JP2005067601A - Satellite attitude control method - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a satellite attitude control method capable of controlling the attitude of a satellite and stabilizing power generating efficiency at the maximum by correctly facing the light receiving surface of a solar battery to the sunlight, without forming the solar battery in a substantially spherical shape. <P>SOLUTION: A solar battery light receiving surface for receiving the sunlight is disposed so that it is substantially parallel with a plane including a first attitude control axis 24x and a second attitude control axis 24y intersecting orthogonally. The second attitude control axis 24y is kept vertically to an ecliptic surface 32 as an orbital surface of the earth moving circumferentially about the sun by the first attitude control axis 24x, and the light receiving surface of the solar battery is correctly faced to the sunlight by the second attitude control axis 24y. <P>COPYRIGHT: (C)2005,JPO&NCIPI

Description

  The present invention relates to a satellite attitude control method in satellite communication in which a satellite having a communication function provides a communication service to the earth while orbiting a quasi-stationary satellite orbit.

  The quasi-stationary satellite orbit is, for example, as disclosed in Non-Patent Document 1, for a geostationary satellite orbit (that is, an orbit having a satellite period of about 24 hours and an orbit altitude of about 36000 km). An orbit of a satellite that orbits the earth at an arbitrary period on a surface inclined at an arbitrary angle, and has a gentle movement of the satellite as viewed from the earth in an arbitrary time zone.

  FIG. 7 is a diagram showing a configuration of a satellite orbit using a conventional quasi-stationary satellite orbit. In FIG. 7, 1 is the earth on which communication service is provided from the satellite 4, 2 is the equator plane of the earth 1, 3 is a quasi-stationary satellite orbit inclined at an arbitrary angle with the equator plane 2, and 4 is a quasi-stationary satellite orbit 3. A satellite that orbits at an arbitrary satellite cycle, 5 is the near point of the quasi-stationary satellite orbit 3 where the altitude from the surface of the earth 1 is the lowest (hereinafter referred to as the Perige point), 6 is the highest altitude from the surface of the earth 1 This is a far point of the quasi-stationary satellite orbit 3 (hereinafter referred to as an apogee point). Reference numeral 7 denotes a point where the equator plane 2 and the quasi-stationary satellite orbit 3 intersect, and is an ascending intersection of the quasi-stationary satellite orbit 3.

Next, the operation will be described.
In FIG. 7, the satellite 4 moves around the earth 1 along the quasi-stationary satellite orbit 3 from the position of the illustrated satellite 4 in the order of the apogee point 6, the ascending point 7, and the peripheral point 5. Orbit to the position of. The satellite period when the satellite 4 orbits is set to an appropriate value according to the form of the communication service.

  Since the quasi-stationary satellite orbit 3 is generally an elliptical orbit, the orbiting speed of the satellite 4 varies depending on the position of the satellite 4 on the quasi-stationary satellite orbit 3. That is, from Kepler's second law, the magnitude of the orbiting speed of the satellite 4 changes so that the area velocity of the line segment connecting the earth 1 and the satellite 4 is constant, and at the margin point 5 and the apogee point 6, The magnitude of the circulation speed is maximum and minimum, respectively.

  Since the orbiting speed of the satellite 4 is minimum at the Apoge point 6, when the satellite 4 is observed from the earth 1, the satellite 4 at the Apoge point 6 appears to be almost stationary. Therefore, in satellite communication using the quasi-stationary satellite orbit 3, communication service can be provided to the earth 1 using the apogee point 6 as an operating point for satellite communication in a form similar to satellite communication using the geostationary satellite orbit.

  When the rotation period of the earth 1 and the satellite period of the satellite 4 coincide with each other, only one apogee point 6 is projected onto the earth 1. On the other hand, when the satellite period of the satellite 4 is shorter than the rotation period of the earth 1, the apogee point 6 has a plurality of points on the earth 1 (hereinafter referred to as apogee point projection points) according to the difference between these periods. ) Is projected.

  Therefore, when the satellite period of the satellite 4 is shorter than the rotation period of the earth 1, the communication service is provided every time the satellite 4 is positioned at the apogee point 6, and the apogee point projection point is switched. When the required number of orbital movements determined by the relationship between the rotation period of the earth 1 and the satellite period of the satellite 4 is performed, the satellite 4 again at the apoge point 6 with respect to the apoge point projection point where the communication service was first provided. Semi-synchronize. As described above, in the satellite communication system using the quasi-stationary satellite orbit, the number of apogee projection points, which are communication service providing areas, determines the form of the communication service.

