US20160376032A1 - Torque generation system, attitude control system for spacecraft, and relative position and velocity control system for spacecraft - Google Patents

Torque generation system, attitude control system for spacecraft, and relative position and velocity control system for spacecraft Download PDF

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Publication number
US20160376032A1
US20160376032A1 US15/150,994 US201615150994A US2016376032A1 US 20160376032 A1 US20160376032 A1 US 20160376032A1 US 201615150994 A US201615150994 A US 201615150994A US 2016376032 A1 US2016376032 A1 US 2016376032A1
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Prior art keywords
solar array
spacecraft
attitude
light receiving
array panels
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Osamu Mori
Junichiro Kawaguchi
Yoji SHIRASAWA
Toshihiro CHUJO
Kosuke AKATSUKA
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Japan Aerospace Exploration Agency JAXA
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Japan Aerospace Exploration Agency JAXA
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Assigned to JAPAN AEROSPACE EXPLORATION AGENCY reassignment JAPAN AEROSPACE EXPLORATION AGENCY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: AKATSUKA, KOSUKE, CHUJO, TOSHIHIRO, KAWAGUCHI, JUNICHIRO, MORI, OSAMU, SHIRASAWA, YOJI
Publication of US20160376032A1 publication Critical patent/US20160376032A1/en
Priority to US15/993,302 priority Critical patent/US10577132B2/en
Abandoned legal-status Critical Current

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/244—Spacecraft control systems
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/34—Guiding or controlling apparatus, e.g. for attitude control using gravity gradient
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
    • B64G1/361—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using star sensors
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
    • B64G1/365—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using horizon or Earth sensors
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40—Arrangements or adaptations of propulsion systems
    • B64G1/407—Solar sailing
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/42—Arrangements or adaptations of power supply systems
    • B64G1/44—Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/42—Arrangements or adaptations of power supply systems
    • B64G1/44—Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
    • B64G1/443—Photovoltaic cell arrays
    • H—ELECTRICITY
    • H02—GENERATION; CONVERSION OR DISTRIBUTION OF ELECTRIC POWER
    • H02S—GENERATION OF ELECTRIC POWER BY CONVERSION OF INFRARED RADIATION, VISIBLE LIGHT OR ULTRAVIOLET LIGHT, e.g. USING PHOTOVOLTAIC [PV] MODULES
    • H02S40/00—Components or accessories in combination with PV modules, not provided for in groups H02S10/00 - H02S30/00
    • H02S40/30—Electrical components
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02E—REDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
    • Y02E10/00—Energy generation through renewable energy sources
    • Y02E10/50—Photovoltaic [PV] energy

Definitions

  • the present invention relates to torque generation systems, attitude control systems for a spacecraft, and relative position and velocity control systems for a spacecraft.
  • the present invention more particularly relates to a torque generation system, an attitude control system for a spacecraft, and a relative position and velocity control system for a spacecraft, which use a thermal radiation pressure generated from a solar array panel.
  • FIG. 1 is an explanatory view of various kinds of disturbance torque based on computation examples of disturbance torque with respect to the altitude of earth orbiting satellites, which is illustrated in “Recent Trends in Three-Axis Stabilized Satellites” by Susumu Nishida in the Journal of the Japan Society for Aeronautical and Space Sciences, Vol. 22, No. 243, 1974, p. 200.
  • aerodynamics torque is a main disturbance in the very low-altitude orbit in an altitude range of about 100 to 200 km
  • gravity gradient torque is the main disturbance in an orbit in an altitude range up to about 10,000 km
  • radiation pressure torque is the main disturbance in a higher altitude orbit.
  • thrusters and reaction wheels are conventionally used (Japanese Laid-Open Patent Publication No. 2000-296800 and Japanese Laid-open Patent Publication No. 2015-703).
  • a technology to control the attitude with a liquid crystal device that changes thrust caused by a solar radiation pressure is recently being developed (Y. Tsuda, Y. Mimasu, R. Funase, T. Saiki, Y. Shirasawa, O. Mori, N. Motooka, and T. Yamamoto, “Challenges and Results on Attitude Control Operation of World's First Solar Power Sail IKAROS,” Journal of Space Technology and Science, Vol. 27, No. 1, 2013, pp. 69-78).
  • the amount of fuel mountable on a spacecraft is limited, and the thruster is not operable once the fuel is exhausted.
  • the reaction wheel sufficiently suppresses the turbulence of its own making, there is a demand for an attitude control device that causes further lower turbulence for use in future earth observation and astronomical observation missions which challenge a diffraction limit in short wavelength bands, for example.
