JP6029114B2 - Electric propulsion device - Google Patents

Description

  The present invention relates to an electric propulsion device that generates a propulsive force in outer space, and more particularly to an electric propulsion device that uses a Hall Effect Thruster that is one of electrostatic acceleration types.

  In recent years, in the field of electric propulsion devices that generate propulsive force in outer space, the purpose of space transport aircraft for construction of large space structures such as solar power generation satellites, inter-orbital transportation near the moon and the earth, or manned planetary exploration Research and development of large-scale electric propulsion devices with the development of the inter-orbit transport system in mind is underway. Electrostatic acceleration rockets such as Hall thrusters and ion engines require an electron source that generates electrons for plasma generation and ion beam neutralization. It greatly affects the service life.

  Patent Document 1 exemplified below discloses an electric propulsion apparatus using such an ion engine, and Patent Document 2 and Non-Patent Document 1 disclose the configuration of an electric propulsion apparatus using a Hall thruster. Has been.

  The ion engine extracts ions from plasma generated using an electrostatic field and accelerates them. Therefore, the ion engine is characterized by high ejection speed (high specific thrust 3000-10000 seconds), that is, low fuel consumption. The ion beam density, which is a reference for thrust per area, is limited. In contrast, the Hall thruster has a propulsive efficiency exceeding 50% in a specific thrust of 1500 to 2000 seconds, and the ion acceleration region is kept electrically quasi-neutral and is not affected by the space charge restriction law. High thrust density can be obtained. For these reasons, the Hall thruster has an excellent balance of performance expressed by thrust, specific thrust, propulsion efficiency, etc., and can be miniaturized, and is attracting the most attention as a future main propulsion device.

Japanese Patent No. 4925132 Japanese Patent No. 5295423

Semenkin, A., Kochergin, A., Garkusha, V., Chislov, G. and Rusakov, A.: RHETT / EPDM Flight Anode Layer Thruster Development, IEPC Paper 97-106, 1997 Yamamoto, N .; , Komurasaki, K. et al. and Arakawa, Y .: Discarge Current Oscillation in Hall Thrusters, Journal of Propulsion and Power, 21 (2005), pp. 870-876 Shigeru Yokota, Shinsuke Yasui, Ken Kumakura, Kimiya Koshiba, Yoshihiro Arakawa: Numerical analysis of sheath structure and discharge current inside anode layer type Hall thruster, Proceedings of the Japan Aerospace Society, 54 (2006), pp.30-44 Tony Schonherr, Rei Kawashima, Hiroyuki Koizumi and Kimiya Komurasaki, “Design and Performance Evaluation of Thruster with Anode Layer UT-58 for High-Power Application,” The 33rd International Electric Propulsion Conference, Oct. 2013, IEPC-2013-242

However, in particular as disclosed in Patent Document 1, such a Hall thruster, in particular an anode layer type Hall thruster also called sheath (T hruster with A node L ayer : TAL) In the discharge oscillation discharge current fluctuates There is a problem that the plasma is not formed stably because the discharge is unstable and is relatively large in most operating parameter regions.

  Many studies have been made so far to elucidate and reduce the oscillation phenomenon of the discharge current. However, in the case of the anode layer type hall thruster, a hollow shape that fits into the annular discharge space is shown. The most typical example is the use of a hollow anode (see Non-Patent Documents 1 to 3). Even when this is used, the discharge vibration can be reduced to some extent, but the vibration can be further reduced. It is desired. In addition, there is a problem that an operating parameter region in which plasma can be stably formed is narrow.

  Further, as a thrust generation process of the Hall thruster, it is necessary to supply electrons into the annular discharge space, and it is also necessary to supply electrons for neutralizing the emitted ion beam. For this reason, the electron source which generates the electron for using for these is provided. In this electron source, a method of extracting electrons from plasma is mainly widely used, and is roughly classified into a direct current discharge type, a high frequency discharge type and a microwave discharge type by thermionic emission. In general, the DC discharge type called Hollow Cathode has been widely adopted because it has a relatively sufficient electron emission capability at a low gas flow rate and low power, and has a track record of operation in outer space. Many.

  However, since the performance and life of the hollow cathode deteriorate due to oxygen or the like, strict management is required. In addition, it has excellent electron emission performance, such as the oxide cathode is likely to deteriorate, there are many factors that limit the lifetime, and there are large individual differences in the lifetime. In particular, there are various restrictions, and as the deterioration progresses, there is a problem that a long time is required for plasma ignition.

  The present invention has been made to solve such problems, and its purpose is to further reduce the discharge oscillation and to change the operating parameters such as the propellant flow rate, the anode voltage, and the magnetic field strength. It is another object of the present invention to provide an electric propulsion device that can provide a stable operation region that is larger than before.

  As a result of intensive studies to solve the above problems, the inventors of the present invention have realized an increase in current of an inductively coupled plasma electron source, which has generally been difficult to increase as an electron source. Paying attention to the disposition relationship with the coupled plasma electron source, disposing the hole thruster and the electron generation source in a specific electrical coupling relationship sufficiently suppresses the above-mentioned discharge vibration and stabilizes The inventor has found that the operating region is expanded, and has completed the present invention.