  Further, since the quasi-stationary satellite orbit 3 is inclined at an arbitrary angle with respect to the equator plane 2, the satellite 4 is compared to the satellites on the equator plane 2 in the geostationary satellite orbit. Since a high latitude area such as Australia is irradiated at a high elevation angle, communication services can be provided to the earth 1 while reducing interference from buildings and mountains existing in the high latitude area.

  Further, in satellite communication such as a medium and low orbit satellite system using a plurality of satellites 4 orbiting a plurality of quasi-stationary satellite orbits 3, a communication function is provided to cope with the failure of the plurality of satellites 4 (not shown). A spare satellite is provided, and the orbit of the spare satellite (hereinafter referred to as a standby orbit) is arranged in an orbit different from the plurality of quasi-stationary satellite orbits 3.

  If there are no failures in the plurality of satellites 4, the spare satellites orbit around the standby orbit and do not perform communication services. When a failure occurs in a plurality of satellites 4, a spare satellite is inserted into the quasi-stationary satellite orbit 3 in which the failed satellite 4 orbited, and the communication service is replaced with the earth 1 in place of the failed satellite 4. provide.

  In addition, the satellite 4 is provided with a solar cell (not shown), and sunlight is converted into electric power by the solar cell and used as a power source of the satellite 4. The larger the light receiving surface of the solar cell, the larger the generated power of the solar cell, and the higher the incident angle of sunlight with respect to the light receiving surface, the higher the power generation efficiency of the solar cell.

  As in the quasi-stationary satellite orbit 3, the orbit of the satellite 4 does not exist on the ecliptic plane (not shown) that is inclined by about 23 ° with respect to the equatorial plane (that is, the orbital plane of the earth 1 that moves around the sun (not shown)). In this case, since the incident angle of sunlight with respect to the light receiving surface of the solar cell changes depending on the position of the satellite 4 on the quasi-stationary satellite orbit 3, there is a problem that the power generation efficiency of the solar cell is not stable. As a method for solving this problem, Patent Document 1 discloses a technique in which a solar cell is formed in a substantially spherical shape so that the incident angle is constant even when the position of the satellite 4 changes.

“Current Status and Trends of Non-stationary Mobile Satellite Communication Systems” (Satellite Communication Research, No. 75, KDD Engineering and Consulting, September 1998) JP-A-1-257700

  The conventional quasi-stationary satellite orbit is configured as described above, but when constructing a satellite communication system that provides communication services using a plurality of quasi-stationary satellite orbits, an apogee point projection point that affects the form of the communication service In the past, an efficient satellite orbit configuration method that correlates parameters such as the satellite period of a satellite that orbits a quasi-stationary satellite orbit and the number of satellites required for the entire system has been established. There was a problem that was not done.

  In addition, the necessity of a standby satellite to cope with the failure of multiple satellites that orbit each of multiple quasi-stationary satellite orbits has been pointed out in the past. However, there has been a problem that the standby orbit of the standby satellite must be determined so that the standby satellite is in an optimal standby state for any of a plurality of quasi-stationary satellite orbits.

  Furthermore, it is difficult to increase the size of the solar cell with the technique of configuring the solar cell in a substantially spherical shape disclosed in Patent Document 1, and is inappropriate as a power source for satellites that require a large amount of power supply. There was a problem of being.

  The present invention has been made to solve the above-described problems, and a plurality of quasi-stationary satellite orbits are formed by relating the number of satellites and the number of satellites corresponding to the number of apogee projection points. It is an object of the present invention to provide a satellite orbit configuration method capable of efficiently arranging satellites in an area where communication services are provided.

  Another object of the present invention is to provide a satellite orbit configuration method and a satellite orbit system in which a standby satellite is in an optimum standby state for any of a plurality of quasi-stationary satellite orbits.

  Furthermore, the present invention can stabilize the power generation efficiency to the maximum while controlling the attitude of the satellite without facing the substantially spherical solar cell and making the light receiving surface of the solar cell directly face sunlight. The purpose is to obtain a satellite attitude control method.

  A satellite attitude control method according to the present invention includes a solar cell light receiving surface that receives sunlight so as to be substantially parallel to a plane including a first attitude control axis and a second attitude control axis that are orthogonal to each other. The second attitude control axis keeps the second attitude control axis perpendicular to the ecliptic plane that is the orbital surface of the earth that orbits around the sun, and the second attitude control axis makes the solar cell The light receiving surface is made to face sunlight.