  • the Solar-B project which is a solar observation satellite project, requires a degree of attitude stabilization of 3.0 arcsec/10 sec and 2.0 arcsec/hour and the Solar-C project requires a degree of attitude stabilization even higher than that in the Solar-B project.
  • a typical satellite-mounted wheel generates turbulence as illustrated in Table 1.
  • the liquid crystal device when the aforementioned technology involving use of the liquid crystal device to change the thrust caused by a solar radiation pressure is used, it becomes possible to control the attitude of spacecraft and to control the relative position and relative velocity of the spacecraft without using a thruster. This leads to implementation of the spacecraft with a lighter weight and a longer life span. Furthermore, since the reaction wheel is not mounted on the spacecraft, the attitude control completely free from causing turbulence can be achieved. However, in order to utilize such a technology, the liquid crystal device needs to be mounted on the spacecraft as an additional component.
  • an object of the present invention is to provide a torque generator capable of generating very small torque by using a component which is also used as a component constituting a spacecraft.
  • Another object of the present invention is to provide an attitude control system for a spacecraft, and a relative position and velocity control system for a spacecraft with high precision and without causing turbulence, the systems being free from the necessity of mounting an additional device on the spacecraft.
  • One aspect of the present invention is to provide a torque generation system, including: a plurality of solar array panels and/or solar array panel divisions; and a torque controller configured to control an electricity generation ratio of each of the plurality of solar array panels and/or solar array panel divisions to generate torque.
  • the electricity generation ratio may be controlled by turning on, turning off, or switching a shunt switch provided for each of the plurality of solar array panels and/or solar array panel divisions.
  • the electricity generation ratio may be controlled by switching a maximum output state and a shunt state in each of solar cell groups whose regions are smaller than each of the plurality of solar array panels and/or solar array panel divisions.
  • the torque controller may put at least one of the plurality of solar array panels and/or solar array panel divisions in a shunt state.
  • the torque controller may put at least one of the plurality of solar array panels and/or solar array panel divisions in a maximum output state.
  • Each of the solar array panels may have a light receiving surface significantly higher in temperature than a surface opposite to the light receiving surface of each of the solar array panels during operation.
  • a member with small thermal conductivity may be arranged between the light receiving surface of each of the solar array panels and the surface opposite to the light receiving surface of each of the solar array panels.
  • the surface opposite to the light receiving surface may be smaller in thermal emissivity than the light receiving surface.
  • the surface opposite to the light receiving surface of each of the solar array panels may be coated with aluminum.
  • Another aspect of the present invention is to provide an attitude control system for a spacecraft, including: a plurality of solar array panels and/or solar array panel divisions; and an attitude controller configured to control an electricity generation ratio of each of the plurality of solar array panels and/or solar array panel divisions to generate torque used to control an attitude of the spacecraft.
  • the electricity generation ratio may be controlled by turning on, turning off, or switching a shunt switch provided for each of the plurality of solar array panels and/or solar array panel divisions.
  • the electricity generation ratio may be controlled by switching a maximum output state and a shunt state in each of solar cell groups whose regions are smaller than each of the plurality of solar array panels and/or solar array panel divisions.
  • the attitude control system may further includes: an attitude detector configured to detect the attitude of the spacecraft; and a target attitude setting unit configured to set a target attitude of the spacecraft, wherein the attitude controller may control the electricity generation ratio of each of the plurality of solar array panels and/or solar array panel divisions to generate torque that decreases a difference between a current attitude detected by the attitude detector and the target attitude.
  • the attitude may be an attitude angle, and the attitude controller may perform feedback control based on the attitude angle.
  • the target attitude may be an attitude whereby the spacecraft points to the sun.
  • the attitude control system may further include a disturbance torque estimation unit configured to estimate disturbance torque, wherein the attitude controller may control the electricity generation ratio of each of the plurality of solar array panels and/or solar array panel divisions to generate torque that suppresses the disturbance torque estimated by the disturbance torque estimation unit.
  • the plurality of solar array panels and/or solar array panel divisions may be arranged to generate torque for rotating the spacecraft around one predetermined axis.
  • the one predetermined axis may extend in a travelling direction of the spacecraft or in a center of the earth direction.
  • the plurality of solar array panels and/or solar array panel divisions may be arranged to generate torque for rotating the spacecraft around two predetermined axes, respectively.
  • the two predetermined axes may extend in a travelling direction of the spacecraft and in a center of the earth direction.
  • the plurality of solar array panels and/or solar array panel divisions may be symmetrically arranged with respect to a body of the spacecraft.
  • Each of the solar array panels may have a light receiving surface significantly higher in temperature than a surface opposite to the light receiving surface of each of the solar array panels during operation.
