JPS63149298A - Space missile charging controller - Google Patents

Abstract

(57) [Summary] This bulletin contains application data before electronic filing, so abstract data is not recorded.

Description

[Detailed Description of the Invention] [Industrial Application Field] The present invention relates to a spacecraft charge control device that controls charging by space environment plasma of a spacecraft such as an artificial satellite, particularly a geostationary orbit satellite. .

[Conventional technology]

Generally, spacecraft are exposed to a space plasma environment while in orbit. Figure 5 shows the state of charging in the unique state of a spacecraft in this space plasma environment. In FIG. 5, 1 is a spacecraft structure, 41 is a plasma quartet, 42 is a plasma positive ion, and 43 is a space plasma consisting of plasma electrons 41 and plasma positive ions 42.
Reference numeral 44 indicates electrical charges accumulated in the spacecraft structure 1.

The current density of plasma electrons 41 in the space environment is usually higher than the current density of plasma positive ions 42, and near geostationary orbit, these current densities have values as shown in Table 1 below.

Table 1 Therefore, the unique state of the earth is such that the spacecraft structure 1 accumulates a negative charge load 44 and becomes negatively charged.

In particular, solar activity increases and geomagnetic winds (substorms) occur.
When this occurs, plasma electrons 41 and plasma positive ions 42 with current densities as shown in the substorm column of Table 1 above flow into the vicinity of the geostationary orbit, and the degree of negative charging of the spacecraft structure 1 becomes large. There is a possibility that a discharge may occur. Such a band 1 of the spacecraft structure 1 causes propellant to adhere to the surface of the cover glass of the solar cell, reducing the efficiency of the solar cell, or causing malfunction or failure of onboard equipment due to discharge. Therefore, in order to extend the lifespan and increase the reliability of the spacecraft body 1, it is necessary to control the spacecraft body 1 to prevent such charging and discharging.

As a conventional spacecraft belt control device for controlling the charging of the spacecraft structure 1 as described above, for example, "Journ" is used.
al of 5pacecraft and l'Lo
ckets j Vol. 18 No. 5, P, 462 (1981
) is known, having the configuration shown in FIG. 7. In FIG. 7, 21a to 21e are metal cesium (CB), and 22a to 22c are cesium 21a.
A vaporizer for vaporizing ~21C, 2.3 a~23c are gas pipes through which the vaporized cesium 21a~21c passes, 2
4a124b is a sheath heater, 25a and 25b are coil heaters, 26a and 26b are anodes, 27 is a probe,
28 is a probe power supply that supplies a negative voltage to the probe 27;
29 is a discharge ffl, 30 is a main anode installed in the discharge chamber 29, 31 is an annular magnet installed in the discharge chamber 29 to increase the discharge efficiency of each cesium 21b, 21c, and 32 is a main anode installed in the discharge chamber 29. An extraction electrode for extracting generated cesium ions, 33 an acceleration electrode for accelerating the cesium ions extracted by the extraction north pole 32, 34 a neutralizer, 35 a cesium ion engine, 36 electrons, 37
is a cesium ion, 38 is an electron 36 and a cesium ion 3
It is a cesium plasma consisting of 7.

Next, the operation of the conventional spacecraft belt '1 control device will be explained. The neutralizer 34 generates cesium plasma 38 so that the spacecraft structure 1 does not cause negative potential fluctuations when a positive cesium ion beam is emitted from the cesium ion engine 35, and
It is used to neutralize the cesium ion beam by mixing it with the cesium ion beam and to prevent potential fluctuations in the spacecraft structure 1. Only the above-mentioned neutralizer 34 is used for the spacecraft 1's spacecraft 1. The cesium 21a in the neutralizer 34 is vaporized by the vaporizer 22a, passes through the gas pipe 23&, is further heated by the coil heater 25a, and reaches the vicinity of the anode 26a. At that time, from the sheath heater cathode 24a to the positive 42
Thermal particles drawn out due to the potential difference between 6a and 6a collide with cesium atoms, ionize the cesium atoms, and form anode 2.
Cesium plasma 38 is formed near 6a. When the spacecraft structure 1 is in a unique state and is negatively charged, d in the cesium plasma 38 due to the potential difference in charge r with space.
Only the particles 36 are ejected from the spacecraft structure 1 into space, and the electric potential of the spacecraft structure 1 changes to the positive side to control charging.