  According to this invention, the solar cell light receiving surface for receiving sunlight is provided so as to be substantially parallel to a surface including the first posture control axis and the second posture control axis orthogonal to each other, and the first posture control is performed. The second attitude control axis is kept perpendicular to the ecliptic plane, which is the orbit of the earth that orbits around the sun, by the axis, and the solar cell light-receiving surface is opposed to sunlight by the second attitude control axis. Since it did in this way, the effect that the attitude | position of a satellite is stabilized with respect to disturbance by the gravity inclination of the earth, the solar wind, etc. by simple biaxial control, and the electric power generation efficiency of a solar cell can be stabilized to the maximum is acquired.

An embodiment of the present invention will be described below.
Embodiment 1 FIG.
As described above, in satellite communications using quasi-stationary satellite orbits, the speed of the satellite is minimized at the apogee point, and high-elevation communication services are provided to the earth in a manner similar to satellite communications using geostationary satellite orbits. Can be provided. Therefore, when designing a quasi-stationary satellite orbit, it is necessary to determine the number of apogee projection points that serve as an area for providing communication services.

Here, the number of apogee point projection points is Na, the earth's rotation period is Te (about 24 hours), and the number of days required for the satellite to be semi-synchronized with the earth, that is, a certain point on the earth becomes the apogee point projection point. If Nc is the number of days (hereinafter referred to as quasi-synchronization required days) from the next to the next apoge point projection point, and Ts is the satellite period required for the satellite to orbit around the quasi-stationary satellite orbit once (1) ) Holds.
Ts = Te × Nc ÷ Na (1)

In equation (1), the number of days required for quasi-synchronization Nc and the number Na of apogee projection points are in an integer relationship that is relatively prime. The reason is as follows. Assuming that the quasi-synchronization required number of days Nc and the number Na of apogee point projection points have a common factor a, it can be expressed as in equation (2).
Nc = Ncx × a, Na = Nax × a (2)

Here, Ncx and Nax have an integer relationship that is relatively prime. Substituting equation (2) into the left side of equation (1), as shown in equation (3) below,
Te × Nc ÷ Na = Te × Ncx ÷ Nax (3)
This is because the number of apogee projection points becomes Nax, resulting in a contradiction.

  The fact that the number of days required for quasi-synchronization Nc and the number Na of apogee point projection points are in a non-prime integer relationship specifically means that the same apogee point projection points are counted repeatedly. It is. For example, if the number Na of apogee point projection points is 12 points, but the number of days required for quasi-synchronization Nc is 9 days that are not prime numbers, the same apogee point projection points are duplicated three times. It means to count. In addition, in order for the satellite to be semi-synchronized with the earth, the satellite cycle Ts should not be an infinite number.

Ns indicates the number of satellites necessary to provide a communication service for about 24 hours a day for any apoge point projection point, and one satellite provides a communication service for any apoge point projection point When the operation time is To, Equation (4) is established.
Ns = Te × Nc ÷ To (4)

  FIG. 1 is a flowchart showing a satellite orbit configuration method according to Embodiment 1 of the present invention. A specific example of the satellite orbit configuration method according to the first embodiment of the present invention will be described below with reference to FIG.

  Here, as an example, let us consider a case where the apogee projection points are arranged at intervals of 30 ° in the longitude direction of the earth. Therefore, the number Na of the apogee projection points is obtained as 12 points by dividing 360 ° (an angle for one round of the earth) by the longitude interval of 30 ° (step ST1). Considering that the number of days required for quasi-synchronization Nc and the number Na of apogee projection points are in a prime integer relationship, the number of days that are candidates for Nc are 5, 7, 11, and so on. When the quasi-synchronization required number of days Nc is 5 days (step ST2), the satellite cycle Ts is calculated as 10 hours from the equation (1) (step ST3).

  Here, when the satellite period Ts is 10 hours, the orbital altitude of the satellite obtained from the satellite period Ts is too low, so it is determined that it is not appropriate for constructing the quasi-stationary satellite orbit (step ST4). Return to ST5. The criteria for determining the satellite's orbital altitude may vary depending on the form of the satellite communication system. Here, as an example, the orbital altitude when Ts = 10 hours is low, which is inappropriate for the satellite communication system considered here. It is determined.

  In step ST5, the quasi-synchronization required number of days Nc is changed to the next candidate 7th, and the process proceeds to step ST3. When the satellite cycle Ts is obtained again from the equation (1), the satellite cycle Ts is calculated as 14 hours (step ST3). It is determined that the orbit of the satellite when the satellite period Ts is 14 hours is a sufficient orbit altitude (step ST4), and the process proceeds to the next step ST6. As described above, as an example, the orbital altitude at Ts = 14 hours is determined to be a sufficient orbital altitude. In step ST6, the satellite operating time To is determined to be, for example, 8 hours.