That is, the present invention provides the following inventions.
1. An electric propulsion device that introduces a propellant from one axial direction of an annular discharge space, accelerates ions generated from the propellant by an electric field, and discharges it from the other output end of the discharge space to obtain thrust There,
A hall thruster having a magnetic circuit for forming a magnetic field in a radial direction of the discharge space and having an anode electrode in the discharge space that contributes to the formation of the electric field;
A cylindrical discharge vessel disposed near the output end of the hall thruster and supplied with a working gas, and an induction coil disposed along the discharge vessel, for supplying high-frequency power to the induction coil The plasma of the working gas in the discharge space in the discharge vessel, and an inductively coupled plasma electron source that emits the generated electrons.
An electric propulsion device, wherein the Hall thruster and the inductively coupled plasma electron source are arranged so that the anode electrode of the Hall thruster and the inductively coupled plasma electron source are electrically capacitively coupled.
2. 2. The electric propulsion device according to 1, wherein the inductively coupled plasma electron source is formed by an inner coil type rf cathode in which the induction coil is disposed in the discharge vessel.
3. 2. The electric propulsion device according to 1, wherein the hall thruster is an anode layer type hall thruster in which the anode electrode is disposed in the vicinity of the output end.
4). 2. The electric propulsion apparatus according to claim 1, further comprising a direct current discharge type electron source that emits thermoelectrons by heating the cathode by direct current discharge.
5. 2. The electricity according to claim 1, wherein a plurality of the Hall thrusters are provided, and one or more inductively coupled plasma electron sources are electrically capacitively coupled to the plurality of Hall thrusters. Propulsion device.

The electric propulsion device according to the present invention further reduces the discharge oscillation, and even when the operation parameters such as the propellant flow rate, the anode voltage, and the magnetic field strength are changed, a more stable operation region can be obtained than before. It is what
More specifically, the electric propulsion apparatus according to the present invention employs an inductively coupled plasma electron source as an electron source. This inductively coupled plasma electron source is generally considered to be difficult to obtain a large amount of electron emission because there is little change in the amount of electrons emitted even when the high frequency power supplied is increased. However, as a result of research, it has been clarified that there is a transition condition in which the amount of emitted electrons increases dramatically when the high-frequency power is increased. Even when emphasizing the electron emission capability by utilizing such properties, it has been found that it is possible to use an inductively coupled plasma electron source instead of a hollow cathode as an electron source. By using it, all the faults at the time of using a hollow cathode as mentioned above can be wiped out. The advantages of the inductively coupled plasma electron source are that it can be ignited immediately, the ignition procedure is simple, the structure is simple, and it can contribute to the reduction of the number of power supplies, and the deterioration easily proceeds in the environmental environment on the ground. Since there is no such a constituent member, it is easy to manage deterioration on the ground, and it is possible to enjoy various advantages such as high power efficiency.

  In the state where the anode electrode of the Hall thruster and the inductively coupled plasma electron source are electrically capacitively coupled, when the plasma is ignited by supplying high frequency power to the inductively coupled plasma electron source, the anode is adjacent on the adjacent Hall thruster side. Even when no power is supplied to the electrodes or the magnetic circuit, plasma brightness is observed in the annular discharge space where the propellant is introduced. In the present invention, by providing an inductively coupled plasma electron source with respect to the Hall thruster so that they are electrically capacitively coupled to each other in this way, the discharge vibration is not generated when the Hall thruster is actually operated. It has been experimentally found that it is greatly reduced. As a result of sufficiently suppressing the discharge oscillation in this way, it becomes possible to form plasma more stably in the discharge space of the Hall thruster, coupled with the effect of obtaining a large amount of electron emission from the inductively coupled plasma electron source, Even when the operation parameters such as the flow rate, the anode voltage, and the magnetic field strength are changed, the stable operation region in which the Hall thruster operates stably can be expanded.

