JP2006117180A - Information processor, orbital control device, artificial satellite, communication device and communication method - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To make a satellite direction seen from an earth core when two satellites fly near an orbital plane intersection line approaching while avoiding collision of the satellites regarding a plurality of non-stationary satellites having crossed satellite orbital. <P>SOLUTION: An amount of difference of a distance from the earth core of two satellites is set and argument of perigee so as to satisfy the amount is determined at the time when respective two satellites of the first orbital satellite, the second satellite and the third satellite can be seen in the approximately same direction when seeing from an earth direction. Eccentricity is determined such that a flight time of the satellite from one orbital plane intersection line to the subsequent orbital plane intersection line becomes 1/3 time of an orbital period. Reference orbital of the first orbital satellite, the orbital of the second satellite and the orbital of the third satellite are set from the argument of perigee and the eccentricity and a retaining range is set around a satellite material point the reference orbital. The retaining range of the respective satellites are overlapped and can pass on the respective orbital surface intersection lines. <P>COPYRIGHT: (C)2006,JPO&NCIPI

Description

  The present invention relates to a reference orbit setting technique for a plurality of non-geostationary satellites whose satellite orbits intersect each other, an orbit control technique for the non-geostationary satellite, a communication technique for communicating with the non-geostationary satellite, and the like.

  As background art, there are techniques disclosed in Japanese Patent Laid-Open Nos. 8-223100 and 8-3331033. These disclose technologies for communicating between a plurality of satellites such as Iridium satellites (66) and the ground, and two satellites meet for a plurality of non-stationary satellites whose satellite orbits intersect each other. A technique for reducing the inter-satellite distance at the time of performing is disclosed.

JP-A-8-223100 JP-A-8-331033

In the case of an iridium satellite or the like that is the subject of the prior art of Patent Document 1 and Patent Document 2 described above, the directivity of the RF signal transmitted from the ground is low, so the communication between two satellites when transferring from one satellite to another There is no major restriction on the distance, and there is no need to bring the satellites close to each other. In addition, it is necessary to see multiple satellites from the ground, but it seems that there are no further restrictions, and the orbital arrangement is considered to be sufficiently flexible.
For this reason, the prior arts of Patent Document 1 and Patent Document 2 described above cannot be applied to satellite communication in which it is necessary to reduce the distance between satellites when communication is transferred from one satellite to another.

  An object of the present invention is to reduce the distance between satellites when two satellites meet while avoiding a collision between the satellites of a plurality of non-stationary satellites whose satellite orbits intersect each other.

An information processing apparatus according to the present invention includes:
Information processing for determining a first satellite orbit that the first artificial satellite orbits, and a second satellite orbit of the second artificial satellite that intersects the first satellite orbit and the satellite direction viewed from the earth center A device,
Orbit for setting the difference between the distance of the first artificial satellite from the earth center and the distance of the second artificial satellite from the earth center at the time of transfer of communication from the first artificial satellite to the second artificial satellite Difference amount setting section,
The orbit holding range of the first artificial satellite and the orbit holding range of the second artificial satellite overlap each other when viewed from the center of the earth based on the near point argument determined based on the difference amount set by the orbit difference amount setting unit. And a reference trajectory setting unit for determining a passing point.

  According to the present invention, it is possible to reduce the direction between each satellite as seen from the ground at the time of handover for switching communication from one satellite to another, and only one antenna for signal transmission on the ground is required. Transmission signal switching control is also facilitated.

Embodiment 1 FIG.
In this embodiment, an information processing apparatus that determines reference orbits of a plurality of non-stationary satellites whose satellite orbits intersect each other will be described.
Here, the reference orbit is a reference orbit that is the target of the orbit of the satellite. In order to ensure that each of the three satellites is in a specific position at a specific time, It means the reference trajectory to be targeted.
Embodiments of the present invention will be described with reference to the drawings.