  • a member with small thermal conductivity may be arranged between the light receiving surface of each of the solar array panels and the surface opposite to the light receiving surface of each of the solar array panels.
  • the surface opposite to the light receiving surface may be smaller in thermal emissivity than the light receiving surface.
  • the surface opposite to the light receiving surface of each of the solar array panels may be coated with aluminum.
  • Another aspect of the present invention is to provide a spacecraft including the attitude control system.
  • Still another aspect of the present invention is to provide a system for controlling a relative position and/or a relative velocity of a first spacecraft and a second spacecraft, including: the first spacecraft and the second spacecraft each including a solar array panel; and a relative position and velocity controller configured to control an electricity generation ratio of each of the solar array panels of the first spacecraft and the second spacecraft to cause each of the first spacecraft and the second spacecraft to generate thrust that changes the relative position and/or velocity of the first spacecraft and the second spacecraft.
  • the system may further include a relative position and velocity setting unit configured to set a target relative position and/or a target relative velocity of the second spacecraft relative to the first spacecraft, wherein the relative position and velocity controller may control the electricity generation ratio of each of the solar array panels of the first spacecraft and the second spacecraft to cause each of the first spacecraft and the second spacecraft to generate thrust that provides the target relative position and/or target relative velocity set by the relative position and velocity setting unit.
  • a relative position and velocity setting unit configured to set a target relative position and/or a target relative velocity of the second spacecraft relative to the first spacecraft
  • the relative position and velocity controller may control the electricity generation ratio of each of the solar array panels of the first spacecraft and the second spacecraft to cause each of the first spacecraft and the second spacecraft to generate thrust that provides the target relative position and/or target relative velocity set by the relative position and velocity setting unit.
  • the first spacecraft and the second spacecraft each may include a position detector configured to detect a position of its own spacecraft and to output position information
  • the system may further includes a relative position and velocity calculation unit configured to calculate the relative position and/or relative velocity of the second spacecraft relative to the first spacecraft based on the position information from the position detectors of the first spacecraft and the second spacecraft
  • the relative position and velocity controller may control the electricity generation ratio of each of the solar array panels of the first spacecraft and the second spacecraft based on the relative position and/or relative velocity of the second spacecraft relative to the first spacecraft calculated by the position and velocity calculation unit.
  • the relative position and velocity controller may put the solar array panel of the second spacecraft in a shunt state.
  • the relative position and velocity controller may put the solar array panel of the first spacecraft in a maximum output state.
  • Each of the solar array panels may have a light receiving surface significantly higher in temperature than a surface opposite to the light receiving surface of each of the solar array panels during operation.
  • a member with small thermal conductivity may be arranged between the light receiving surface of each of the solar array panels and the surface opposite to the light receiving surface of each of the solar array panels.
  • the surface opposite to the light receiving surface may be smaller in thermal emissivity than the light receiving surface.
  • the surface opposite to the light receiving surface of each of the solar array panels may be coated with aluminum.
  • An attitude of the first spacecraft and/or the second spacecraft may be controlled to cause the solar array panels of the first spacecraft and/or the solar array panels of the second spacecraft to point to the sun.
  • a term “electricity generation ratio” refers to a ratio of output electricity of solar cell or cells to incoming solar energy W.
  • a term “solar array panel” refers to an electricity generator including a plurality of solar cells arrayed on at least part of the surface of a support.
  • the solar cells may be any appropriate solar cells, such as crystalline solar cells, amorphous solar cells, and thin-film solar cells.
  • the support may have rigidity or may have flexibility.
  • the plurality of solar cells may be arrayed on almost the entire surface of the support, or may be arrayed in one or more subregions on the surface of the support as disclosed in, for example, “Challenges and Results on Attitude Control Operation of World's First Solar Power Sail IKAROS.”
  • the torque generator according to the present invention can generate very small torque by using solar array panels that are also used as components constituting a spacecraft.
  • the attitude control system for a spacecraft and the relative position and velocity control system for a spacecraft can control the attitude of the spacecraft and control the relative position and velocity of the spacecraft with high precision and without causing turbulence, the attitude and the relative position and velocity being controlled by using the solar array panels mounted on the spacecraft without the necessity of mounting an additional device thereon.
  • FIG. 1 is an explanatory view of various kinds of disturbance torque with respect to the altitude of earth orbiting satellites
  • FIG. 2 illustrates the principle of generation of a thermal radiation pressure
  • FIG. 3 illustrates the principle of generation of thermal radiation torque
  • FIG. 4 is an external view of a satellite according to a first embodiment of the present invention.
