of a part broken. Moreover, FIG. 4 is a schematic diagram of a propellant gas introduction system of the second embodiment. Further, FIG. 5 is a schematic diagram of a propellant gas introduction system of a conventional ion thruster.

Claims (6)

[Claims] (1) A discharge vessel, a propellant gas introduction system that introduces propellant gas into the discharge vessel, and a hollow cathode that forms part of the propellant gas introduced into the discharge vessel into ionized plasma by electric discharge. , an accelerating electrode provided at the opening of the discharge vessel to accelerate and discharge ionized plasma formed within the discharge vessel, and a power supply system for supplying power to the hollow cathode, the discharge vessel, and the accelerating electrode. In the above-described propellant gas introduction system, the propellant gas introduction system selectively introduces multiple types of propellant gases into the discharge vessel and adjusts the flow rate of the propellant gas to be introduced in accordance with the type of propellant gas introduced. In addition, the above power supply system controls the discharge power and acceleration beam amount in accordance with the type and flow rate of the propellant gas introduced, and maintains a constant thrust no matter which propellant gas is introduced. An ion thruster that maintains (2) The ion thruster according to claim 1, wherein the propellant gas introduction system selectively supplies a plurality of types of inert gases as the plurality of types of propellant gases. (3) The propellant gas introduction system described above uses A as the propellant gas.
The system selectively introduces at least two types of gases among r, Kr, and When Xe gas is introduced as a propellant gas, the discharge voltage is 1.8 times and 2.2 times that of the case where Xe gas is introduced as a propellant gas, and when Kr gas is introduced as a propellant gas, the beam current and acceleration current are 3. The ion thruster according to claim 2, wherein the ion thruster has a discharge voltage of 1.3 times and a discharge voltage of 1.4 times that of the case where the gas is introduced as a gas, and the thrust force is constant in any case. (4) The ion thruster according to claim 1, wherein the propellant gas introduction system introduces oxygen gas as the propellant gas. (5) When introducing oxygen gas, the propellant gas introduction system introduces the inert gas through the hollow cathode, and the oxygen gas is introduced into the discharge vessel without going through the hollow cathode. The ion thruster according to claim 4, wherein the ion thruster is introduced into the ion thruster. (6) When controlling an ion thruster that selectively introduces multiple types of gases as propellant gases, the smaller the electric power that can be supplied from the space structure to which the ion thruster is attached, the more gases with higher atomic numbers are used. A method for controlling an ion thruster, characterized in that the larger the power that can be supplied, the smaller the gas is introduced as a propellant gas.