FIG. 1 is a diagram illustrating a satellite orbit example of a quasi-zenith satellite that is an example of a non-geostationary satellite that is a target of the information processing apparatus according to the first embodiment.
FIG. 1 shows a situation in which three quasi-zenith satellite orbits and the movement of the satellite in an inertial space centered on the earth are viewed from the celestial North Pole. As shown in FIG. 1, the quasi-zenith satellite is three artificial satellites that orbit the first orbit 10, the second orbit 20, and the third orbit 30. The satellites flying in the first orbit 10 and the satellites flying in the second orbit 20 are simultaneously present on the intersection line A of their orbital planes at a certain time, and the satellite directions seen from the center of the earth are the same direction. The satellites flying in the orbit 20 and the satellites flying in the third orbit 30 are simultaneously present on the intersection line B of their orbital planes at a certain time, and the satellite directions viewed from the center of the earth are the same direction. The satellites flying in the first orbit 10 and the satellites flying in the first orbit 10 simultaneously exist on the intersection line C of their orbital planes at a certain time, and the satellite directions viewed from the center of the earth are the same. When two satellites exist at the same time on the intersection of their orbital planes at the same time and the satellite directions viewed from the center of the earth are the same, this is called direction association. Lines (intersection line A, intersection line B, intersection line C) are called direction meeting lines. The quasi-zenith satellite starts from one direction meeting line and arrives at the next direction meeting line, and it takes about 8 hours. The satellite in the third orbit and the satellite in the first orbit meet at the direction meeting line C. After about 8 hours, the satellites in the first orbit and the satellites in the second orbit meet at the direction meeting line A. In addition, after about 16 hours, the satellites in the second orbit and the satellites in the third orbit meet in the direction meeting line B. Then, after about 24 hours, that is, one day later, the satellite in the third orbit and the satellite in the first orbit meet in the direction meeting line C. This repetition is continued.
From the direction meeting line C to the direction meeting line A, the satellite in the first orbit that orbits the first orbit 10 can communicate with the ground station, and the second orbit that orbits the second orbit 20 on the direction meeting line A. The direction meeting with the satellite and after the direction meeting line A, the satellite communicating with the ground station is switched to the satellite in the second orbit. The satellite in the second orbit is communicable with the ground station from the direction meeting line A to the direction meeting line B, and is directionally associated with the satellite in the third orbit around the third orbit 30 on the direction meeting line B. After the meeting line B, the satellite communicating with the ground station is switched to the satellite in the third orbit. The satellite in the third orbit can communicate with the ground station from the direction meeting line B to the direction meeting line C, and directionally meets with the satellite in the first orbit on the direction meeting line C. The satellite for communication is switched to the satellite in the first orbit.

The orbital elements of each satellite are characterized as follows.
The orbital length radius is common to all three orbits and is about 42164 km, that is, a period of about 23 hours and 56 minutes, the eccentricity is common to all three orbits between about 0.099 and about 0.101, and the orbit inclination angle is common to all three orbits. The ascending intersection red longitude is set at an interval of 120 degrees, the peripoint argument is common to all three orbits between about 260 degrees and about 280 degrees, and the average near point separation angle is 240 degrees, 120 degrees, 0 at the same time. Degree. The satellite direction meeting is characterized in that the trajectories of the respective orbits physically intersect the direction meeting line, and at the time of the intersection, the two satellites occupy the same direction as viewed from the earth center. Providing this feature is based on setting a specific eccentricity as a function of the perimeter argument. When the perimeter argument takes a common value between about 260 degrees and about 280 degrees, the eccentricity should take a value between about 0.099 and about 0.101 as a function of that value. If the eccentricity takes a different value other than this value, the trajectories of the orbits physically intersect each other at the direction meeting line, but two predetermined satellites are simultaneously on the direction meeting line and have the same direction as seen from the center of the earth. It does not happen. This eccentricity is mathematically determined from other defined orbital elements.