  FIG. 2 is a world map projecting a quasi-stationary satellite orbit in which the satellite orbits when the number of apoge point projection points Na is 12, the satellite period Ts is 14 hours, and the satellite operating time To is 8 hours. In FIG. 2, the horizontal axis represents longitude, the vertical axis represents latitude, 100 represents a quasi-stationary satellite orbit projected on a world map, and 200 represents an apogee point projection point. The specifications of this quasi-stationary satellite orbit 100 have a perigi point of altitude of about 9000 km, an apogee point of altitude of about 37000 km, and an eccentricity representing the degree of ellipse of the quasi-stationary satellite orbit is 0.48. From FIG. 2, it can be seen that satellites can be seen at intervals of 30 ° longitude between 30 ° and 45 ° north latitude, and there are 12 apogee point projection points 200.

  Under the above conditions, the number of satellites Ns required to provide a communication service of about 24 hours per day to any apogee projection point is calculated using equation (4). The number Ns is 21 (step ST7). Therefore, the number of satellites with respect to one apoge point projection point is 1.75, and it can be seen that the satellites can be efficiently arranged considering the number of satellites with respect to the area where the communication service is provided.

  As described above, according to the first embodiment, the step of determining the quasi-synchronized required number of days Nc so as to have a relatively prime integer relationship corresponding to the number Na of apogee projection points, A step of obtaining a satellite cycle Ts based on the rotation period Te, the number of days required for quasi-synchronization Nc, and the number Na of apogee projection points, and a step of determining the validity of the orbit altitude of the satellite corresponding to the satellite cycle Ts; Since the rotation period Te of the earth, the required number of quasi-synchronization days Nc, and the step of obtaining the number Ns of satellites based on the operation time To of the satellite, the satellite corresponding to the number Na of the apogee projection points is provided. A plurality of quasi-stationary satellite orbits can be configured by relating the period Ts and the number Ns of satellites, so that the satellites can be efficiently arranged in the area where the communication service is provided. The effect is obtained that that.

  In addition, when the orbital altitude of the satellite is inappropriate, the number of days of quasi-synchronization Nc, which is an integer that is relatively prime to the number of apogee projection points Na, is changed to another number of days. The quasi-synchronized required number of days Nc corresponding to the orbital altitude of the satellite can be determined, and the satellite obtained based on the rotation period Te of the earth, the quasi-synchronized required number of days Nc, and the number Na of the apogee projection points. The effect is obtained that the period Ts can be obtained in correspondence with the orbital altitude of an appropriate satellite.

Embodiment 2. FIG.
FIG. 3 is a diagram showing the configuration of a satellite orbit according to Embodiment 2 of the present invention. In FIG. 3, 11 is the earth that receives the communication service from the satellites 14a, 14b, and 14c, and 12 is a standby orbit in which the spare satellite 14d provided to cope with the failure of the satellites 14a, 14b, and 14c moves around. The standby orbit 12 in the second embodiment is a geostationary satellite orbit that exists in the equatorial plane of the earth 11 and has a satellite period of about 24 hours and an orbital altitude of about 36000 km.

  13a, 13b, and 13c are quasi-stationary satellite orbits inclined at an arbitrary angle with respect to the equator plane, 14a, 14b, and 14c are satellites that orbit around the quasi-stationary satellite orbits 13a, 13b, and 13c, and 14d is a satellite 14a, 14b, and It is a spare satellite having a communication function equivalent to 14c and orbiting the standby orbit 12.

  Reference numerals 15a, 15b, and 15c denote the peripheral points of the quasi-stationary satellite orbits 13a, 13b, and 13c, reference numerals 16a, 16b, and 16c denote the apogee points of the quasi-stationary satellite orbits 13a, 13b, and 13c, and reference numerals 17a, 17b, and 17c respectively. This is the ascending intersection of the tracks 13a, 13b, 13c. The three peripheral points 15a, 15b, 15c all have a 120 ° longitude interval, the longitude interval of the three apoge points 16a, 16b, 16c, and the three ascending intersection points 17a, 17b, 17c. In the same manner, both have a longitude interval of 120 °. Further, the inclination angles formed by the quasi-stationary satellite orbits 13a, 13b, and 13c and the equator plane are all the same value.