[Problem that the invention seeks to solve]

Since the conventional spacecraft charge control device is configured as described above, cesium ions 37 in the cesium plasma 38 accumulate in the spacecraft structure 1 and adhere to the surface of the spacecraft groove book 1. This poses a problem in that it causes deterioration of the characteristics of typical surface materials such as thermal 1U control materials and cover glasses of solar cells. Furthermore, since cesium is once vaporized and then used to generate discharge plasma, there is also the problem that the response speed of the neutralizer 34 is slow.

This invention was developed in order to solve these problems, and it is possible to avoid deterioration of the characteristics of the surface material of the spaceflight object due to contamination due to adhesion of cesium ions, etc., and to have a fast response speed. The object is to obtain a body belt '1 control device.

[Means for solving problems]

The spacecraft charge control device according to the present invention includes an electrically insulated hollow cathode provided on the surface of the spaceflight object, four keeper poles arranged in front of the hollow cathode, and a rare gas stored. a gas tank supplied from the gas tank, a flow rate regulator that adjusts the gas flow t+ supplied from the gas tank, the hollow cathode and the keeper electrode, and a voltage is applied between the hollow cathode and the keeper electrode to discharge the rare gas supplied through the inside of the hollow cathode. The spacecraft is equipped with a keeper power source that generates plasma, and an ammeter connected to the keeper power source and the space flight vehicle.

[Effect]

In the spacecraft belt 11E11tIJ control device of the present invention, rare gas plasma is generated by the voltage applied between the hollow cathode and the keeper power source, and the electrons in the plasma thus generated are emitted into space. , This controls the charging of the spaceflight object.

〔Example〕

FIG. 1 is a schematic configuration diagram showing a spacecraft charge control device which is an embodiment of the present invention. In Fig. 1, l is a space flight suspension body, 2 is a hollow cathode, 3 is a keeper electrode with a small hole in the center provided on the front surface of the hollow cathode 2, and 4 is between the hollow cathode 2 and the keeper electrode 3. A keeper power supply that applies voltage, 5 a heater that heats the inner wall cathode 2 to facilitate the supply of active elements, 6 a heater power supply that supplies current to the heater 5;
7 is an insulating tube that electrically insulates the hollow cathode 2 from the space flight object 1; 8 is a gas tank for storing gas; 9 is a rare gas atom such as xenon, argon, neon, krypton, etc. supplied from the gas tank 8; 10 1 is a valve that opens and closes the gas tank 8; 11 is a valve control mechanism that controls opening and closing of the valve 10; 12 is a pulp control mechanism of the valve control mechanism 11; 13 is a gas pipe through which rare gas atoms 9 pass; 14 is a rare gas atom 9. 15 is a flow rate regulator power supply for controlling the flow rate regulator 14; 16 is an ammeter connected to the keeper power supply 4 at one end and to the spaceflight body 1 at the other end; 17 is an ionized rare gas ion, 18 is an electron,
19 is a rare gas plasma consisting of rare gas ions 17 and electrons 18, and 20 is a plasma generation source. Note that the keeper power supply 4. Heater power supply 6° Valve control power supply 12. The operation and stopping of the flow regulator power supply 15 are controlled by commands from the ground, and signals from the flow regulator power supply 15 and the ammeter 16 are transmitted to the ground as telemetry.

FIG. 2 is a diagram illustrating a procedure for controlling the charging of a spacecraft in the spacecraft belt charging control device of FIG. 1.