Here, the process of deriving the eccentricity as a function of the near-field argument for two adjacent satellites to directionally meet on the orbital direction meeting line and three satellites to sequentially meet on the three orbital direction meeting lines is described below. Shown in
FIG. 8 shows the situation of the intersection of the equatorial plane and each of the first and second orbits using the spherical trigonometry. The equator plane is indicated by a disk W, and its center point is indicated by O. Point O is the center of the earth. A hemispherical surface T having the same center and the same diameter is provided on the disk W, and its apex is A (points A, B, and C shown in FIGS. 8 to 11 are direction meeting lines A and B shown in FIG. , Different from C). FIG. 8 is a view of the equatorial plane W and the hemispherical surface T as viewed from directly above, so that the points O and A appear to occupy the same position.
An intersection line between the track surface of the first track and the hemispherical surface T is indicated by U. U is a semicircular arc, and the intersection with the equator plane W is represented by F and G. The line segment FG passes through the center O. Let C be a point where a straight line included in the first orbital plane perpendicular to the line segment FG from the point O intersects the arc U. Apparently, the point C is the highest point of the arc U, that is, the point on the arc U farthest from the equator plane W.
An intersection line between the orbital surface of the second orbit and the hemispherical surface T is indicated by V. V is an arc of a semicircle, and the intersection with the equator plane W is represented by H and I. The line segment HI passes through the center O. Let D be the point where a straight line from the point O perpendicular to the line segment HI and included in the second orbital plane intersects the arc V. Apparently, the point D is the highest point of the arc V, that is, the point on the arc V farthest from the equator plane W.
A surface obtained by rotating the first track surface by 120 degrees counterclockwise is the second track surface. The rotation of this surface is made around an axis perpendicular to the equator plane W, that is, a straight line OA. The orbit inclination angles of the first and second orbits are the angles formed by the orbital angular momentum vector (perpendicular to the orbital plane) of each orbital plane and the north vertical axis of the equatorial plane W, that is, the straight line OA. The angle is represented by i. In a specific example, i = 45 degrees.
Here, let B be the intersection of arc U and arc V. Since the trajectory inclination angles of the two trajectories are the same, the arcs U and V and the points C and D are in symmetrical positions with reference to the arc AB in FIG. This can also be seen from the fact that the spherical triangle BIG is an isosceles triangle. Because of this symmetry, it can be seen that the spherical angle BAD formed by the plane OBA and the plane ODA is equal to the spherical angle BAC formed by the plane OBA and the plane OCA, and the sum thereof is 120 degrees. This can also be seen from the fact that the point C is rotated 120 degrees around the point A and the point D is obtained by rotating the arc U by 120 degrees. Accordingly, the spherical angles BAD and BAC are 60 degrees.
Further, since the plane OCA is perpendicular to the line segment FG, it is perpendicular to the plane OCF, that is, OCB. This indicates that the spherical angle ACB is a right angle. Similarly, since the plane ODA is perpendicular to the line segment HI, the plane ODA is perpendicular to the plane ODH, that is, ODB. This indicates that the spherical angle ADB is a right angle.

