WO2022249428A1 - 測距機能を備える衛星減速装置 - Google Patents

測距機能を備える衛星減速装置 Download PDF

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Publication number
WO2022249428A1
WO2022249428A1 PCT/JP2021/020319 JP2021020319W WO2022249428A1 WO 2022249428 A1 WO2022249428 A1 WO 2022249428A1 JP 2021020319 W JP2021020319 W JP 2021020319W WO 2022249428 A1 WO2022249428 A1 WO 2022249428A1
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Prior art keywords
satellite
light
light output
signal
output
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English (en)
French (fr)
Japanese (ja)
Inventor
貴敬 鈴木
俊行 安藤
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Mitsubishi Electric Corp
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Mitsubishi Electric Corp
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Priority to PCT/JP2021/020319 priority Critical patent/WO2022249428A1/ja
Priority to JP2021556702A priority patent/JPWO2022249428A1/ja
Publication of WO2022249428A1 publication Critical patent/WO2022249428A1/ja
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G3/00—Observing or tracking cosmonautic vehicles

Definitions

  • the technology disclosed herein relates to a satellite deceleration device having a ranging function.
  • a satellite deceleration device is a device that decelerates a target satellite by applying power from afar. The purpose of slowing down the satellite is to plunge it into the atmosphere and burn it up.
  • Patent Document 1 As a conventional technique for annihilating space debris in the earth's atmosphere, there is known a space debris annihilation apparatus that includes an electron beam generator and an electric field generator (for example, Patent Document 1).
  • a space debris incineration apparatus illustrated in Patent Document 1 is used by being mounted on an artificial satellite.
  • a space debris incinerator irradiates an electron beam toward space debris. Space debris receives an electron beam and becomes negatively charged.
  • the electric field generator of the space debris annihilator generates a negative electric field. Space debris receives electrostatic force from a negative electric field, changes its trajectory due to electrostatic repulsion, and finally enters the atmosphere and burns up.
  • the space debris destruction apparatus is a timing calculation for calculating the beam irradiation timing for irradiating an electron beam and the electric field generation timing for generating an electric field based on space debris orbit information and artificial satellite orbit information. has a department. Further, the space debris extermination apparatus according to Patent Document 1 calculates the beam irradiation direction of the electron beam based on the orbital information of the space debris and the orbital information of the artificial satellite, and changes the direction of the electrostatic force of the electric field to the direction of the electric field. It has a direction calculation unit that calculates as
  • the space debris destruction apparatus illustrated in Patent Document 1 is based on the premise that the trajectory of space debris can be accurately obtained as described above.
  • Translating space debris into satellites means that satellite decelerators need to accurately obtain the orbits of the satellites.
  • Satellite orbital information is generally described in a format called Two Line Elements (TLE). Satellite orbit information increases in error over time if it is not updated. Also, satellite orbit information is not accurate without frequent updates, especially for erratically moving satellites. That is, the conventional technology has a problem that it is difficult to decelerate a satellite whose orbital information is not accurate.
  • a satellite deceleration device having a ranging function includes a pulse light source, a second light output unit that outputs light output from the pulse light source to space, and a target satellite that is output from the second light output unit. a second optical detector that converts the reflected light reflected by the second optical detector into an electrical signal; and the distance to the target satellite from the time difference between the trigger signal from the pulse light source and the electrical signal from the second optical detector. and a first signal processor for calculating.
  • the satellite deceleration device with the ranging function according to the technology disclosed herein has the above configuration, it is possible to confirm the orbit change effect of beam ablation based on actual measurement values. As a result, the satellite deceleration device having the ranging function according to the technology disclosed herein can reliably decelerate even a satellite whose orbital information is not accurate.
  • FIG. 1 is a configuration diagram showing a configuration example of a satellite deceleration device having a ranging function according to Embodiment 1.
  • FIG. FIG. 2 is a configuration diagram showing a configuration example of the phase error detection section 5 of the satellite deceleration device having the ranging function according to the first embodiment.
  • FIG. 3 is a schematic diagram of the light output unit 6 according to the first embodiment.
  • FIG. 4A is a schematic diagram of a satellite deceleration device having a ranging function according to Embodiment 1 capturing a target satellite by spiral search.
  • FIG. 4B is a flow chart showing an operation example of the second signal processing section 9 according to the first embodiment.
  • FIG. 5A is a schematic diagram showing the irradiation operation of the satellite deceleration device having the ranging function according to Embodiment 1.
