JPH086680B2 - Ion thruster and control method thereof - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION Object of the Invention (Industrial field of application) The present invention relates to a spacecraft and a space platform for navigating earth orbits, geostationary orbits, orbits reciprocating between them, interplanetary orbits, and the like. TECHNICAL FIELD The present invention relates to an ion thruster used for propulsion or orbit control of space structures such as the like and its control method.

(Prior Art) It has been attempted to use an ion thruster for the purpose of orbit control and propulsion of a space structure as described above. This ion thruster generates a propellant gas into an ionized plasma and electrically accelerates the propellant ions to obtain thrust. A conventional ion thruster is constructed, for example, as shown in FIG. Reference numeral 3 in the drawing is a discharge vessel, and an accelerating electrode composed of a screen grid 4 and an accelerating grid 5 is provided at the opening of the discharge vessel 3. Further, a part of the propellant gas is introduced into the discharge vessel 3 through the gas insulator 1a and the hollow cathode 2, and the remaining propellant gas is directly introduced into the discharge vessel through the gas insulator 1b. It is configured to be introduced. The propellant gas introduced through the hollow cathode 2 is formed into ionized plasma by discharge when passing through the hollow cathode. The electrons extracted from the ionized plasma are accelerated by the discharge vessel biased to the anode potential, collide with the propellant gas introduced into the discharge vessel through the gas insulator 1b, and enter the discharge vessel 3. Generates ionized plasma. The ionized propellant ions are accelerated by the accelerating electrodes 4 and 5 to be 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. Also, a magnet (not shown) is provided on the inner surface of the discharge vessel.
Is provided to enclose the ionized plasma by forming a cusp magnetic field and prevent the loss from the wall surface of the discharge vessel 3.

Further, 10 in the figure is a gas cylinder of a propellant gas introduction system, and after the pressure of the propellant gas supplied from this gas cylinder 10 is adjusted by the pressure reducing valve 11, each flow rate regulator 12a,
The hollow cathode 2, the gas isolator 1b, and the neutralizer 6 are supplied via 12b and 12c, respectively. Further, although not shown, a power supply system for supplying electric power to the hollow cathode 2, the discharge container 3, the acceleration electrodes 4, 5, the neutralizer 6 and the like is provided. Further, the pressure reducing valve 11, the flow rate adjusting valve and the like are controlled by a controller 13, and the 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 an ion thruster,
For example, it is considered to be used for the propulsion engine of an inter-orbital transport aircraft that reciprocates between a space station in an orbit around the earth and a stationary platform in a geostationary orbit, and for orbit control of a geostationary platform. When it is used for such an application, 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 also increases. By the way, the present and ion thrusters use Xe gas as a propellant. This Xe gas has a large atomic number, high efficiency, and since it is an inert gas, it does not damage each device of the ion thruster even when ionized, and it is an excellent propellant for such an ion thruster. . However, since this Xe gas is scarce in resources and expensive, it is disadvantageous in terms of resources and cost when used in a large amount as described above.

In addition to this Xe gas, it is possible to use an inert gas such as Ar or Kr gas as a propellant gas. However, since the atomic number of these gases is smaller than that of Xe gas, there is a problem that the efficiency is lowered. In addition, it is expected that semiconductor manufacturing factories will operate in outer space in the future, but in such a case a large amount of oxygen gas will be emitted from this space factories, so if this oxygen gas can be used as a propellant. Very advantageous in terms of resources and costs. However, since the plasma of oxygen gas has a high chemical activity, it causes a problem of damaging each device of the ion thruster.

Therefore, it is desired to overcome the above problems and use a substance other than Xe gas, preferably oxygen gas, as a propellant in response to a large thrust of an ion thruster 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 make it possible to use a gas other than Xe gas as a propellant. Another object of the present invention is to provide an ion thruster that can be operated as efficiently as possible as a whole propulsion system in consideration of a decrease 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 capable of preventing damage to equipment 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 first object described above, the propellant gas supply system of the ion thruster of the present invention can selectively introduce a plurality of kinds of gases. In addition, the power supply system changes the electric power to be supplied according to the type of propellant gas to be introduced, and the thrust is constant no matter which type of gas is introduced. Even if the kind of gas is changed, it is not necessary to change the structure of the ion thruster, for example, the distance between accelerating electrodes.