It is a mimetic diagram showing the whole system configuration of an electric propulsion device. It is a perspective view showing the appearance of an anode layer type hall thruster. It is an axial section of an anode layer type hall thruster, and is a sectional view showing an enlarged view of only one side in the radial direction from the central axis. It is a perspective view which shows each member which comprises a pressurization chamber. It is a graph which shows the relationship between the high frequency electric power supplied to RF cathode, and electron emission amount (anode current). It is a graph which shows the change of the electron emission amount (anode current) of RF cathode when the kind of working gas and an anode voltage are changed. It is the observation photograph which copied each emission side end face of a TAL and RF cathode. It is a graph which shows the result of having measured the discharge current and anode voltage which flow between an anode electrode and ground potential. It is a chart which shows the operation | movement sequence of an electric propulsion apparatus. It is the observation photograph which copied each emission side end face of a TAL and RF cathode. It is the observation photograph which copied each emission side end face of a TAL and RF cathode. (A), (b) is the observation photograph which copied the mutual emission side end surface of TAL and RF cathode. It is the observation photograph which copied each emission side end face of a TAL and RF cathode. (A) is a graph which shows the change of the discharge current in RF cathode, (b) is a graph which shows the change of the discharge current in a hollow cathode. It is a graph which shows the stable operation area | region of TAL at the time of using RF cathode / hollow cathode for an electron source. It is a graph which shows a stable operation area | region. (A)-(d) is a graph which shows the result of having evaluated operation stability by using magnetic flux density and an anode voltage as parameters, when each supplying a predetermined propellant flow rate. In addition, (d) is a result at the time of using a hollow cathode. (A)-(c) is a graph which shows the result of having evaluated operation stability by setting magnetic flux density and a propellant flow volume as a parameter, when a predetermined anode voltage is given, respectively. In addition, (c) is a result at the time of using a hollow cathode. (A), (b) is a graph which shows the result of having evaluated the operation stability by fixing the propellant flow volume and anode voltage to be given to predetermined levels, and using magnetic flux density and discharge current as parameters. In addition, (b) is a result at the time of using a hollow cathode. (A)-(c) is a graph which shows the result of having evaluated the operation | movement stability by setting magnetic flux density and a xenon flow rate as a parameter, when each given anode voltage is given. It is a schematic diagram which shows roughly the arrangement | positioning state of a Hall thruster, RF cathode, and a hollow cathode. It is a schematic diagram which shows roughly the arrangement | positioning state of a Hall thruster and RF cathode. It is a schematic diagram which shows roughly the arrangement | positioning state of a Hall thruster and RF cathode. It is a schematic diagram which shows roughly the arrangement | positioning state of a Hall thruster and a hollow cathode. It is a schematic diagram which shows roughly the arrangement | positioning state of a Hall thruster, RF cathode, and a hollow cathode.

1: electric propulsion device, 10: vacuum chamber, 20: DC power supply, 30: xenon tank, 40: high frequency power supply, 50: DC power supply, 100: anode layer type Hall thruster (TAL), 110: inner magnetic pole, 120: Outer magnetic pole, 140: output end, 160: anode electrode, 190: discharge space, 200: RF cathode (inductively coupled plasma electron source), 300: hollow cathode

  FIG. 1 schematically shows a system configuration of the electric propulsion apparatus according to the embodiment. The electric propulsion device 1 includes an anode layer type hall thruster (hereinafter referred to as “TAL”) 100 in a main body portion that generates thrust, supplies electrons to the TAL 100, and generates an ion beam emitted from the TAL 100. The electron source to be summed includes an RF cathode 200 as an inductively coupled plasma electron source, and these main components are arranged in the vacuum chamber 10.

  As shown in FIGS. 2 and 3, the TAL 100 has a cylindrical inner magnetic pole 110 (see FIG. 3) in the center, and an annular ring is provided between the outer periphery and the inner magnetic pole 110. The six outer magnetic poles 120 are arranged at equal intervals and in such a manner that the six outer magnetic poles 120 are positioned on one circumference. The inner magnetic pole 110 is configured by winding a solenoid coil 114 around an iron core 112, and the outer magnetic pole 120 is similarly configured by winding a solenoid coil 124 around an iron core 122.

  One end side in the axial direction of the iron core 112 of the inner magnetic pole 110 is fixed to an iron disc-shaped return yoke 130 having a larger diameter than the inner magnetic pole 110, and the other end side of the inner magnetic pole 110 is fixed to the inner magnetic pole 110. An iron disk-shaped inner pole piece 132 formed to be approximately the same size as the pole diameter is fixed. Further, one axial end side of the iron core 122 of each outer magnetic pole 120 is also fixed to the return yoke 130, and the other end side of each iron core 122 is fixed to an outer pole piece 134 formed in an iron flat ring shape. doing.

  Each solenoid coil 114, 124 is individually connected to a DC power source 20 (see FIG. 1, and the connection form is omitted because a known connection form can be used without particular limitation). , 124 is supplied with a coil current, whereby the inner magnetic pole 110 and each outer magnetic pole 120 are excited. With such a configuration, the inner magnetic pole 110, the outer magnetic pole 120, the return yoke 130, the inner pole piece 132, and the outer pole piece 134 constitute a magnetic circuit, and between the inner pole piece 132 and the outer pole piece 134. In the annular gap formed in the above, a radial magnetic field can be formed radially between the inner peripheral side and the outer peripheral side. The annular gap is a part that forms a discharge space as will be described later, and the part located on the other axial side (the other end side) is an output end 140 from which an ion beam is output from the TAL 100.

  In the annular space formed between the inner pole piece 132 and the outer pole piece 134, the propellant gas supplied from one place is uniform in the circumferential direction in order from the return yoke 130 side, as will be described later. And an anode electrode 160 that forms an annular discharge space that is open at one end, both of which are made of a copper material. Further, the copper preload chamber 150 and the anode electrode 160, and the inner magnetic pole 110, the outer magnetic pole 120, and the return yoke 130 disposed so as to surround them are covered with a ceramic material 170, The preload chamber 150 and the anode electrode 160 are electrically insulated from other peripheral components.