Next, the operation will be described.
In FIG. 3, three satellites 14a, 14b, and 14c provide communication services to the earth 11 while orbiting quasi-stationary satellite orbits 13a, 13b, and 13c, respectively. When the communication service by the three satellites 14a, 14b, and 14c is performed without any problem, the standby satellite 14d is in standby orbit with a satellite cycle of about 24 hours to cope with the failure of the three satellites 14a, 14b, and 14c. No. 12 does not perform communication services.

  When one of the three satellites 14a, 14b, 14c fails, the standby satellite 14d orbiting the standby orbit 12 moves from the standby orbit 12 to the quasi-stationary satellite orbit where the failed satellite occurred. A communication service is provided to the earth 11 in place of the failed satellite.

  Since the standby satellite 14d is arranged in the standby orbit 12 (that is, the geostationary satellite orbit existing on the equator plane) from the three quasi-stationary satellite orbits 13a, 13b, and 13c with the same inclination angle and orbital altitude, The orbit correction amount can be made equal for any orbit, and any failure of the three satellites 14a, 14b, and 14c can be handled equally.

  A specific input operation of the spare satellite 14d will be described. Here, a case where a failure occurs in the satellite 14a among the three satellites 14a, 14b, and 14c is considered. In order to provide a communication service in place of the failed satellite 14a, the standby satellite 14d is introduced into the quasi-stationary satellite orbit 13a. As a point on the quasi-stationary satellite orbit 13a (hereinafter referred to as an orbit correction point) to which the spare satellite 14d is introduced, for example, the ascending intersection point 17a is selected.

  When the standby satellite 14d reaches the ascending intersection 17a from the point where the orbit correction on the standby orbit 12 is started, the same orbital motion parameter as that of the failed satellite 14a is given, and from the ascending intersection 17a on the quasi-stationary satellite orbit 13a. Start a circular motion. Then, instead of the failed satellite 14a, the standby satellite 14d starts a communication service and provides a communication service to the earth 11 together with other non-failed satellites 14b and 14c while moving around the quasi-stationary satellite orbit 13a. To do.

  When the spare satellite 14d is turned on, it is turned on in consideration of prevention of accidents such as a collision between the failed satellite 14a and the spare satellite 14d and the operational relationship with other non-failed satellites 14b and 14c. The Further, in order not to disturb the communication service of the spare satellite 14d, appropriate measures are taken for the failed satellite 14a depending on the failure status, such as suspension of the communication function or exclusion from the quasi-stationary satellite orbit 13a. .

Here, the operation of inserting the standby satellite 14d when the satellite 14a has failed has been described. However, even when the satellites 14b and 14c have failed, the standby satellite 14d can be input by a similar operation, and three quasi It can be seen that the orbit correction amounts for the geostationary satellite orbits 13a, 13b and 13c are equal.
In addition, as an example, the ascending intersection 17a is selected as the trajectory correction point at the time of input, but the trajectory correction point is selected according to the situation of the entire system.

  As described above, according to the second embodiment, the standby orbit 12 in which the standby satellite 14d orbits is arranged from the three quasi-stationary satellite orbits 13a, 13b, and 13c to the stationary satellite orbit having the same inclination angle and the same orbital height. As a result, the orbit correction amount when the standby satellite 14d is inserted is equal for any of the three quasi-stationary satellite orbits 13a, 13b, and 13c, resulting in failure of the three satellites 14a, 14b, and 14c. On the other hand, it is possible to obtain an effect that the standby satellite 14d can be brought into an optimum standby state.

In the second embodiment, the case of three quasi-stationary satellite orbits 13a, 13b, and 13c has been described. However, the number of quasi-stationary satellite orbits is not limited to this, and a plurality of quasi-stationary satellite orbits are used. All cases can be handled.
In the second embodiment, the number of spare satellites 14d is one. However, the number of spare satellites 14d is not limited to this, and a plurality of spare satellites 14d are arranged in a geostationary satellite orbit as standby orbit 12. May be arranged.

Embodiment 3 FIG.
In the second embodiment, a case where a geostationary satellite orbit is used as the standby orbit 12 in which the standby satellite 14d orbits is shown. In the third embodiment, a case will be described in which the orbit altitude of the standby orbit of the standby satellite 14d is made equal to the altitudes of the three peripheral points 15a, 15b and 15c, and the standby orbit is arranged on the equator plane.