Next, the operation of the spacecraft charge control device which is an embodiment of the present invention will be described. When the space flight object 1 enters the earth's unique state, the space flight object 1 becomes negatively charged by the space plasma 43 as shown in FIG. Power supply 1
2. Flow regulator' power supply 15. Heater source 6. and sends a command to activate the keeper power supply 4, and the gas tank 8
The rare gas atoms 9 are passed through the hollow cathode 2 from the inside of the hollow cathode 2.
and the key chisel electrode 3. When the heater 5 is heated and thermoelectrons are easily emitted from the tip of the hollow cathode 2, the DC voltage, AC voltage, or high frequency voltage applied between the hollow cathode 2 and the keeper electrode 3 from the keeper '1 source 4 A discharge occurs and rare gas plasma 19 is generated. The ions 18 in the generated rare gas plasma 19 are emitted into space due to the potential difference between space and the negatively charged space flight object 1, and the rare gas ions 17 are It remains at 1, and the 1st place of space flight suspension class 1 moves to the positive side. At this time, the ammeter 16 measures a negative current. Data on the current value measured by the ammeter 16 is transmitted to the ground as telemetry, and is used for checking the charging control of the space flight object 1. In this way, the 4th position of space flight cancellation body 1 shifts to the positive side, and finally the potential difference with space disappears. At this time 1. ,! Since the L particles 18 and the rare gas ions 17 both diffuse into space, the first flow total 16
The current value becomes zero. This state continues until the space flight rod structure 1 is removed from the unique state of the earth. When leaving the unique state of the earth, send a command from the ground to the space flight rod structure l, and send the keypad source 4. Heater power supply 6. Flow regulator power supply 15. And the valve control power supply 12 is stopped. In addition,
Even if the heater power supply 6 is stopped at the time when the discharge occurs and the rare gas plasma 19 is generated, the discharge continues. The procedure for controlling the charging of the space flight rod structure 1 as described above is shown in FIG. Here, the above keeper power supply 4. Heater power supply 6.
Flow regulator area source 15. The command to put each power source such as the valve control power source 12 into operation may be issued before the space flight rod structure 1 enters the unique state of the earth.

Above, we have explained the charging control when the space flight rod structure 1 enters the unique state of the earth. A potential difference occurs between the surface of the insulator and the space flight rod structure 1 (this is called a local band), and this local band can also be controlled by putting the plasma generation source 2.0 into operation.

The state of this control will be explained with reference to FIG. In FIG. 6, reference numeral 45 denotes a diode emitted from the surface of the space flight rod structure 1 by the incidence of sunlight, and 46 represents an insulating surface material attached to a part of the space flight rod structure 1. Even when the space plasma 43 is incident on the space flight rod structure 1 during sunlight, the Yuanseiko 45 is emitted from the surface of the space flight rod structure 1, and the density of the photoelectrons 45 emitted is approximately equal to that of the plasma electrons 41. Since the current density is approximately the same as the current density, charging generally does not occur. However, when a part of the surface of the space flight rod structure 1 is made of the insulating surface material 46 and this part is in the shadow, the original
Since the electrons 45 are not emitted, band charges 44 caused by the plasma electrons 41 are accumulated on the insulating surface material 46 and become negatively charged, so that a potential difference is created between the insulating surface material 46 and the space flight rod structure 1. arise. In this situation,
When the plasma generation source 20 installed in the spacecraft structure L is put into operation by a command from the ground, the rare gas ions 17 in the generated rare gas plasma 19 cause the insulating surface material 46 to become negatively charged. , the negative band 4 exchanges charge with the charge 44, and the charge is neutralized.

In the above embodiment, one gas tank 8 is made up of one gas tank 8.
Although the case of one plasma generation source 20 is shown, as in another embodiment of the present invention shown in FIG. 3, a plurality of plasma generation sources 20 are provided in each part of the space flight rod structure 1, 20 may be provided with a gas tank 8, and as in another embodiment of the present invention shown in FIG.
A configuration may be adopted in which one gas tank 8 is provided for a plurality of plasma generation sources 20.