In order to achieve the second object, the ion thruster of the present invention introduces an inert gas from the hollow cathode when introducing oxygen as a propellant gas, and the oxygen gas does not pass through the hollow cathode. Directly introduced into the discharge vessel,
Electrons extracted from the plasma of the inert gas ionized by the discharge of the hollow cathode are accelerated by the bias potential of the discharge vessel and collide with the oxygen gas, and the oxygen gas is turned into plasma. This prevents the plasma of the oxygen gas from directly contacting the hollow cathode and other devices, and prevents the devices from being damaged by the chemically active oxygen plasma.

Further, in order to achieve the third object, the ion thruster of the present invention uses a gas having a large atomic number as a propellant when the electric power supplied from the space structure on which the ion thruster is mounted has no margin. When the electric power has a margin, the gas having a small atomic number is controlled to be used as the propellant. As a result, the entire propulsion system composed of this ion thruster can be operated as efficiently as possible.

(Example) Hereinafter, an example of the present invention will be described with reference to the drawings.

1 to 3 show a first embodiment of the present invention. The first embodiment is a case where a plurality of types of inert gas are used as the propellant gas.

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

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

An acceleration electrode is provided at the opening of the discharge container 3. This acceleration electrode is composed of a screen grid 4 and an acceleration grid 5. These grids 4,5
Is a thin metal plate in which a large number of electrode holes are formed, and they are arranged with a small gap therebetween. Electric power is supplied to the accelerating electrode from a power supply system described later, and the screen grid 4 is biased to a positive potential and the accelerating grid 5 is biased to a negative potential. And the above-mentioned discharge vessel 3
The ions of the propellant gas generated inside are accelerated by the potential difference between these grids when passing through the electrode holes of the screen grid 4 and the acceleration grid 5, are given kinetic energy, and are released to give thrust. Is configured to occur.

In addition, a plurality of annular magnets 7 are provided on the inner surface of the discharge vessel 3, a cusp magnetic field is formed by these magnets, and the generated electromagnetic plasma is confined, and the plasma contacts the wall of the discharge vessel 3. Is configured to prevent loss from this wall.

A neutralizer 6 is provided near the opening of the discharge vessel 3, and electrons are emitted from the neutralizer to neutralize the ions emitted from the opening of the discharge vessel. Devices such as the discharge container 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 described with reference to FIG. This propellant gas introduction system is provided with three propellant gas cylinders 10a, 10b, 10c. Then, the first cylinder 10a is filled with Xe gas, the second cylinder 10b is filled with Kr gas, and the third cylinder 10a is filled with Kr gas.
The cylinder 10c is filled with Ar gas. The gas supplied from these cylinders 10a, 10b, 10c is a pressure reducing valve that also serves as a switching valve provided corresponding to each of these cylinders.
11a, 11b and 11c are combined into a common pipe, and then branched to supply the hollow cathode 2, the gas diffuser 9 and the neutralizer 6 through the flow rate regulators 12a, 12b and 12c, respectively. It is configured to be. These pressure reducing valves 11a,
11b, 11c and the flow rate regulators 12a, 12b, 12c are configured to be controlled by the controller 13, respectively.
Is configured to operate according to commands from the body of the space structure on which this ion thruster is mounted.

Then, when instructed 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 closed,
Xe gas is supplied from the cylinder 10a. Also other Kr
Similarly, when instructed to use gas or Ar gas, only gas of the instructed type is selectively supplied.