  • FIG. 5 illustrates the configuration of an attitude control system for a satellite according to the first embodiment of the present invention
  • FIG. 6 is a block diagram of the attitude control system for a satellite according to the first embodiment of the present invention.
  • FIG. 7 illustrates the relationship between a satellite and a coordinate system according to the first embodiment of the present invention
  • FIG. 8 illustrates a result of simulation of a change in the angle of a satellite 1 in a gravity gradient stable state
  • FIGS. 9( a ) and 9( b ) illustrate results of simulation of an attitude control system for a satellite according to the first embodiment of the present invention
  • FIG. 10 illustrates a result of simulation of the attitude control system for a satellite according to the first embodiment of the present invention
  • FIGS. 11( a ) to 11( d ) illustrate results of simulation of the attitude control system for a satellite according to the first embodiment of the present invention
  • FIG. 12 illustrates external appearance and arrangement of a first satellite and a second satellite according to a second embodiment of the present invention.
  • FIG. 13 illustrates the entire configuration of a system for controlling a relative position and velocity of the first satellite and the second satellite according to the second embodiment of the present invention.
  • the solar array panels absorb the solar energy by the amount corresponding to an absorption ratio C abs of the solar array panels.
  • the solar array panels have an electricity generation ratio expressed as C abs ⁇ .
  • some of the thermal energy is consumed for increasing the temperature of the bodies of the solar array panels, most of the thermal energy is emitted to the outside from the surfaces of the solar array panels as thermal radiation.
  • the amounts of energy W out front and W out back radiated from the light receiving surface of the solar array panel and from the surface opposite to the light receiving surface are expressed based on the Stefan-Boltzmann law as described below:
  • ⁇ is a Stefan-Boltzmann constant
  • ⁇ f and ⁇ b are thermal emissivities from the light receiving surface and from the surface opposite to the light receiving surface, respectively.
  • f front B f ⁇ ⁇ f ⁇ ⁇ ⁇ T f 4 c ⁇ [ N ⁇ / ⁇ m 2 ] ( 6 )
  • f back B b ⁇ ⁇ b ⁇ ⁇ ⁇ T b 4 c ⁇ [ N ⁇ / ⁇ m 2 ] ( 7 )
  • a difference between the pressures acts on the solar array panels as a following thermal radiation pressure in a direction of a normal to the light receiving surface, and a direction from the light receiving surface of the solar array panel to the surface opposite to the light receiving surface:
  • B f are Lambertian coefficients of the light receiving surface and the surface opposite to the light receiving surface, respectively.
  • Such a mechanism makes it possible to generate thrust with use of the thermal radiation pressures from the solar array panels.
  • the thermal radiation pressure in a direction from the light receiving surface of the solar array panel toward the surface opposite to the light receiving surface at a right angle with respect to the light receiving surface may be increased by (i) increasing the temperature of the light receiving surface during operation since the thermal radiation pressure is proportional to the 4th power of temperature, (ii) increasing a temperature difference between the light receiving surface and the surface opposite to the light receiving surface, and (iii) providing a large difference in thermal emissivity ⁇ between the light receiving surface and the surface opposite to the light receiving surface.
  • the light receiving surface and the surface opposite to the light receiving surface are made of materials with a high thermal emissivity in order to implement active emission of heat from their surfaces.
  • the light receiving surface of the solar array panels has a thermal emissivity ⁇ f of 0.8
  • the surface opposite to the light receiving surface is made of carbon fiber reinforced plastics (CFRP) having a thermal emissivity ⁇ b of 0.8.
  • An aluminum honeycomb with high heat conductivity is provided between the light receiving surface and the surface opposite to the light receiving surface. Since the aluminum honeycomb is thick to some extent, a temperature T f of the light receiving surface is higher than a temperature T b of the surface opposite to the light receiving surface during operation of the solar array panels.
  • the temperature T f of the light receiving surface is higher than the temperature T b of the surface opposite to the light receiving surface as described before.
  • the solar array panels having such a configuration are referred to as “solar array panels having configuration 1.”
  • the solar array panels are made to have a sufficiently small thickness and a material with large thermal conductivity is used, so that the light receiving surface and the surface opposite to the light receiving surface have the same temperature and their time constants for temperature changes are made sufficiently small.
  • the solar array panels are made to have an absorption ratio C abs of 0.8, and their light receiving surfaces have a thermal emissivity ⁇ f of 0.8.
  • thermo emissivity ⁇ b 0.05
  • heat emission from the surface opposite to the light receiving surface is suppressed to increase the temperature of the solar array panels and to thereby increase the thermal radiation pressure.
  • the solar array panels having such a configuration are referred to as “solar array panels having configuration 2.”