  • FIG. 5B is a first flowchart showing an operation example of the first signal processing section 8 according to the first embodiment.
  • FIG. 6 is a flow chart 2 showing an operation example of the first signal processing section 8 according to the first embodiment.
  • FIG. 7 is a flow chart 3 showing an operation example of the first signal processing section 8 according to the first embodiment.
  • target satellite a target satellite whose distance is to be measured
  • space debris those that are slowed down and burned up in the atmosphere are called “space debris” or simply “debris.”
  • FIG. 1 is a configuration diagram showing a configuration example of a satellite deceleration device having a ranging function according to Embodiment 1.
  • the satellite deceleration device having the ranging function according to the first embodiment includes a reference light source section 1, a signal distribution section 2, an element unit section 3, an optical amplification section 4, and a phase error detection section 5. , an optical output unit 6 , an optical receiving unit 7 , a first signal processing section 8 , a second signal processing section 9 , and a pulse light source 39 .
  • the first signal processing unit 8 may be connected to the information output interface 12, for example.
  • the second signal processing unit 9 may be connected to the satellite orbit input interface 10 and the gimbal 11, for example.
  • the signal distribution unit 2 includes an optical demultiplexer 21 and a signal distribution coupler 22 .
  • the element unit section 3 according to Embodiment 1 includes a phase modulator 23, an optical amplifier 24, and a beam splitter 25.
  • phase error detection unit 5 will become clear from the explanation along with FIG. 2 below.
  • the light output unit 6 includes a first light output section 26 and a second light output section 27 .
  • the optical receiver unit 7 includes an optical receiver 28, an optical demultiplexer 29, a first photodetector 30, and a second photodetector 31.
  • the reference light source unit 1 outputs light for irradiating the satellite and causing plasma ablation.
  • the output light may be a single wavelength reference light.
  • the reference light source unit 1 may be, for example, a laser light source.
  • the optical demultiplexer 21 of the signal distribution unit 2 splits the light output from the reference light source unit 1 into signal light and local oscillation light (hereinafter referred to as "local light").
  • the optical demultiplexer 21 may be, for example, a 1 ⁇ 2 polarization-maintaining filter coupler.
  • the signal distribution coupler 22 distributes the signal light branched by the optical demultiplexer 21 into a plurality of signal lights.
  • the number of signal distributions of the signal distribution coupler 22 should be the same as the number of the element unit sections 3 .
  • the distributed signal lights are output to different element unit sections 3, respectively.
  • the signal distribution unit 2 has an optical demultiplexer 21 and a signal distribution coupler 22 .
  • the signal distribution unit 2 separates the reference light output from the reference light source unit 1 into a plurality of signal lights and a single local light.
  • phase modulator 23 phase-modulates the light output from the signal distribution coupler 22 and input to the element unit section 3 .
  • Phase modulator 23 may be, for example, an LN modulator or an AO modulator.
  • the optical amplifier 24 amplifies the light phase-modulated by the phase modulator 23 .
  • Optical amplifier 24 may be, for example, a semiconductor optical amplifier (SOA).
  • SOA semiconductor optical amplifier
  • the beam splitter 25 splits the input light into two. One of the two split lights is output to the phase error detector 5 . The other of the two-branched light is output to the light output unit 6 .
  • the optical amplifier 4 amplifies local light output from the optical demultiplexer 21 .
  • the optical amplifier section 4 may be, for example, an erbium-doped optical fiber amplifier (EDFA).
  • EDFA erbium-doped optical fiber amplifier
  • FIG. 2 is a configuration diagram showing a configuration example of the phase error detection unit 5 of the satellite deceleration device having the ranging function according to the first embodiment.
  • the phase error detector 5 includes a coupler 32, a photodetector 33, a reference signal generator 34, a phase shifter 35, a phase comparator 36, a loop filter 37, and a VCO 38. include.
  • Coupler 32 multiplexes the light output from the element unit section 3 and the light output from the optical amplification section 4 .
  • Coupler 32 may be, for example, a 1 ⁇ 2 polarization-maintaining filter coupler.
  • the photodetector 33 converts the light output from the coupler 32 into an electrical signal. Specifically, the photodetector 33 performs heterodyne detection on the light output from the element unit section 3 and the light output from the optical amplification section 4 to detect a beat signal. Photodetector 33 may be, for example, a photodiode.
  • the reference signal generation source 34 generates an electrical signal.