FIG. 4 is a diagram showing the configuration of a satellite orbit according to Embodiment 3 of the present invention. In FIG. 4, reference numeral 22 denotes a standby orbit in which a spare satellite 14d provided for coping with a failure of three satellites 14a, 14b, and 14c orbits. The orbital height of the standby orbit 22 in the third embodiment is as follows. It is equal to the altitude of the three peripheral points 15a, 15b, 15c, and the standby track 22 is arranged on the equator plane.
Other configurations are denoted by the same reference numerals as those in FIG. 3 in the second embodiment, and description thereof is omitted.

Next, the operation will be described.
In FIG. 4, three satellites 14a, 14b, and 14c provide communication services to the earth 11 while orbiting quasi-stationary satellite orbits 13a, 13b, and 13c, respectively. When the communication service by the three satellites 14a, 14b, and 14c is performed without any problem, the standby satellite 14d can receive an arbitrary satellite period (i.e., period) to cope with the failure of the three satellites 14a, 14b, and 14c. A satellite cycle corresponding to the altitude of the point) orbits the standby orbit 22, and no communication service is performed.

  When one of the three satellites 14a, 14b, 14c fails, the standby satellite 14d orbiting the standby orbit 22 moves from the standby orbit 22 to the quasi-stationary satellite orbit where the failed satellite occurred. A communication service is provided to the earth 11 in place of the failed satellite.

  Since the standby satellite 14d is arranged in the standby orbit 22 having the same inclination angle and the same orbit altitude from the three quasi-stationary satellite orbits 13a, 13b, and 13c, the orbit correction amount at the time of introduction is the same for any orbit. It is possible to cope with any failure of the three satellites 14a, 14b, and 14c.

  A specific input operation of the spare satellite 14d will be described. Here, a case where a failure occurs in the satellite 14a among the three satellites 14a, 14b, and 14c is considered. In order to provide a communication service in place of the failed satellite 14a, the standby satellite 14d is introduced into the quasi-stationary satellite orbit 13a. Perige point 15a is selected as the orbit correction point where spare satellite 14d is introduced.

  When the standby satellite 14d reaches the peripheral point 15a from the point where the orbit correction on the standby orbit 22 is started, the same orbital motion parameters as the failed satellite 14a are given, and the preliminary satellite 14d is on the quasi-stationary satellite orbit 13a. Start a circular motion. Then, instead of the failed satellite 14a, the standby satellite 14d starts a communication service and provides a communication service to the earth 11 together with other non-failed satellites 14b and 14c while moving around the quasi-stationary satellite orbit 13a. To do.

  When the standby satellite 14d is inserted, as in the second embodiment, an accident such as a collision between the failed satellite 14a and the standby satellite 14d is prevented, and the operation with the other non-failed satellites 14b and 14c. It is input after considering the relationship. Further, in order not to disturb the communication service of the spare satellite 14d, appropriate measures are taken for the failed satellite 14a depending on the failure status, such as suspension of the communication function or exclusion from the quasi-stationary satellite orbit 13a. .

  Here, the operation of inserting the standby satellite 14d when the satellite 14a has failed has been described. However, even when the satellites 14b and 14c have failed, the standby satellite 14d can be input by a similar operation, and three quasi It can be seen that the orbit correction amounts for the geostationary satellite orbits 13a, 13b and 13c are equal.

  In the third embodiment, since the standby satellite 14d orbits the standby orbit 22 at the satellite period corresponding to the altitude of the three peripheral points 15a, 15b, 15c, the three satellites 14a, 14b, 14c and The relative positional relationship with the satellite 14d changes with time. Therefore, according to the satellite period of the three satellites 14a, 14b, and 14c and the satellite period of the standby satellite 14d, the auxiliary satellite 14d is orbited appropriately and the timing of insertion is measured, so that the three satellites 14a, 14b, For any case of 14c failure, three peripheral points 15a, 15b, 15c are selected as trajectory correction points.

  Further, the orbital altitude of the standby orbit 22 is equal to the altitudes of the three perigee points 15a, 15b and 15c which are lower than the orbital altitude of about 36000km of the geostationary satellite orbit, so that the three perigee points where the altitude from the surface of the earth 11 is the lowest. The spare satellite 14d may be launched to an altitude of 15a, 15b, 15c, and the fuel required when the spare satellite 14d is thrown from the earth 11 into the standby orbit 22 can be saved.

  Further, since the standby orbit 22 of the standby satellite 14d is different from the geostationary satellite orbit, when the standby satellite 14d orbits the standby orbit 22, it is different from the existing geostationary satellite orbiting the geostationary satellite orbit. Interference does not occur, and the effect that the standby satellite 14d can be arranged in the standby orbit 22 without being restricted by the geostationary satellite on the geostationary satellite orbit is obtained.