The configuration of the power supply system of this ion thruster will be described with reference to FIG. A heater power source 21 is provided in this power source system, and the heater power source 21 supplies power to the heater for ignition of the hollow cathode. Further, 20 is a keeper power supply, and the keeper power supply 20 maintains the discharge in the hollow cathode 2. Further, 22 is a discharge power source,
The discharge vessel 3 is driven by the electric power supplied from the discharge power source 22.
The main discharge is performed inside. 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. Further, in the neutralizer 6, the ignition heater is heated by the heater power supply 25 for the neutralizer, and the discharge in the neutralizer 6 is maintained by the electric power supplied from the keeper power supply 26 for the neutralizer. Is configured. Each of these power sources is configured to be controlled by the controller 13. As described above, the 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 of the power supplies described above. It is controlled so that the thrust is constant regardless of which gas is used as the propellant. The heater power source 21, the keeper power source 20, and the discharge power source 22 are connected to the beam power source 23 on the negative electrode side.
Since it is connected to the positive electrode side of, the electrical insulator 27 is provided.

Next, the operation of the propellant gas introduction system and the power supply system when the type of gas used as the propellant is switched will be described. First, when using Xe gas, cylinder 10a
Only the pressure reducing valve 11a 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 that the maximum efficiency corresponding to this Xe gas can be achieved as in the conventional case.

Next, when Kr gas is used, only the pressure reducing valve 11b of the cylinder 10b is opened and the other pressure reducing valves are closed to supply this Kr gas. In this case, each of the above flow rate regulators 12a.12b, 12
Adjust c to make the flow rate of these gases about 1.3 times that of Xe gas. In addition, the discharge voltage of the discharge power supply 22 of the power supply system
Approximately 1.4 times that for gas. As a result, the ion current density of the ionization plasma in the discharge vessel is 1.3
It will be about double. By thus increasing the ion current density in the discharge vessel by about 1.3 times, the beam current of the beam power supply 23 and the acceleration current of the acceleration power supply 24 are about the same as those in the case of Xe.
It becomes 1.3 times.

When using Ar gas as the propellant, only the pressure reducing valve 11c of the cylinder 10c is opened and the other pressure reducing valves are closed,
This Ar gas is supplied. And in this case, the flow rate of each of the above flow rate regulators 12a, 12b, 12c is set to about 1.8 in the case of Xe gas.
Double. Further, the discharge voltage of the discharge power supply 22 of the power supply system is set to about 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 about 1.8 times that of Xe gas. As a result, the beam current of the beam power source 23 and the acceleration current of the acceleration power source 24 are about 1.8 times the case of Xe gas.

By the control as described above, the thrust force can be maintained substantially constant regardless of which type of propellant gas is introduced. That is, the thrust of such an ion thruster, when the type of the propellant gas to be introduced is changed and the flow rate is the same, the thrust of the thrust and the charge and mass of the plasma particles of the propellant gas. It is approximately proportional to the square root of the ratio of the ionized mass, which is the ratio. Therefore, when the type of propellant gas is changed as described above, the flow rate is controlled so as to be inversely proportional to the square root of the ratio of the ionized masses of those propellant gases, and the ion current corresponding to this flow rate is also changed. By controlling the density, that is, the beam current so as to be proportional to this flow rate, the thrust can be maintained substantially constant regardless of the type of propellant gas introduced.

Therefore, by appropriately selecting the type of propellant gas according to the power that can be supplied, the consumption of expensive Xe gas can be reduced, and the thrust of this ion thruster can be maintained constant. it can.

In this embodiment, since the inert gas is used as the plurality of types of propellant gas, the plasma of these propellant gases does not damage each device of this ion thruster.

In addition, these inert gases rarely form clusters of atoms even when formed into plasma, and the ionized mass of the plasma particles substantially corresponds to the atomic weight thereof. Therefore, even when the type of propellant gas introduced is changed, the flow rate of the propellant gas is controlled as described above in accordance with the atomic weight ratio, and the state of plasma flow of this propellant gas is controlled. Can be approximately similar. Therefore, it is not necessary to change the size and the like of each part of the ion thruster according to the type of gas, and different types of gas can be used as a propellant in the same ion thruster. This feature is particularly effective in that it is not necessary to change the spacing between the accelerating electrodes according to 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 electrode holes of the accelerating electrode screen grid 4 and accelerating grid 5, and when the type of propellant gas is changed and the state of the ion flow is changed, this It is necessary to change the distance between the screen grid 4 and the acceleration grid 5 in accordance with the above. Generally, the distance between the screen grid 4 and the acceleration grid 5 is as small as about 1 mm, so it is not easy to change the distance between them accurately and mechanically. However, in the case of this embodiment, since the ion flow state is the same regardless of the type of gas as described above, it is not necessary to change the interval between the screen grid 4 and the acceleration grid 5. It is extremely advantageous in that respect.