  As shown in FIG. 4, the supply preparatory chamber 150 is a chamber formed of an annular bottom plate 152, a cylindrical inner wall plate 154 provided on the inner peripheral side thereof, and a cylindrical outer wall plate 156 provided on the outer peripheral side thereof. It is demarcated by members. Further, an annular partition plate 158 (see FIG. 3) is disposed on the anode electrode 160 side. As shown in FIG. 3, the supply preliminary chamber 150 partitions the open end of the chamber forming member shown in FIG. Comparted in a state of being blocked by the plate 158. The bottom plate 152 is formed with an opening 151 for introducing a propellant into the inside, and the opposing partition plate 158 has a small diameter that penetrates the partition plate 158 in the thickness direction as shown in FIG. The introduction holes 158a are formed at a plurality of locations along the circumferential direction at regular intervals.

  The opening 151 of the bottom plate 152 is connected to a xenon tank 30 filled with xenon as a propellant via a gas supply line system 36 (see FIG. 1), and the xenon guided from the xenon tank 30 is Then, the gas is introduced into the supply preliminary chamber 150 from the opening 151 via the flow rate adjusting unit 32, the high-frequency leakage blocking unit 34, the gas isolator 38, and the like. The xenon introduced into the supply preliminary chamber 150 is uniformly supplied into the discharge space 190 inside the anode electrode 160 from each through hole 158a of the partition plate 158.

  The anode electrode 160 includes a cylindrical inner annular electrode 162 provided on the inner peripheral side and a cylindrical outer annular electrode 164 provided on the outer peripheral side. The outer annular electrode 164 is provided on the supply preliminary chamber 150 side. A step portion is provided in the middle so that the diameter on the output end 140 side is smaller than that in FIG. The inner annular electrode 162 and the outer annular electrode 164 constitute the aforementioned output end 140 by fixing the end portion on the preload chamber 150 side to the partition plate 158 and closing it, and opening the other end side. Therefore, since the partition plate 158 is configured structurally and electrically integrally with the inner annular electrode 162 and the outer annular electrode 164, the partition plate 158 functions as a so-called partition plate and serves as a bottom portion of the anode electrode 160. Also works. An anode voltage is applied to the anode electrode 160 configured in this manner by the DC power supply 50, and an annular discharge space 190 is formed by a gap between the inner annular electrode 162 and the outer annular electrode 164, and its output end 140. An acceleration channel is formed in the vicinity.

  In addition, guard rings 180 extending in the discharge space 190 along the inner annular electrode 162 and the outer annular electrode 164 are provided on the inner peripheral side and the outer peripheral side of the output end 140. The two guard rings 180 are both formed in an annular shape made of stainless steel (SUS316L), and the guard ring 180 on the inner peripheral side is on the inner peripheral side of the inner annular electrode 162 and the base end is fixed to the inner pole piece 132. The guard ring 180 on the outer peripheral side is on the outer peripheral side of the outer annular electrode 164 and the base end portion is fixed to the outer pole piece 134. The two guard rings 180 are both electrically connected to the cathode side and kept at the cathode potential. Since the acceleration channel wall in the vicinity of the output end 140 is maintained at the cathode potential by the guard ring 180 in this way, the electrons act so as to go deeper in the discharge space 190 without colliding with the acceleration channel wall in the vicinity. .

On the other hand, the RF cathode 200 serving as the electron source of the TAL 100 has a simple component structure as schematically shown in FIG. 1, and an induction coil is formed inside the cylindrical discharge vessel 210 along the cylindrical shape. 220 is an inner coil type. In the present embodiment, the discharge vessel 210 and the induction coil 220 are both made of a nonmagnetic metal material.
A high frequency power supply 40 is connected to the induction coil 220 via the matching box 42, and high frequency power of 13.56 MHz is supplied to the induction coil 220. The induction coil 220 is fixed to the discharge vessel 210 side through a ceramic, and the discharge vessel 210 and the induction coil 220 are electrically insulated. Further, the same xenon as the propellant is used as the working gas, and the xenon gas is supplied into the discharge vessel 210 through the gas supply line system 36 connected to the xenon tank 30. In the discharge vessel 210, xenon is ionized to generate inductively coupled plasma, and the generated electrons are emitted to the outside of the discharge vessel 210 through the orifice 230. The generated ions are captured in the discharge vessel 210 and output to the outside as an RF electron current. As the working gas, it is possible to form a gas supply system different from the propellant and use a gas type different from the propellant.

  Here, the large current mode of the RF cathode 200 will be described with reference to FIG.