  • the electrical signal generated by reference signal source 34 is output to phase shifter 35 as a reference signal.
  • Reference signal source 34 may be, for example, a signal generator.
  • the phase shifter 35 shifts the phase of the electrical signal generated by the reference signal source 34 .
  • the phase shifter 35 also performs phase adjustment according to the control signal output from the second signal processing section 9 .
  • the phase comparator 36 compares the electrical signal whose phase has been adjusted by the phase shifter 35 and the beat signal detected by the photodetector 33 to detect the phase difference.
  • the phase difference detected here is referred to as a phase error signal for convenience.
  • a loop filter 37 smoothes the phase error signal detected by the phase comparator 36 .
  • the VCO 38 generates a signal for matching the frequency of the beat signal with the frequency of the electrical signal whose phase is adjusted by the phase shifter 35 according to the phase error signal smoothed by the loop filter 37 .
  • the signal generated here is output to the phase modulator 23 of the element unit section 3 .
  • FIG. 3 is a schematic diagram of the light output unit 6 according to Embodiment 1.
  • the light output unit 6 has a plurality of light output ends arranged in an array and outputs light toward the target satellite.
  • the light output unit 6 has a plurality of first light output sections 26 and at least one second light output section 27 .
  • the number of first light output sections 26 is equal to the number of element unit sections 3 .
  • the first light output section 26 outputs the light output from the element unit section 3 to space.
  • the light output to the space by the first light output section 26 is the light output from the phase error detection section 5 to the element unit section 3 .
  • the second light output unit 27 outputs the light output from the pulse light source 39 to space.
  • the pulsed light source 39 outputs pulsed light with a wavelength different from that of the light output from the reference light source section 1 .
  • the light receiving unit 28 receives reflected light reflected by the target satellite.
  • the received reflected light is output to the optical demultiplexing section 29 .
  • the optical receiver 28 may be, for example, a large telescope.
  • the optical demultiplexing unit 29 separates the optical path on the output side so as to change the output destination according to the wavelength of the input light.
  • the reflected light output from the first light output section 26 and reflected by the target satellite is output to the first light detection section 30 .
  • the reflected light output from the second light output section 27 and reflected by the target satellite is output to the second light detection section 31 .
  • the optical splitter 29 may be, for example, a dichroic mirror.
  • the first photodetector 30 converts the reflected light output from the optical demultiplexer 29 into an electrical signal. This electrical signal is output to the second signal processing section 9 as a signal for initial acquisition.
  • the first photodetector 30 may be, for example, a CCD image sensor.
  • the second photodetector 31 converts the reflected light output from the optical demultiplexer 29 into an electrical signal. This electrical signal is output to the first signal processing section 8 as a signal for satellite ranging.
  • the second photodetector 31 may be, for example, a photon counting detector.
  • the first signal processing unit 8 calculates the distance to the target satellite from the time difference between the trigger signal from the pulse light source 39 and the electrical signal from the second light detection unit 31.
  • the calculated distance information about the target satellite is output to the satellite orbit input interface 10 as real-time satellite orbit information.
  • the first signal processing unit 8 also generates a control signal for controlling the orbital position of its own station based on the calculated distance information.
  • the control signal generated here is output to the information output interface 12 .
  • the second signal processing unit 9 uses the satellite position information obtained from the first signal processing unit 8 and the second light detection unit 31 and the public satellite orbit information output from the satellite orbit input interface 10. , to generate beam direction control signals.
  • the generated irradiation direction control signal is output to the gimbal 11 and the phase shifter 35, respectively.
  • the irradiation direction control signal output to the gimbal 11 is used for coarse tracking. Also, the irradiation direction control signal output to the phase shifter 35 is used for fine tracking.
  • the satellite orbit input interface 10 receives public satellite orbit information from the outside. As described above, public satellite orbit information is expressed in, for example, the TLE format. Public satellite orbit information is output to the second signal processing unit 9 .
  • the gimbal 11 controls the azimuth and elevation of the telescope on which the light output unit 6 is mounted based on the control signal from the second signal processing section 9, and scans the output light.
  • azimuth and elevation are one of methods of expressing the position of the satellite as seen from the observer.
  • Azimuth is the azimuth angle of the satellite.
  • Elevation is the elevation angle of the satellite.
  • the information output interface 12 outputs a control signal to the drive mechanism of the own station based on the control signal for controlling the track position of the own station output from the first signal processing unit 8 .