  As described above, according to the third embodiment, the orbit altitude of the standby orbit 22 of the standby satellite 14d is made equal to the altitudes of the three peripheral points 15a, 15b and 15c, and the standby orbit 22 is arranged on the equator plane. Thus, as in the second embodiment, the orbit correction amount when the standby satellite 14d is inserted into any of the three quasi-stationary satellite orbits 13a, 13b, and 13c becomes equal, and the three satellites For the failure of 14a, 14b, 14c, an effect is obtained that the standby satellite 14d can be put in an optimum standby state.

  Further, since the orbital altitude of the standby orbit 22 is equal to the altitudes of the three peripheral points 15a, 15b and 15c which are lower than the orbital altitude of about 36000km of the geostationary satellite orbit, the standby satellite 14d is at an altitude of the three peripheral points 15a, 15b and 15c. Thus, it is possible to save fuel when the standby satellite 14d is thrown from the earth 11 into the standby orbit 22.

  Further, since the standby orbit 22 of the standby satellite 14d is different from the geostationary satellite orbit, when the standby satellite 14d orbits the standby orbit 22, it is different from the existing geostationary satellite orbiting the geostationary satellite orbit. Interference does not occur, and the effect that the standby satellite 14d can be arranged in the standby orbit 22 without being restricted by the existing geostationary satellite on the geostationary satellite orbit (not shown) is obtained.

In the third embodiment, the case of three quasi-stationary satellite orbits 13a, 13b, and 13c has been described. However, the number of quasi-stationary satellite orbits is not limited to this, and a plurality of quasi-stationary satellite orbits are used. All cases can be handled.
In the third embodiment, the number of spare satellites 14d is one. However, the number of spare satellites 14d is not limited to this, and a plurality of spare satellites 14d are arranged in standby orbit 22. May be.

Embodiment 4 FIG.
FIG. 5 is a diagram for explaining the relative relationship between the satellite and the earth according to the fourth embodiment of the present invention, and FIG. 6 is a diagram for explaining the attitude control of the satellite according to the fourth embodiment of the present invention.
In FIG. 5, 11 is the earth on which the communication service is provided from the satellite 24, 24 is a satellite orbiting the quasi-stationary satellite orbit on the quasi-stationary satellite orbit plane 34, and 24c is illustrated for explaining the attitude direction of the satellite. The satellite attitude axis, 31 is the earth axis of the earth 11, 32 is the ecliptic plane of the earth 11, and the sun (not shown) exists in the right direction of FIG. 33 is the equator plane of the earth 11, and 34 is a quasi-stationary satellite orbit plane in which a quasi-stationary satellite orbit in which the satellite 24 orbits exists.

  In FIG. 6, 24 a is a deployable large planar solar cell light receiving surface fixedly provided on the satellite 24, and 24 b is a directional antenna provided on the satellite 24. The desired direction of the antenna beam is predicted from the shape of the geostationary satellite orbit, and the desired antenna beam is obtained by a mechanical drive system in which the antenna operates or an electronic scanning system using a phased array antenna, independently of the main body of the satellite 24. It can be controlled in the direction.

  Furthermore, in FIG. 6, 24c is a satellite attitude axis shown for explaining the attitude direction of the satellite, 24x and 24y are a first attitude control axis and a second attitude control axis for controlling the attitude of the satellite 24, respectively. is there. The first attitude control axis 24x and the second attitude control axis 24y are a roll axis and a pitch axis, respectively, of three axes orthogonal to each other in the three-axis stabilization method as a conventional satellite attitude control method.

  In general, in the conventional satellite attitude control method using the three-axis stabilization method, the roll axis is set in the satellite traveling direction in the satellite orbital plane, the pitch axis is set in the direction perpendicular to the satellite orbital plane, and the third axis (not shown) orthogonal to the roll axis and the pitch axis. The yaw axis that is the axis of the quasi-stationary satellite orbital plane 34 is set to the earth direction. These three axes are controlled by rotation of a wheel (not shown) provided in the satellite 24, injection of a thruster, and the like, and function to stabilize the attitude of the satellite without rotating the satellite body.

  The satellite attitude axis 24c is substantially collinear with the second attitude control axis 24y. The solar cell light receiving surface 24a is fixedly installed on the satellite 24 so as to be substantially parallel to a surface including the first attitude control axis 24x and the second attitude control axis 24y.