  When a high frequency power (13.56 MHz) supplied to the induction coil 220 was gradually increased using a xenon gas having a flow rate of 0.6 mg / s as the working gas with respect to the RF cathode 200, the high frequency power exceeded 200 W. Even in the region, the change in the anode current indicating the electron emission amount is small and is about several A. Conventionally, it has been generally considered that it is difficult to obtain a large amount of electron emission from an RF cathode due to such a phenomenon. However, as a result of experiments, it has become clear that there are transition conditions that dramatically increase the amount of electrons emitted. That is, as shown in FIG. 5, as an example, when the flow rate of xenon gas is 0.7 mg / s and the supplied high frequency power is increased to 300 W, the amount of electron emission increases dramatically, and the anode current is 20 A or more. It has been confirmed that a large current mode that suddenly increases appears.

  Therefore, based on this result, the high frequency power supplied to the RF cathode is set to 300 W, and the result when the type of the working gas and the anode voltage are changed in the state where the RF cathode is operated in such a large current mode is shown in FIG. It is shown in FIG. When xenon is used as the working gas and supplied at a flow rate of 1.0 mg / s and the anode voltage is increased, the anode current also tends to increase. For example, the anode voltage is about 25 V and the anode current is 20 A. . This sufficiently clears the electron emission amount of about 15 A required for a 5 kW class hall thruster that will be the base of future large electric propulsion devices. In addition, it was confirmed that an electron emission amount of 20 A can be obtained even when argon, which is easily available and expected to be practical, is used as a working gas, compared to rare and expensive xenon.

<Example>
Various measurements were performed using the RF cathode 200 operated in such a large current mode as an electron source of the TAL100. The diameter of the used TAL 100 is 174 mm, and the diameter of the discharge vessel 210 in the RF cathode 200 is 51 mm. The distance between the center portions of the emission side end faces of the TAL 100 and the RF cathode 200 was set to 90 mm.

  Further, at this time, the TAL 100 and the RF cathode 200 are disposed so as to be electrically capacitively coupled to each other. Here, the capacitive coupling state is energy transmission in which an electrical energy change on one side is transmitted as an electrical energy change to the other between two conductive members, and is similar to a capacitor in an electric circuit. It is a working state. Specifically, when the TAL 100 and the RF cathode 200 are electrically capacitively coupled, the phenomenon described below is observed.

FIG. 7 is an observation photograph showing the emission side end surfaces of the TAL 100 and the RF cathode 200. In the drawing, the side indicated as RF / C is the electron emission surface of the RF cathode 200, and the side indicated as HET is the emission surface on which the output end 140 of the TAL 100 is formed. In this state, the working gas xenon is supplied to the RF cathode 200 at a flow rate of 0.6 mg / s, and high frequency power of 13.56 MHz and 300 W is supplied and ignited. The brightness of the inductively coupled plasma generated in the discharge vessel 210 drawn out from the central orifice 230 is observed. In the opposite TAL100, no xenon is supplied, no anode voltage is applied, and no Hall magnetic field is formed, and no glitter is observed.
FIG. 8 shows the measurement result of the discharge current flowing between the anode electrode 160 and the ground potential and the anode voltage in this state.
From this result, the anode voltage is maintained at approximately 0 V, but the discharge current periodically changes with an amplitude of about 0.3 A, and the period is about 70 ns. This period substantially coincides with 73.74 ns, which is one period of the frequency 13.56 MHz of the high-frequency power source 40 operating on the RF cathode 200 side. This is considered to be a state in which the TAL 100 and the RF cathode 200 are capacitively coupled in the same manner as the capacitor in the electric circuit, and the vibrational electric energy change occurring on one RF cathode 200 side is the anode of the opposing TAL 100. It is considered that it was transmitted to the electrode 160 and detected as a change in the discharge current.

  FIG. 9 shows an operation sequence of the electric propulsion apparatus using the TAL 100 and the RF cathode 200 arranged in such a capacitively coupled state. Hereinafter, it demonstrates in order according to this operation | movement sequence.

  First, in S10, xenon is supplied to TAL100 at a flow rate of 2.0 mg / s. At this time, no voltage is applied to the anode electrode 160 on the TAL 100 side, and no current is supplied to the solenoid coils 114 and 124. In addition, no high-frequency power is applied to the RF cathode 200, and no light emission phenomenon due to plasma is observed.

  In subsequent S12, the RF cathode 200 is ignited. At this time, the working gas xenon is supplied to the RF cathode 200 at a flow rate of 0.5 mg / s, and high frequency power of 300 W is supplied from the high frequency power source 40. The results observed in this situation are shown in FIG. The brightness of inductively coupled plasma generated in the discharge vessel 210 is observed through the orifice 230 on the RF / C side in the figure. Further, at this time, a ring-shaped light plasma brightness is also observed in the discharge space 190 of the TAL 100 to which xenon is supplied. This ring-shaped light plasma brightness is much lower than the brightness when the ion beam is emitted from the output end 140 in a steady state after the TAL 100 operates normally. At this stage, xenon is supplied on the TAL 100 side, but the other power supply system is in a stopped state, and since the anode voltage is not supplied, no electric field is formed and no Hall magnetic field is formed. . However, as described above, since the TAL 100 side and the RF cathode 200 side are capacitively coupled, the capacitively coupled plasma that is capacitively coupled with the inductively coupled plasma formed in the RF cathode 200 is induced. Is considered to be excited in the discharge space 190 of the TAL 100 as a plasma with a low ionization degree.