  • the driving mechanism of the local station may be, for example, an engine.
  • FIG. 4A is a schematic diagram of a satellite deceleration device having a ranging function according to Embodiment 1 capturing a target satellite by spiral search.
  • FIG. 4B is a flow chart showing an operation example of the second signal processing section 9 according to the first embodiment.
  • FIG. 5A is a schematic diagram showing the irradiation operation of the satellite deceleration device having the ranging function according to Embodiment 1.
  • FIG. 5B is a flowchart 1 showing an operation example of the first signal processing section 8 according to the first embodiment.
  • the satellite deceleration device having the ranging function according to Embodiment 1 performs initial acquisition of the target satellite by, for example, spiral scanning.
  • Spiral scanning can be achieved by scanning the beams by controlling the gimbal 11 and controlling the relative phases of each beam.
  • the range over which the beam is scanned may be an indefinite range. If the target satellite exists within an indefinite range, the irradiated beam is reflected by the target satellite, and the reflected light arrives.
  • the first photodetector 30 acquires the reflected light. As described above, the first photodetector 30 converts the reflected light output from the optical demultiplexer 29 into an electrical signal. The converted electrical signal is output to the second signal processing section 9 as the position information of the target satellite.
  • the processing steps of the second signal processing unit 9 include a step of acquiring the position information of the target satellite sent from the first photodetector unit 30 (ST41), a step of comparing the satellite orbit information from 10 (ST42); a step of determining whether or not pointing correction is necessary (ST43); and a step of controlling the phase shifter 35 and the gimbal 11 (ST44). have.
  • the second signal processing section 9 When it is determined in ST43 that pointing correction is unnecessary, the second signal processing section 9 continues to control the gimbal 11 based on public satellite orbit information. When it is determined in ST43 that pointing correction is necessary, the second signal processing section 9 outputs control signals for the phase shifter 35 and the gimbal 11 to perform initial acquisition and tracking.
  • the satellite deceleration device having the ranging function according to Embodiment 1 is capable of phase control of each beam in addition to mechanical control of the gimbal 11 as described above. Phase control of each beam enables directivity angle control by electronic scanning, and has the effect of realizing high-speed scanning.
  • the satellite deceleration device having the ranging function outputs a ranging pulse laser and a satellite deceleration beam from the optical output unit 6 .
  • the wavelength of the pulsed laser for distance measurement is ⁇ 0 .
  • the wavelength of the satellite deceleration beam is ⁇ 1 . That is, the pulse laser for range finding and the beam for satellite deceleration have different wavelengths.
  • the satellite deceleration beam is directed toward the debris and decelerates the debris by plasma ablation.
  • the pulsed laser for ranging is reflected by the target satellite and received by the optical receiving unit 7 .
  • the optical receiving unit 7 may be co-located with the optical output unit 6 .
  • the first signal processing unit 8 calculates the distance to the satellite from the time difference between the trigger signal of the pulse laser and the arrival of reflected light from the satellite.
  • the technique of measuring the distance to an object using the time of flight of light in this way is called the Time of Flight technique (hereinafter "ToF technique").
  • the processing steps of the first signal processing unit 8 include a step of acquiring a pulse laser trigger signal (ST51), a step of acquiring reflected light from the target satellite (ST52), and a target It has a step of calculating the distance to the satellite (ST53) and a step of outputting the position information (ST54).
  • the satellite deceleration device with the ranging function according to Embodiment 1 Since the satellite deceleration device with the ranging function according to Embodiment 1 has the above configuration, it has the effect of continuing to acquire orbit information even after the target satellite changes its orbit.
  • the satellite deceleration device having the ranging function according to Embodiment 1 may have the optical output unit 6 placed on the ground, or may have a station provided on an orbiting satellite and placed there.
  • the satellite deceleration device having the ranging function according to the first embodiment can approach debris based on the ranging result and efficiently perform ablation if a station is provided on an orbiting satellite.
  • FIG. 6 is a second flowchart showing an operation example of the first signal processing section 8 according to the first embodiment.
  • Formula (1) shown below expresses the conditions for developing ablation.
  • Ep is the pulse energy
  • ⁇ is the wavelength
  • M2 is the beam quality
  • D is the output beam diameter
  • L is the distance from the satellite
  • Fth is the ablation threshold.
  • Orbit change by plasma ablation must satisfy the conditional expression (1) that the beam irradiated to the satellite causes ablation.