  • a material with small thermal conductivity may be used for the member placed between the light receiving surface and the surface opposite to the light receiving surface or the solar array panels may be made to have a larger thickness, for example. As a result, a temperature difference between the light receiving surface and the surface opposite to the light receiving surface becomes larger, so that a larger thermal radiation pressure can be generated.
  • thermal radiation torque which is a product of a difference f TRP between thermal radiation pressures, a surface area of the light receiving surface of the solar array panel, and an arm length from the center of rotation O to the center of pressure of the solar array panel.
  • FIG. 4 is an external view of a satellite 1 as a spacecraft according to a first embodiment of the present invention.
  • the satellite 1 is a 2t-class satellite, which includes a satellite body 10 and two solar array panels 21 and 22 .
  • the satellite body 10 has a cubic shape with one side being 2 m and the center of the cube being the center of mass.
  • the solar array panels 21 and 22 have rectangular light receiving surfaces 210 and 220 with a shorter side being 3 m and a longer side being 10 m.
  • the solar array panels 21 and 22 have the aforementioned configuration 2 with aluminum being vapor-deposited on the surface opposite to the light receiving surface.
  • the solar array panels 21 and 22 are completely axially symmetrically arranged for coordinate axes of a later-described body-fixed frame, with respect to the center of mass of the satellite body 10 .
  • the solar array panels 21 and 22 include solar cell arrays 212 , 214 , 222 , and 224 which are divisions of the solar array panels.
  • the solar cell arrays each have a longer side parallel to the longer side of the light receiving surface 210 of the solar array panel 21 , and their light receiving surfaces are equal in area.
  • each of the solar cell arrays 212 , 214 , 222 , and 224 concentrates on a central one point (center of pressure) of each of the solar cell arrays 212 , 214 , 222 , and 224 .
  • a straight line connecting the centers of pressure of the solar cell arrays 212 and 224 and a straight line connecting the centers of pressure of the solar cell arrays 214 and 222 each pass through the center of mass. Therefore, in the satellite 1 according to the first embodiment, the solar array panels 21 and 22 mutually negate influence of a solar radiation pressure, which makes it possible to disregard the influence of disturbance by the solar radiation pressure torque.
  • the attitude of the satellite 1 can be controlled with respect to two axes around a travelling direction and a center of the earth direction.
  • Table 2 indicates moment of inertia of the satellite 1 .
  • Table 3 indicates gravity gradient torque with respect to the altitude acting on the satellite 1 .
  • FIG. 5 illustrates the configuration of an attitude control system 3 of the satellite 1 .
  • FIG. 6 is a block diagram of the attitude control system 3 of the satellite 1 .
  • An attitude detector 31 includes a star sensor and an earth sensor to detect an attitude angle of the satellite body 10 .
  • a position detector 33 includes a GPS positioning device to detect a position of the satellite 1 itself. The detected attitude angle is input into the attitude controller 35 from the attitude detector 31 . The detected position of the satellite 1 itself is input into the attitude controller 35 from the position detector 33 . The attitude detected by the attitude controller 35 may be an appropriate parameter other than the attitude angle.
  • a torque generator 36 includes a thruster 361 , a wheel 362 , and the solar array panels 21 and 22 .
  • a target attitude setting unit 34 sets a target attitude angle that is a target attitude.
  • a target attitude an appropriate parameter other than the target attitude angle may be used.
  • the target attitude to be set is typically an attitude of the solar array panels 21 and 22 which points to the sun for a later-described reason.
  • the target attitude is not limited thereto.
  • the attitude controller 35 gives an attitude change amount as a controlled variable to one unit selected out of the thruster 361 , the wheel 362 , and the solar array panels 21 and 22 included in the torque generator 36 .
  • the attitude controller 35 may simultaneously control a plurality of units selected out of the thruster 361 , the wheel 362 , and the solar array panels 21 and 22 included in the torque generator 36 .
  • the solar cell arrays 212 , 214 , 222 , and 224 each include a shunt switch to cause short-circuit as necessary.
  • the attitude controller 35 can individually command turning on, turning off, and switching (repetition of turning on/off) of each shunt switch so as to continuously change the electricity generation ratio of each of the solar cell arrays 212 , 214 , 222 , and 224 in the range of 0 to 30%.
  • the attitude controller 35 gives an attitude change amount to the torque generator 36 as a controlled variable based on an attitude angle from the attitude detector 31 and a target attitude angle from the target attitude setting unit 34 . As a consequence, the attitude of the satellite 1 is controlled.
  • the attitude controller 35 may be provided in a terrestrial station to control the attitude of the satellite 1 through communication between the terrestrial station and the satellite 1 .