  In subsequent S14, a coil current is supplied to the solenoid coils 114 and 124 by the DC power source 20, and a magnetic field B having a magnetic flux density of 20 mT is applied to the annular discharge space 190 in the radial direction. As a result, as shown in FIG. 11, the ring-like brightness observed in the discharge space 190 of the TAL 100 disappears, and the diffusion state of the plasma brightness extracted from the orifice 230 changes on the RF cathode 200 side. ing. FIGS. 12A and 12B show the result of observing the emission side end face of the TAL 100 in which the output end 140 is formed from the front while changing the observation angle. The magnetic flux density is shown with the direction from the center to the outside as positive, and as shown in FIGS. 12A and 12B, the diffusion direction of the plasma brightness extracted from the RF cathode 200 when the direction of the magnetic field B is reversed. Is also reversed. This also indicates that the TAL 100 side and the RF cathode 200 side are capacitively coupled. Although the TAL 100 side and the RF cathode 200 side are considered to be capacitive coupling based on the observed phenomenon and measurement result, inductive coupling is also generated in a complex manner under a situation involving high frequency. There is also a possibility, and it does not exclude that there is an electrical coupling relationship other than capacitive coupling.

  In continuing S16, the direct-current power supply 50 is operated and the application of an anode voltage is started. That is, the output voltage of the DC power supply 50 is increased and the voltage applied to the anode electrode 160 is increased. When the output voltage of the DC power supply 50 rises to about 300V to 400V, ignition of the main discharge in the discharge space 190 of the TAL 100 is confirmed as shown in subsequent S18. When the discharge is normally ignited, an ion beam starts to be emitted from the output end 140 of the discharge space 190, and the state shifts to a steady operation state. The ring-like brightness of the discharge space 190 that disappeared in the previous S14 is observed again in S18. The ring-like brightness observed at this time is a brightness due to normal plasma discharge, and is a dim brightness as shown in FIG. Instead, the brightness is extremely high as shown in FIG.

  FIG. 14A shows the oscillation waveform of the discharge current in the TAL 100 in the steady operation state as described above, and FIG. 14B shows the discharge when a hollow cathode is used instead of the RF cathode 200 as an electron source. The vibration waveform of an electric current is shown. Comparing these results, it is clear that when the RF cathode 200 is used, the oscillation amplitude of the discharge current showing the discharge oscillation can be suppressed to about 1/3 to 1/4 compared with the case of using the hollow cathode. It was.

  FIG. 15 shows a stable operation region in which the discharge of the TAL 100 can be stably performed without changing when the propellant flow rate, the anode voltage, and the magnetic field strength, which are the operation parameters of the TAL 100, are changed. Indicates the magnetic field intensity, and the vertical axis indicates the product of the anode voltage and the propellant flow rate. In this graph, the region surrounded by the dotted line was verified, and the RF cathode 200 operated stably in the entire region. On the other hand, the region surrounded by the solid line is the stable operation region when the hollow cathode is used, and as a result, the stable operation region is sufficiently expanded when the RF cathode 200 is used compared to the hollow cathode. I understand that

  In Non-Patent Document 1, a Hall thruster using a hollow cathode as an electron source is disclosed, and comparison results with various operating parameters disclosed in Non-Patent Document 1 are shown in FIGS. In any result, the Hall thruster of this embodiment using the RF cathode is more stable in the operating parameter region where the operating conditions are stricter than the Hall thruster of Non-Patent Document 4 using a hollow cathode as the electron source. Can be seen to work. In the graph of FIG. 16, the horizontal axis indicates the magnetic field strength, the vertical axis indicates the product value of the anode voltage and the propellant flow rate, and “TMU-HET” indicates the electric propulsion device 1 shown in the embodiment. “UT-58” is data of the electric propulsion apparatus shown in Non-Patent Document 4.

In the embodiment described above, the RF cathode 200 is used as a stationary electron source of the Hall thruster. However, the RF cathode 200 can also be used as an ignition device for the main discharge of the Hall thruster. In this case, the propellant gas is supplied into the discharge space of the Hall thruster, the anode voltage is not applied, and no Hall magnetic field is formed or a very weak Hall magnetic field is formed. When the RF cathode is operated in this state, capacitively coupled plasma that emits a faint radiance shown in FIG. 10 is formed in the discharge space of the Hall thruster.
With such a usage method, the RF cathode can be used for plasma ignition of the main discharge. Then, after the main discharge is normally ignited, the RF cathode is turned off to stop all functions and reduce power consumption. By operating in this way, the main discharge can be ignited in a very short time compared to the case where a hollow cathode is used.