Further, the ion thruster of the present invention can use not only an inert gas but also oxygen gas as a propellant.
Since this oxygen gas is discharged as a by-product from a space semiconductor factory or the like, which is expected to be constructed in the future, the use of this oxygen gas as a propellant is extremely advantageous in terms of cost. However, the ions of this oxygen gas are
Due to its high chemical activity, it can damage the equipment of various parts of this ion thruster and must be considered in this regard.

FIG. 4 shows the propellant gas introduction system of the ion thruster of the second embodiment using oxygen gas as the propellant. This propellant gas introduction system is provided with an oxygen gas cylinder 10d and an inert gas cylinder 10e, the oxygen gas cylinder 10d is filled with oxygen gas, and the inert gas cylinder 10e is provided with Ar, K.
It is filled with an inert gas such as r or Xe. 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 source pressure valve 11e and the flow rate adjusters 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 structure as the above-mentioned embodiments except for the above points. Further, only the oxygen gas introduction system and the system related thereto are shown in FIG. 4, but when the inert gas other than the oxygen gas is selectively used as the propellant, the above-mentioned system is used. Of course, a similar propellant gas introduction system is added.

In this embodiment, ionized plasma of an inert gas is generated at the hollow cathode 2, and the electrons extracted from this plasma are accelerated by the bias potential of the discharge vessel and collide with oxygen gas, resulting in ionized plasma of this oxygen gas. To produce oxygen ions. The pressure in the hollow cathode 2 is about 1 torr, while the pressure in the discharge vessel 3 is 1 × 10 ⁻⁴ torr and the pressure in outer space is 5 × 10 ⁻⁶ torr.
The oxygen ions do not flow back into the hollow cathode 2 and only the inert gas ions come into contact with the hollow cathode 2. Therefore, the hollow cathode is not damaged by oxygen ions.

In addition, such an ion thruster controls which kind of propellant gas is used according to the electric power that can be supplied from the space structure on which the ion thruster is mounted, and the most efficient propulsion system as a whole. Is operated so that That is, when there is little margin in the power that can be supplied from the power source of the space structure on which this ion thruster is mounted, such as a solar cell, a propellant with a larger atomic number, such as Xe gas, is used. If the supplied electric power has a margin, a propellant having a smaller atomic number such as Kr or Ar gas is used. For example, using this Ar gas requires about 1.9 times more electric power to obtain the same thrust than using Xe gas, but in this case there is a margin of electric power, so the maximum You can get the thrust of. Therefore, by performing such control, the maximum thrust can always be obtained, and the entire propulsion system including this ion thruster can be operated with the maximum efficiency. Also, when using multiple ion thrusters in a cluster,
Different kinds of propellants may be used for some of the ion thrusters and other ion thrusters, and more flexible operation control can be performed corresponding to the power margin.

The present invention is not limited to the above-mentioned embodiment, and various modifications are possible. 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 the cusp magnetic field, etc. are specified. It can be appropriately changed according to the above. For example, when it is not necessary to change the type of propellant frequently, only one cylinder may be provided in the propellant gas introduction system and the type of propellant gas to be charged may be changed. Also, the means for generating the ionized plasma is not limited to that of the above-described embodiment, and is not limited to the electron impact type of cusp magnetic field confinement as described above, and may be a high-frequency type provided with a high-frequency power source.

[Advantages 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 are selected and used according to the margin of the power that can be supplied. It is possible to eliminate the resource limitation and cost limitation of the propellant. In addition, when changing the propellant, it is not necessary to change the ion thruster itself, and the structure is simple. Further, according to the control method of the present invention, the type of this propellant is selected according to the atomic number of the propellant in accordance with the surplus of the power that can be supplied, and the efficiency of the entire propulsion system including this ion thruster is maximized. The effect is great, such as maintaining the efficiency of.