FIG. 9 is a diagram showing the geometric relationship shown in FIG. 8 for the first trajectory and the second trajectory. In FIG. 9, elements W and O represented by the same symbols as in FIG. 8 indicate the same as in FIG.
U ′ is a closed curve indicating the first trajectory, and an ascending intersection crossing the equator plane W from the south to the north is G ′, and a descending intersection crossing the equator plane W from the north to the south is F ′. The point G ′ is on the straight line OG in FIG. 8, and the point F ′ is on the straight line OF in FIG. When the near point argument of the first trajectory is 270 degrees, the angle from the ascending intersection point G ′ to the far point C ′ of the trajectory is 90 degrees, and the point C ′ is on the straight line OC in FIG.
V ′ is a closed curve indicating the second trajectory, and an ascending intersection that crosses the equator plane W from south to north is H ′, and a descending intersection that crosses the equator plane W from north to south is I ′. The point H ′ is on the straight line OH in FIG. 8, and the point I ′ is on the straight line OI in FIG. When the near point argument of the second trajectory is 270 degrees, the angle from the ascending intersection H ′ to the far point D ′ of the trajectory is 90 degrees, and the point D ′ is on the straight line OD in FIG.
When the upside line connecting the far point C ′ and the near point of the first trajectory U ′ is J, when the near point argument of the first trajectory is 270 degrees, the straight line J is perpendicular to the straight line F′G ′ and the straight line FG. Since the straight line J is an upside line, the trajectory U ′ is symmetric with respect to the straight line J. Similarly, assuming that the upside line connecting the far point D ′ and the near point of the second trajectory V ′ is K, when the near point argument of the second trajectory is 270 degrees, the straight line K is a straight line H′I ′ and a straight line HI. Since the straight line J is an upside line, the trajectory V ′ is symmetric with respect to the straight line K. Since the first trajectory and the second trajectory are in a relationship rotated by 120 degrees around the straight line OA, when both the near-point argument of the first trajectory and the near-point argument of the second trajectory are 270 degrees, these trajectories The first trajectory U ′ and the second trajectory V ′ always cross each other due to the symmetry of the above, and if the intersection on the upper side (north side) of the equator plane W is B ′, the point B ′ is the first trajectory. Are on a straight line intersecting the orbital plane of the second and the orbital plane of the second orbit. This means that the point B ′ is on the straight line OB in FIG. The angle B′OD ′ and the angle B′OC ′ are equal, and the angle is a.
In addition, when both the near point of the first trajectory and the near point argument of the second trajectory are other than 270 degrees, the first trajectory U ′ and the second trajectory V ′ do not intersect each other, but each trajectory. Always intersects the straight line OB in FIG. This is because the straight line OB is included in each track surface.

FIG. 10 is a diagram in which a geometrical relationship is extracted with the spherical triangle ABD in FIG. 8 as the center. 10, elements A, B, D, U, and V represented by the same symbols as in FIG. 8 indicate the same as in FIG. Point E indicates the point at which the orbital angular momentum vector of the second orbit, that is, the straight line perpendicular to the orbital plane of the second orbit intersects the hemispherical surface T in FIG. Since the orbit inclination angle i is defined as an angle formed by a straight line perpendicular to the equator plane W and the orbital angular momentum vector, the angle formed by the side AE on the spherical surface is the orbit inclination angle i. Further, since the angle of the side DE is 90 degrees, if the angle of the side AD is b, the angle i of the side AE is 90 deg−b. Further, as shown in the description of FIG. 8 above, the angle BAD is 60 degrees and the angle ADB is 90 degrees. The angle of the side BD is a shown in FIG.
The spherical triangle ABD is a right triangle, and the following equation is obtained from the right spherical triangle formula.
When the angle A is the angle BAD, that is, 60 degrees, the angle a is the angle of the side BD, and the angle b is the angle of the side AD,
tan A = (tan a) / (sin b)
Therefore,
tan a = tan A × sin b
= (√3) × sin (90 deg−i)
= (√3) x cos i
As a specific example, when i is 45 degrees, cos i = 1 / √2, so
It can be seen that tan a = √ (3/2) and a = 50.768848 degrees.