  • pulse energy (E p ), wavelength ( ⁇ ), beam quality (M 2 ), and output beam diameter (D) are determined by component specifications. Therefore, the distance (L) to the satellite becomes a variable parameter for the occurrence of ablation.
  • a satellite deceleration device having a ranging function according to the technology disclosed herein calculates the distance (L) to the satellite using the ranging function, and checks whether or not the ablation occurrence condition is satisfied.
  • the satellite deceleration device having the range finding function according to the technology of the present disclosure may be programmed to start beam irradiation (when "none" is set in step ST61 in FIG. selected).
  • the satellite deceleration device with the ranging function according to the technology of the present disclosure may be programmed to approach the target satellite (if "yes" in step ST61 in FIG. 6 is selected).
  • the approach to the target satellite may be realized, for example, by sending a control signal from the first signal processing section 8 to the information output interface 12 and operating a driving mechanism such as an engine based on the control signal (ST62 in FIG. 6). steps shown).
  • FIG. 7 is a third flowchart showing an operation example of the first signal processing section 8 according to the first embodiment.
  • FIG. 7 shows the action of the first signal processing section 8 when verifying the trajectory change effect due to ablation in particular.
  • Equation (2) expresses the relationship between the velocity of the satellite and the orbit radius.
  • ⁇ r represents the change in the orbit radius
  • ⁇ v represents the change in the velocity of the satellite.
  • the unit of ⁇ r is [km]
  • the unit of ⁇ v is [m/s].
  • a satellite deceleration device having a ranging function according to the technology disclosed herein inputs orbit information after deceleration obtained by ranging to the first signal processing unit 8, and is given by the input measured value and equation (2). It may be programmed to compare with the calculated value (step indicated by ST71 in FIG. 7). Further, the satellite deceleration device having the ranging function according to the technology disclosed herein may be programmed to calculate error information from the actual measurement value and the calculated value compared in ST71 (step indicated by ST72 in FIG. 7). By being programmed in this manner, the satellite deceleration device having the ranging function according to the technology of the present disclosure can confirm the orbit change effect of beam ablation based on actual measurements.
  • the satellite deceleration device with the ranging function according to Embodiment 1 has the above configuration, it is possible to confirm the orbit change effect of beam ablation based on actual measurement values. As a result, the satellite deceleration device having the ranging function according to the first embodiment can reliably decelerate even a satellite whose orbital information is not accurate.
  • the disclosed technology can be applied to space debris extermination equipment and has industrial applicability.
  • 1 reference light source section 1 reference light source section, 2 signal distribution section, 3 element unit section, 4 optical amplification section, 5 phase error detection section, 6 optical output unit, 7 optical reception unit, 8 first signal processing section, 9 second signal processing section, 10 satellite orbit input interface, 11 gimbal, 12 information output interface, 21 optical splitter, 22 signal distribution coupler, 23 phase modulator, 24 optical amplifier, 25 beam splitter, 26 first optical output unit, 27 second optical Output section, 28 optical receiver, 29 optical demultiplexer, 30 first photodetector, 31 second photodetector, 32 coupler, 33 photodetector, 34 reference signal source, 35 phase shifter, 36 phase comparator device, 37 loop filter, 38 VCO, 39 pulse light source.

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PCT/JP2021/020319 2021-05-28 2021-05-28 測距機能を備える衛星減速装置 Ceased WO2022249428A1 (ja)

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WO2020152744A1 (ja) * 2019-01-21 2020-07-30 スカパーＪｓａｔ株式会社 宇宙機、制御システム

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DE3943374A1 (de) * 1989-12-30 1991-07-04 Deutsche Forsch Luft Raumfahrt Verfahren und einrichtung zum entfernen von weltraumtruemmern
JPH09318743A (ja) * 1996-05-30 1997-12-12 Toshiba Corp 距離測定装置
JP2010127818A (ja) * 2008-11-28 2010-06-10 Mitsubishi Electric Corp レーザー照射装置
RU2505461C1 (ru) * 2009-11-25 2014-01-27 Поулос Эйр Энд Спейс Стабилизация движения неустойчивых фрагментов космического мусора
JP6233606B2 (ja) * 2015-08-04 2017-11-22 日本電気株式会社 目標識別レーザ観測システム
US20170067996A1 (en) * 2015-09-04 2017-03-09 U.S.A. As Represented By The Administrator Of The National Aeronautics And Space Administration Ground-based laser ranging system for identification and tracking of orbital debris

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