  In addition to the RF cathode exemplified in the embodiment, a hollow cathode may be provided and both of them may be used as an electron source. As an example of use in this case, an RF cathode that can be operated immediately by power supply is used at the initial stage of ignition of the main discharge, and at the timing when electrons are emitted from the hollow cathode that takes a long time to rise after shifting to a steady operation state. The RF cathode is stopped and the hollow cathode is used continuously thereafter. By using in this way, an RF cathode with high power consumption can be used only at the initial stage of main discharge ignition, and the main discharge can be ignited immediately, and then the operation is switched to the hollow cathode to reduce power consumption. it can.

  Further, a plurality of Hall thrusters are arranged around the one RF cathode exemplified in the embodiment so as to surround the periphery, and the RF cathode and each Hall thruster are installed so as to be electrically capacitively coupled. It is also possible to use the RF cathode for main discharge ignition of each Hall thruster arranged around. As an example of use in this case, a propellant gas is supplied into the discharge space of each Hall thruster, an anode voltage is not applied, and a Hall magnetic field is not formed or a very weak Hall magnetic field is formed. In this state, the central RF cathode is operated. In the discharge space of each hole thruster, plasma emitting a faint radiance as shown in FIG. 10 is formed. The so-formed plasma is used as a seed fire for main discharge ignition. Thereafter, an anode voltage is first applied to the anode electrode to form an electric field, and then a magnetic field is formed by a magnetic circuit. As a result, plasma serving as a main discharge can be formed in the discharge space. As a procedure for terminating the main discharge of the Hall thruster, the magnetic field may be turned off first after the electric field is turned off.

Further, as shown in FIG. 21, in addition to the RF cathode 200 exemplified in the embodiment as an electron source of the TAL 100, as another electron source, a hollow cathode 300 that emits electrons by heating the cathode is disposed. Both can also be used as electron sources. As disclosed in Patent Document 2 and the like, the hollow cathode 300 has been conventionally used as an electron source, and is a direct current discharge type in which the cathode is heated by direct current discharge to emit thermoelectrons. As an example of use in this case, an RF cathode 200 that can be operated immediately by power supply is used at the initial stage of main discharge ignition, and after shifting to a steady operation state, electrons are emitted from the hollow cathode 300 that takes time to rise. At the timing, the RF cathode 200 is stopped, and the hollow cathode 300 is continuously used thereafter. By using in this way, the RF cathode 200 with high power consumption can be used only at the initial stage of the main discharge ignition, and the main discharge can be immediately ignited, and then the operation is switched to the hollow cathode 300 to reduce the power consumption. be able to.
Furthermore, as shown in FIG. 22, a plurality of hole thrusters 100 are arranged so as to surround the periphery of one RF cathode 200 exemplified in the embodiment, and the RF cathode 200 and each Hall thruster 100 are electrically connected. It is also possible to install so as to be capacitively coupled, and one RF cathode 200 can be used for the main discharge ignition of each Hall thruster 100 disposed around. As an example of use in this case, a propellant gas is supplied into the discharge space of each Hall thruster 100, an anode voltage is not applied, a Hall magnetic field is not formed, or a very weak Hall magnetic field is formed. In this state, the central RF cathode 200 is operated. In the discharge space of each hole thruster 100, the plasma that emits the faint radiance shown in FIG. 10 is formed. The so-formed plasma is used as a seed fire for main discharge ignition. Thereafter, an anode voltage is first applied to the anode electrode to form an electric field, and then a magnetic field is formed by a magnetic circuit. As a result, plasma serving as a main discharge can be formed in the discharge space. As a procedure for terminating the main discharge of the Hall thruster 100, the magnetic field may be turned off first after the electric field is turned off.