  • the attitude controller 35 may also be provided in another spacecraft which is not illustrated.
  • FIG. 7 illustrates the relationship between a satellite orbiting a celestial object and a coordinate system.
  • the satellite 1 can generate thermal radiation torque whose maximum value is about 1.27 ⁇ 10 ⁇ 4 N ⁇ m around an i B axis of the later-described body-fixed frame and about 3.17 ⁇ 10 ⁇ 5 N ⁇ m around a k B axis.
  • thermal radiation torque whose maximum value is about 1.27 ⁇ 10 ⁇ 4 N ⁇ m around an i B axis of the later-described body-fixed frame and about 3.17 ⁇ 10 ⁇ 5 N ⁇ m around a k B axis.
  • These values represent microscopic power between that approximately equal to and 1/100 of the values of the gravity gradient torque of Table 3.
  • typical application examples of the attitude control using thermal radiation torque include but are not limited to the case of damping oscillation of the satellite 1 which is caused by disturbance applied to the satellite 1 in the gravity stable state. In this case, the longer side of the solar array panel needs to extend in the center of the earth direction.
  • the satellite 1 Since the thermal radiation torque is generated by short-circuiting the solar array panel, the satellite 1 needs a margin of electricity. Therefore, without being limited thereto, the most desirable orbit of the satellite is a sun-synchronous orbit in which the solar array panels point to the sun and the satellite is always exposed to the sun.
  • FIG. 7 illustrates an example of the orbit of the satellite 1 in consideration of the above conditions.
  • the orbit-fixed frame i O j O k O has an origin at the center of mass of the entire satellite 1 .
  • the orbit-fixed frame i O j O k O includes an i O axis in an orbiting direction, an ⁇ j O axis in the sun direction, and a k O axis in the center of the earth direction.
  • the body-fixed frame i B j B k B is fixed to the satellite body 10 .
  • the body-fixed frame i B j B k B has an origin at the center of mass of the entire satellite 1 .
  • the body-fixed frame matches the orbit-fixed frame. That is, the attitude angle of the satellite 1 is a parameter representing a difference in direction between the orbit-fixed frame and the body-fixed frame.
  • 2-1-3 Euler angle transformation is used to express deviation of the body-fixed frame from the orbit-fixed frame.
  • a rotation angle around the travelling direction is ⁇
  • a rotation angle around the solar direction is ⁇
  • a rotation angle around the center of the earth direction is ⁇ .
  • the attitude controller 35 can individually control each of the electricity generation ratios of the solar cell arrays 212 , 214 , 222 , and 224 as described before. Therefore, according to the aforementioned theory of operation of the torque generation system, it becomes possible to control the attitude of the satellite 1 around the i O axis in the travelling direction by controlling each of the electricity generation ratios of an entire group of the solar cell arrays 212 and 214 and an entire group of the solar cell arrays 222 and 224 . Moreover, it becomes possible to control the attitude of the satellite 1 around the k O axis in the center of the earth direction by controlling each of the electricity generation ratios of an entire group of the solar cell arrays 212 and 222 and an entire group of the solar cell arrays 214 and 224 .
  • the attitude controller 35 can control the attitude of the satellite 1 around two axes including the i O axis in the travelling direction and the k O axis in the center of the earth direction.
  • a current attitude angle is detected by the attitude detector 31 provided in the satellite body 10 .
  • a target attitude angle is set by the target attitude setting unit 34 .
  • the current attitude angle output from the attitude detector 31 is fed back to calculate an error between the current attitude angle and the target attitude angle output from the target attitude setting unit 34 .
  • the attitude controller 35 performs attitude control calculation, such as proportional-derivative (PD) control.
  • PD proportional-derivative
  • one unit selected out of the thruster 361 , the wheel 362 , and the solar array panels 21 and 22 included in the torque generator 36 is operated to perform attitude control. Selection of the thruster 361 , the wheel 362 , and the solar array panels 21 and 22 may be determined based on the value of the controlled variable and the like.
  • the attitude controller 35 controls opening and closing of each of the shunt switches of the solar cell arrays 212 , 214 , 222 , and 224 based on the result of attitude control calculation so as to control the electricity generation ratios of the respective arrays.
  • an attitude angular velocity detector may further be provided to perform feedback control based on an attitude angle and an attitude angular velocity.
  • the attitude controller 35 performs feedback control.
  • the feedback control feedforward control may be performed.
  • a disturbance torque estimation unit for estimating disturbance torque may be provided, and an electricity generation ratio of the solar array panels and/or the solar array panel divisions may be controlled to generate torque that suppresses the disturbance torque estimated by the disturbance torque estimation unit.