[Brief description of drawings]

1 to 3 show a first embodiment of the present invention. FIG. 1 is a schematic diagram of a propellant gas introduction system, FIG. 2 is a schematic diagram of a power supply system, and FIG. 3 is a main body of an ion thruster. It is a perspective view which fractures | ruptures and shows a part of part. FIG. 4 is a schematic diagram of the propellant gas introduction system of the second embodiment. Further, FIG. 5 is a schematic view of a propellant gas introduction system of a conventional ion thruster. 2 ... Hollow cathode, 3 ... Discharge vessel, 4 ... Screen grid, 5 ... Acceleration grid, 9 ... Gas diffuser, 10a-10e ... Gas cylinder, 12a-12c ... Flow rate regulator, 13 ... Controller, 22 …… Discharge power supply, 23 …… Beam power supply, 24 …… Acceleration power supply

Claims (6)

[Claims] 1. A discharge vessel, a propellant gas introduction system for introducing a propellant gas into the discharge vessel, and a hollow for forming a part of the propellant gas introduced into the discharge vessel into ionized plasma. A cathode, an accelerating electrode that is provided in the opening of the discharge vessel and accelerates and emits ionized plasma formed in the discharge vessel, and a power supply system that supplies power to the hollow cathode, the discharge vessel, and the accelerating electrode, respectively. In the ion thruster provided with, the above-mentioned propellant gas introduction system includes a plurality of gas cylinders, a plurality of pressure reducing valves, and a plurality of flow rate regulators, and by selectively opening these pressure reducing valves, It is possible to selectively introduce different types of propellant gases from the gas cylinders into the discharge vessel, and the flow rate of the propellant gas to be introduced into the discharge vessel by the flow rate regulator. The amount can be adjusted according to its type, and the power supply system includes a discharge power supply that supplies discharge power to the hollow cathode and a beam power supply that supplies beam power to the acceleration electrode. The discharge power source and the beam power source are capable of adjusting the power to be supplied, and is provided with a controller for controlling the above-mentioned propellant gas introduction system and power supply system. Of the propellant gas to be introduced among the plurality of types of propellant gas, the signal for selectively opening the pressure reducing valve corresponding to the type of propellant gas to be introduced is output, and the propellant gas introduced to the above flow rate regulator is also output. Outputs a signal that adjusts the flow regulator so that the flow rate of the beam is inversely proportional to the square root of its ionized mass, and the controller supplies the beam power source described above to the discharge power source of the power source system. Propellant And it outputs a signal for controlling the discharge power to be proportional to the scan of the flow rate, ion thrusters, characterized in that thereby maintaining thrust even when introducing any propellant gas constant. 2. The ion thruster according to claim 1, wherein the propellant gas introduction system supplies plural kinds of inert gases as plural kinds of propellant gases. 3. The propellant gas introduction system introduces at least one kind of gas of an inert gas such as Ar, Kr, or Xe gas as a propellant gas, and the controller controls these gases. The beam current and the discharge voltage of the beam power of the above power supply system are controlled according to the square root of the atomic weight ratio of the propellant gas which is an inert gas, and the beam current is changed when Ar gas is introduced as the propellant gas. When Xe gas is introduced as a propellant gas, it is about 1.8 times, the discharge voltage is about 2.2 times, and when Kr gas is introduced as a propellant gas, the beam current is when Xe gas is introduced as a propellant gas. About 1.3
3. The ion thruster according to claim 2, wherein the discharge voltage is set to about 1.4 times. 4. The ion thruster according to claim 1, wherein the propellant gas introduction system introduces oxygen gas as one of the propellant gases. 5. The propellant gas introduction system introduces an inert gas through the hollow cathode, and introduces oxygen gas into the discharge vessel without going through the hollow cathode. The ion thruster according to claim 4. 6. An ion thruster in which one type is selected from a plurality of types of gas as a propellant gas and is introduced into the discharge vessel, the atomic number becomes smaller as the electric power that can be supplied from the structure in which the ion thruster is mounted is smaller. A large amount of gas is introduced as a propellant gas, and a gas having a smaller atomic number is introduced as a propellant gas as the power that can be supplied is increased.