Next, the operation will be described.
The attitude of the satellite 24 orbiting the quasi-stationary satellite orbit on the quasi-stationary satellite orbital plane 34 is disturbed by disturbance due to the gravity inclination of the earth 11 or the solar wind. Further, as the position of the quasi-stationary satellite orbit of the satellite 24 changes, the incident angle from sunlight (not shown) with respect to the solar cell light receiving surface 24a changes. In order to cope with the disturbance and the change in the incident angle, the attitude of the satellite 24 is controlled by two-axis control using the first attitude control axis 24x and the second attitude control axis 24y.

  That is, the sun from sunlight (not shown) is maintained by attitude control using the second attitude control axis 24y while maintaining the satellite attitude axis 24c perpendicular to the ecliptic plane 32 by attitude control using the first attitude control axis 24x. The incident angle with respect to the battery light receiving surface 24a is kept at 90 °. In this way, by simple biaxial attitude control using the first attitude control axis 24x and the second attitude control axis 24y, the attitude of the satellite 24 is kept perpendicular to the ecliptic plane 32 and from sunlight. By maintaining the incident angle with respect to the solar cell light receiving surface 24a at 90 °, the attitude of the satellite 24 is stabilized against disturbance due to the gravitational inclination or solar wind of the earth 11, and the power generation efficiency of the solar cell provided in the satellite 24 is improved. Maximum stability can be achieved.

  At this time, since the attitude of the satellite 24 with respect to the earth 11 varies depending on the position of the satellite 24 in the quasi-stationary satellite orbit on the quasi-stationary satellite orbital plane 34, the relative relationship between the directional antenna 24 b provided on the satellite 24 and the earth 11. Also changes. However, since the directional antenna 24b can control the antenna beam in a desired direction, the relative relationship between the directional antenna 24b and the earth 11 is predicted from the shape of the quasi-stationary satellite orbit on the quasi-stationary satellite orbit plane 34, and the directivity A communication service can be provided to the earth 11 by controlling the antenna beam of the antenna 24b in a desired direction.

  As described above, according to the fourth embodiment, the solar cell light receiving surface 24a is fixed to the satellite 24 so as to be substantially parallel to the surface including the first attitude control axis 24x and the second attitude control axis 24y. The satellite attitude axis 24c of the satellite 24 is kept perpendicular to the ecliptic plane 32 by the first attitude control axis 24x, and the solar cell light receiving surface 24a is adjusted to sunlight by the second attitude control axis 24y. As a result of the pairing, the effect of stabilizing the attitude of the satellite 24 against the disturbance due to the gravitational inclination of the earth 11 or the solar wind and the power generation efficiency of the solar cell can be maximized by simple two-axis control. Is obtained.

It is a flowchart which shows the satellite orbit structure method by Embodiment 1 of this invention. 2 is a world map in which a quasi-stationary satellite orbit according to the first embodiment is projected. It is a figure which shows the structure of the satellite orbit by Embodiment 2 of this invention. It is a figure which shows the structure of the satellite orbit by Embodiment 3 of this invention. It is a figure explaining the relative relationship between the satellite and the earth by Embodiment 4 of this invention. It is a figure explaining the attitude | position control of the satellite by Embodiment 4 of this invention. It is a figure which shows the structure of the satellite orbit using the conventional quasi-stationary satellite orbit.

Explanation of symbols

  11 Earth, 12, 22 Standby orbit, 13a, 13b, 13c Quasi-stationary satellite orbit, 14a, 14b, 14c Satellite, 14d Spare satellite, 15a, 15b, 15c Perigee point, 16a, 16b, 16c Apogee point, 17a, 17b, 17c Ascending intersection, 24 satellites, 24a solar cell light receiving surface, 24b directional antenna, 24c satellite attitude axis, 24x first attitude control axis, 24y second attitude control axis, 31 earth axis, 32 ecliptic plane, 33 equatorial plane, 34 Quasi-stationary satellite orbit plane, 100 quasi-stationary satellite orbit, 200 apogee projection point.

Claims (1)

In a satellite attitude control method for providing a communication service to the earth,
A solar cell light receiving surface for receiving sunlight is provided so as to be substantially parallel to a surface including a first posture control axis and a second posture control axis that are orthogonal to each other, and the second posture is controlled by the first posture control axis. The attitude control axis of the solar cell is kept perpendicular to the ecliptic plane, the orbital plane of the earth moving around the sun, and the solar cell light-receiving surface is opposed to sunlight by the second attitude control axis. A satellite attitude control method characterized by

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