  • the electricity generation ratios of the solar cell arrays 212 , 214 , 222 , and 224 are each controlled to damp the oscillation with the thermal radiation pressure.
  • FIGS. 9( a ), 9( b ) and 10 illustrate the result of performing PD control by the attitude controller 35 .
  • FIGS. 9( a ) and 9( b ) illustrate comparison between temporal changes in uncontrolled and controlled ⁇ and ⁇ .
  • FIG. 10 illustrates temporal changes in ⁇ and ⁇ .
  • FIGS. 11( a ) to 11( d ) illustrate changes in ⁇ in each of the solar cell arrays 212 , 214 , 222 , and 224 (when the absorption ratio C abs of solar cells is 0.8, ⁇ corresponds to 0 to 0.375 with the electricity generation ratio being in the range of 0 to 30% according to Expression (3)).
  • each of the solar cell arrays 212 , 214 , 222 , and 224 It takes about 6 to 7 minutes for each of the solar cell arrays 212 , 214 , 222 , and 224 to reach an equilibrium temperature from the time of switchover from the maximum output state to the shunt state. Since the time is sufficiently small as compared with a time constant for the attitude change, the time taken for temperature change in each of the solar cell arrays 212 , 214 , 222 , and 224 is negligible.
  • FIGS. 9( a ), 9( b ) and 10 indicate that the attitude control of the present embodiment suppresses the oscillation of the satellite 1 fairly well.
  • FIG. 10 also indicates that an angle change rate of the aforementioned magnetic wheel, which is 360 arcsec/1 hour, is suppressed to about 1 arcsec/1 hour in attitude control of the present embodiment. This indicates that very high-precision attitude control is accomplished.
  • the solar array panels having configuration 2 in which aluminum is vapor-deposited on the surface opposite to the light receiving surface are used.
  • the solar array panels having configuration 1 that are conventional solar array panels in which the light receiving surface and the surface opposite to the light receiving surface have the same thermal emissivity.
  • maximum values of the thermal radiation torque when the solar array panels having the same shape as that in the above embodiment are used are 7.70 ⁇ 10 ⁇ 6 N ⁇ m around the k B axis and 1.93 ⁇ 10 ⁇ 6 N ⁇ m around the k B axis in the center of the earth direction, which are small to resist the gravity gradient torque illustrated in Table 3.
  • the thermal radiation torque is proportional to the area and the arm length of the light receiving surface of solar array panel, and therefore if the area of the light receiving surface of the solar array panel is enlarged such that at least the arm length is increased, the thermal radiation torque is increased at least in proportion to the square of the arm length. As a result, it becomes possible to resist the gravity gradient torque. For example, in the case of using solar array panels with a shorter side of 5 m and a longer side of 15 m, thermal radiation torque of 4.0 ⁇ 10-5 N ⁇ m can be obtained, which is enough to resist the gravity gradient torque.
  • the unit of the control of the electricity generation ratio is the solar cell array, which is one example of the solar array panel divisions, having a shape which is formed by dividing the solar array panel into two, and the control is executed by continuously changing the electricity generation ratio for each of the entire solar cell arrays.
  • the electricity generation ratios of the solar array panels or the solar array panel divisions may be controlled by executing control for switching the maximum output state and the shunt state for each of a plurality of solar cell groups each of which has a region smaller than the solar array panels or the solar array panel divisions that are the units of the control of the electricity generation ratio, and controlling the areas and the positions of the solar cell groups in the maximum output state and the solar cell groups in the shunt state.
  • the solar cell groups in the shunt state are arranged on the side far from the satellite body, it becomes possible to generate, with the same electricity generation ratio, larger torque than that in the case where control is executed to continuously change the electricity generation ratio of each of the entire solar array panels or the solar array panel divisions. This means that the same torque can be generated at a smaller electricity generation ratio, so that a required margin of electricity can be reduced.
  • the electricity generation ratios of the solar cell or cells may also be controlled per region by attaching a film with a variable transmittance to each region on the light receiving surface of the solar array panel and controlling the transmittance of each region.
  • each of the embodiments decreasing the area of each of the solar array panel divisions and/or the area of each of the solar cell groups enables more detailed and complicated control to be executed.
  • One of such examples is to arrange solar array panel divisions and/or solar cell groups, which are each identical in shape and small in area, in a matrix form. It goes without saying that the shapes and areas of the solar array panel divisions and/or the solar cell groups may be different from each other.
  • the satellite body 10 has a cubic shape with one side being 2 m.
  • the satellite 1 can be made to have any appropriate shape and size.
  • any appropriate configuration, such as the configuration 1, may be adopted.