As another embodiment, an example in which a redundant configuration of devices is assembled will be described. For devices that cannot be repaired and repaired easily after being launched, such as artificial satellites and spacecraft, for important parts that are essential to the operation of the device, in the unlikely event that the component breaks down, A redundant configuration is used in which several parts are prepared and switched to a spare part when a failure occurs. For example, in the electronic propulsion apparatus 1 exemplified in the embodiment, it is conceivable to provide a plurality of Hall thrusters 100 and a plurality of RF cathodes 200.
For example, referring to FIG. 22 shown above, six Hall thrusters 100 are provided, and one RF cathode 200 is provided in the center. Similar to the configuration described above, by operating the central RF cathode 200, it is possible to stably ignite the surrounding six Hall thrusters 100 simultaneously. Furthermore, when six Hall thrusters 100 are prepared as a redundant system, for example, when only two of them are to be operated, only two of them are operated by applying a voltage to the two anode electrodes. It is possible that the other Hall thrusters 100 do not ignite.
Similarly, in FIG. 23, two RF cathodes 200a and 200b are provided for one Hall thruster 100. That is, this is a case where a redundant system is assembled for the RF cathode 200. First, for comparison, consider a case where a redundant system is formed by a conventional DC discharge type hollow cathode 300 with reference to FIG. That is, this is a case where the hollow cathode 300 a and the hollow cathode 300 b are provided for one hall thruster 100. Of these, it is assumed that the cathode of the hollow cathode 300a is heated and operated by applying a voltage, and the hollow cathode 300b is not operated. When the Hall thruster 100 is ignited in this state, electrons flow into the anode electrode of the Hall thruster 100, so that electrons are extracted from the hollow cathode 300a. At this time, the hollow cathode 300 b is not heated, but the potential is the same as the cathode potential, that is, the same potential as the hollow cathode 300 a, and has a potential difference with respect to the anode electrode of the Hall thruster 100. Due to this potential difference, a part of the electrons to be output from the hollow cathode 300a may start to be output from the hollow cathode 300b. That is, electrons flow from the cathode of the hollow cathode 300b, and the cathode is heated by the current. This heating makes it easier for the cathode of the hollow cathode 300b to emit electrons, and more electrons are emitted. In other words, even if a plurality of hollow cathodes 300 are prepared for the purpose of a redundant system and some of them are made to stand by, if the Hall thruster 100 is ignited and part of the hollow cathode 300 starts emitting electrons, The hollow cathode 300 may also start to emit electrons, and the cathode of the standby hollow cathode 300 may deteriorate over time. That is, a redundant system cannot be configured simply by arranging the hollow cathodes 300 in parallel.
On the other hand, in the case of the RF cathode 200, when the RF cathode 200a is operated and the RF cathode 200b is kept stopped, the RF cathode 200a is activated by applying high-frequency power to ignite the Hall thruster 100, Although the operation becomes possible, there is no influence on the RF cathode 200b not operating at this time. That is, since the spare RF cathode 200b is not affected by the operation of the Hall thruster 100, there is no fear of deterioration over time. This is a very important feature when forming a redundant system.
When a plurality of RF cathodes 200 are used, a high frequency magnetic field generated by a certain RF cathode 200 may affect other RF cathodes 200. Therefore, particularly when a plurality of RF cathodes 200 are used as described above, if the discharge vessel 210 (FIG. 1) serving as the outer shell of the RF cathode 200 is made of a magnetic material such as iron, a high-frequency magnetic field is generated from the RF cathode 200. There is no leakage to the outside, and the influence on other RF cathodes 200 is small. Of course, the influence of electromagnetic noise on other equipment can also be reduced.
In this manner, the discharge vessel 210 and the induction coil 220 can be made of a magnetic material, and thereby the above-described effects can be achieved.

  Further, the above examples can be used in combination. For example, as shown in FIG. 25, an RF cathode 200 and a hollow cathode 300 are provided for two Hall thrusters 100, and the RF cathode 200 is electrically capacitively coupled to the two Hall thrusters 100. Install as follows. With this arrangement, one RF cathode 200 is used for main discharge ignition of each Hall thruster 100, or the RF cathode 200 is used only at the initial stage of main discharge ignition, and then switched to the hollow cathode 300. There are various possibilities such as operating one by one, or providing one as the other redundant system.

In each of the examples described above, the anode layer type hall thruster is exemplified as the hole thruster. However, the anode electrode is disposed in the deep part of the discharge space, and the acceleration channel has a channel length longer than the channel width. It is also possible to employ a linear Hall thruster called a “Stationary Plasma Thruster” type. That is, even when a linear Hall thruster is used, the anode electrode and the inductively coupled plasma electron source in the Hall thruster are arranged so as to be electrically capacitively coupled to each other, so that the same as in the present embodiment. It is thought that the effective effect is demonstrated.

Claims (5)

An electric propulsion device that introduces a propellant from one axial direction of an annular discharge space, accelerates ions generated from the propellant by an electric field, and discharges it from the other output end of the discharge space to obtain thrust There,
A hall thruster having a magnetic circuit for forming a magnetic field in a radial direction of the discharge space and having an anode electrode in the discharge space that contributes to the formation of the electric field;
A cylindrical discharge vessel disposed near the output end of the hall thruster and supplied with a working gas, and an induction coil disposed along the discharge vessel, for supplying high-frequency power to the induction coil The plasma of the working gas in the discharge space in the discharge vessel, and an inductively coupled plasma electron source that emits the generated electrons.
An electric propulsion device, wherein the Hall thruster and the inductively coupled plasma electron source are arranged so that the anode electrode of the Hall thruster and the inductively coupled plasma electron source are electrically capacitively coupled. The inductively coupled plasma electron source is
The electric propulsion device according to claim 1, wherein the electric propulsion device is formed by an inner coil type RF cathode in which the induction coil is disposed in the discharge vessel.   The electric propulsion apparatus according to claim 1, wherein the hall thruster is an anode layer type hall thruster in which the anode electrode is disposed in the vicinity of the output end.   2. The electric propulsion apparatus according to claim 1, further comprising a direct current discharge type electron source that emits thermoelectrons by heating the cathode by direct current discharge. 2. A plurality of the Hall thrusters are provided, and one or more inductively coupled plasma electron sources are arranged to be electrically capacitively coupled to the plurality of Hall thrusters. Electric propulsion device.