〔Effect of the invention〕

As explained above, this invention is a spacecraft charging control device in which a hollow cathode is provided on the surface of a spacecraft groove body in an electrically insulated manner, a keeper electrode is disposed in front of the hollow cathode, and a rare gas A gas tank that stores gas, a flow rate regulator that adjusts the gas flow rate supplied from the gas tank, and one pressure applied between the hollow cathode and the keeper polarization, and the dielectric gas supplied through the inside of the hollow cathode The configuration includes a keeper 1 that discharges gas to generate plasma as a source, and an ammeter connected to this keeper power source and the space flight rod structure, so that contamination due to adhesion of cesium ions etc. can cause space flight. This has the excellent effect of not causing any deterioration of the characteristics of the surface material of the rod structure, allowing a fast operational response speed, and also allowing the construction of a long-life and highly reliable spacecraft system.

[Brief explanation of the drawing]

FIG. 1 is a schematic configuration diagram showing a spacecraft charge control device which is an embodiment of the present invention, and FIG. 2 shows a procedure for controlling the charge of a spacecraft rod structure in the spacecraft charge control device shown in FIG. 3 and 4 are schematic configuration diagrams showing a spacecraft charging control device which are other embodiments of the present invention, and FIG. 5 is a diagram showing the state of charging in a unique state of a spacecraft. , Figure 6 is a diagram showing the local band during sunlight of a spacecraft partially covered with an insulating surface material, Figure 7
The figure is a schematic configuration diagram showing a conventional spacecraft charge control device. In the figure, 1... Spacecraft fin body, 2... Hollow cathode, 3... Keeper electrode, 4... Keeper power source, 5
... Heater, 6... Heater'; Source, 7... Insulating tube, 8... Gas tank, 9... Rare gas atom, 10...
・Valve, 11...Valve control mechanism, 12...Valve control power supply, 13...Gas pipe, 14...Flow rate regulator, 15...Flow i-adjustment regulator Hill source 16... Ammeter, 17...Rare gas ion, 18.9. Electronic, 19...
・Rare gas plasma, 20... plasma generation source, 21a
~21c... Cesium, 22a~22c... Vaporizer, 23a ~23e-... Gas pipe, 24a, 24b...
- Sheath heater cathode, 25a, 25b... coil heater, 26a, 26b... anode, 27... probe,
28...Probe power supply, 29...Discharge chamber, 30...
- Main anode, 31... Annular magnet, 32... Extraction electrode, 33... Accelerating electrode, 34... Neutralizer, 35...
Cesium ion engine, 36...electron, 37...
Cesium ion, 38... Cesium plasma, 41.
...Plasma electron, 42...Plasma positive ion, 43
... Space plasma, 44... Straight line charge, 45...
Photoelectron, 46... Insulating surface material. In each figure, the same reference numerals indicate the same or complementary parts.

Claims (3)

[Claims] (1) A hollow cathode provided electrically insulated on the surface of the spacecraft structure, a keeper electrode provided in front of this hollow cathode, a gas tank in which gas is stored, and gas supplied from this gas tank. a flow rate regulator that adjusts the flow rate of the gas; a keeper power source that applies a voltage between the hollow cathode and the keeper electrode to discharge the gas supplied through the hollow cathode to generate plasma; A spacecraft charge control device comprising: an ammeter connected to the keeper power source, and an ammeter whose other end is connected to the spacecraft structure. (2) The spacecraft charge control device according to claim 1, wherein the gas supplied from the gas tank is a rare gas such as xenon, argon, neon, or krypton. (3) The spacecraft charging control device according to claim 1 or 2, wherein the voltage applied from the keeper power source is a DC voltage, an AC voltage, or a high frequency voltage.