  • FIG. 12 illustrates external appearance and arrangement of a first satellite 41 as a first spacecraft and a second satellite 42 as a second spacecraft according to a second embodiment of the present invention.
  • the first satellite 41 and the second satellite 42 are solar-sailing ultra-small satellites, which include satellite bodies 410 and 420 and solar array panels 411 and 421 , respectively.
  • the satellite bodies 410 and 420 each have a cubic shape with one side being 0.1 m.
  • the solar array panels 411 and 421 have square light receiving surfaces 4110 and 4210 with one side being 0.75 m, respectively.
  • the satellite bodies 410 and 420 are arranged in the center of surfaces opposite to the light receiving surfaces of the solar array panels 411 and 421 , respectively.
  • FIG. 13 illustrates the entire configuration of a system 4 for controlling a relative position and/or relative velocity of the satellite 41 as a first spacecraft and the satellite 42 as a second spacecraft according to the second embodiment of the present invention.
  • the first satellite 41 has the solar array panel 411 , a controller 412 , an attitude detector 413 , a position detector 414 , and a communication unit 415 .
  • the second satellite 42 has the solar array panel 421 , a controller 422 , an attitude detector 423 , a position detector 424 , and a communication unit 425 .
  • the attitude detectors 413 and 423 include a star sensor and an earth sensor to detect attitude angles of the satellite bodies 410 and 420 , respectively.
  • the position detectors 414 and 424 include a GPS positioning device to detect the positions of the satellite 41 and the satellite 42 themselves.
  • the attitude detected by the attitude detectors 413 and 423 may be an appropriate parameter other than the attitude angle.
  • a terrestrial station 43 includes a relative position and velocity setting unit 431 , a relative position and velocity calculation unit 432 , a relative position and velocity controller 433 , and a communication device 434 .
  • the relative position and velocity calculation unit 432 and the relative position and velocity controller 433 may be provided in the first satellite 41 , the second satellite 42 , and/or other spacecraft which are not illustrated.
  • the communication units 415 and 425 of the first satellite 41 and the second satellite 42 and the communication device 434 of the terrestrial station 43 each have an antenna which is not illustrated, and mutually exchange information through the antennas.
  • the terrestrial station 43 transmits control information to the communication units 415 and 425 .
  • the communication units 415 and 425 transmit information, such as attitude angles detected by the attitude detectors 413 and 423 and positions of the satellite 41 and the satellite 42 themselves detected by the position detectors 414 and 424 , to the terrestrial station 43 , respectively.
  • a target relative position and a target relative velocity of the second satellite 42 relative to the first satellite 41 can be set.
  • the relative position and velocity calculation unit 432 calculates the relative position and velocity of the second satellite 42 relative to the first satellite based on the information on the attitude angles and the positions received from the first satellite 41 and the second satellite 42 .
  • the relative position and velocity controller 433 Based on the relative position and/or the relative velocity calculated by the relative position and velocity calculation unit 432 and on the target relative position and/or the target relative velocity set by the relative position and velocity setting unit 431 , the relative position and velocity controller 433 generates control information for controlling the electricity generation ratio of each of the solar array panels 411 and 421 of the first satellite 41 and the second satellite 42 so as to cause each of the first satellite 41 and the second satellite 42 to generate thrust whereby the relative position and/or relative velocity become the target relative position and/or target relative velocity through execution of control such as feedback control and feedforward control. The relative position and velocity controller 433 then transmits the generated control information to the first satellite 41 and the second satellite 42 through the communication device 434 .
  • the controllers 412 and 422 of the first satellite 41 and the second satellite 42 set the electricity generation ratios of the solar array panels 411 and 421 based on the control signals sent from the relative position and velocity controller 433 and received by the communication units 415 and 425 , respectively. As a consequence, the thrust based on the thermal radiation pressure generated from the solar array panels 411 and 421 is controlled.
  • the attitudes of the first satellite 41 and the second satellite 42 are controlled by the controllers 412 and 422 so that the solar array panels 411 and 421 point to the sun.
  • the control of the attitudes of the first satellite 41 and the second satellite 42 is not limited thereto.
  • the solar array panel 411 of the first satellite 41 is controlled to be in a maximum output state to generate a minimum thermal radiation pressure while the solar array panels 421 of the second satellite 42 is controlled to be in a shunt state to generate a maximum thermal radiation pressure.
  • the satellite bodies 410 and 420 of the first satellite 41 and the second satellite 42 have a cubic shape with one side being 0.1 m.
  • the first satellite 41 and the second satellite 42 can be made to have any appropriate shape and size.
  • any appropriate configuration, such as the configuration 1, may be adopted.

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