FIG. 11 is a diagram showing direction association lines with the near point, the far point, the ascending intersection, the descending intersection, and the orbit of the first satellite when the near point argument is 270 degrees with respect to the second orbit. This geometric relationship holds true for the first trajectory and the third trajectory. This is because the orbital planes of the three orbits are separated by 120 degrees around the equator plane perpendicularly, and the orbital length radius, that is, the orbital period, the eccentricity, the orbital inclination angle, and the near point argument are equal. 11, elements O, B ′, D ′, H ′, I ′ and a represented by the same symbols as in FIGS. 8 and 9 are the same as those in FIGS. Point L indicates the near point of the trajectory, and line OM indicates a direction meeting line with the next or third trajectory. The point M is symmetric with the point B ′ with respect to the upside line D′ L. The angle f is an angle formed from the near point L to the point B ′, and is an angle called a near point separation angle.
Regardless of whether or not the perimeter argument is 270 degrees, three satellites sequentially meet in the direction meeting line when viewed in the second orbit, the first orbiting satellite and the second When the orbiting satellite is directionally met and then separated from the first orbiting satellite and the second orbiting satellite reaches the direction meeting line OM via the far point D ′, the third orbiting satellite is on the line OM. The event that appears in the direction and meets direction is repeated for every three satellites in turn. Therefore, in order for this event to occur, the satellite flight time from the direction meeting line OB ′ to the next direction meeting line OM may be one third of the orbital period. On the other hand, if the satellite flight time is different from one third of the orbital period, directional meetings will not occur sequentially. This is because the orbits of the three satellites have the same period of 23 hours and 56 minutes as the geostationary orbits, and the locus of the point immediately below the satellites is constant on the ground surface and does not change.
In FIG. 11, when the near point argument is 270 degrees, the required time from the near point L to the far point D ′ is equal to the required time from the far point D ′ to the near point L and is a half of the orbital period. The required time from point B ′ to point D ′ is equal to the required time from point D ′ to point M and is one-sixth of the orbital period. Further, the required time from the point L to the point B ′ is equal to the required time from the point M to the point L and is one third of the orbital period. From this, a typical example of the arrangement of the three satellites for direction meeting is a case where the satellites are at point B ′, point M and point L of each orbit at a certain point in time.
The trajectory length radius, eccentricity, trajectory tilt angle, and peripheral arguments are the same for the three trajectories, and the trajectory length radius and trajectory tilt angle are given, and the perimeter arguments are 270 degrees. The trajectory planes of the three trajectories are perpendicular to the equatorial plane. In order for the time required from the direction meeting line OB ′ to the next direction meeting line OM to be one third of the orbital period when the surroundings are 120 degrees apart, that is, the ascending intersection is 120 degrees apart. The required time from the near point L to the direction meeting line OB ′ only needs to be one third of the orbital period, and the eccentricity at which it is established becomes a certain fixed value. The method for obtaining the eccentricity is shown below.
Since the required time from the near point L to the point B ′ is one third of the orbital period, the average near point separation angle m of the point B ′ is
m = (2 × π) / 3
It is. This is because the orbital period expressed by the average near point separation angle is 2 × π.
As a relational expression of the average near point separation angle m, the eccentric near point separation angle g, and the perfect near point separation angle f, the following equation is well known.
tan (f / 2)
= √ ((1 + e) / (1-e)) × tan (g / 2)
m = g−e × sing
Here, e is an eccentricity to be obtained.
The near point separation angle f is expressed as follows using the angle a.
f = π-a
In summary, the angle a is obtained from the orbit inclination angle i, the near point separation angle f is obtained from a, and e is obtained from f and m. As can be seen from these equations, the equation for obtaining e from f and m is a transcendental equation and cannot be obtained directly analytically but by numerical calculation such as an iterative method.
As a specific example, when i = 45 degrees, a = 50.768848 degrees, and it can be understood that e = 0.099231 by numerical calculation.

The above is a calculation of the eccentricity for the three satellites to sequentially directionally meet when the perimeter argument is 270 degrees. Next, in the case where the perimeter argument is different from 270 degrees, the eccentricity for the three satellites to sequentially meet in direction is obtained.
FIG. 12 is a diagram showing the near-point, far-point, direction meeting line with the first satellite orbit, and direction meeting line with the third satellite orbit when the near-point argument is different from 270 degrees with respect to the second orbit. is there.
This geometric relationship holds true for the first trajectory and the third trajectory. This is because the orbital planes of the three orbits are separated by 120 degrees around the equator plane perpendicularly, and the orbital length radius, that is, the orbital period, the eccentricity, the orbital inclination angle, and the near point argument are equal. In FIG. 12, elements O, B ′, M and a represented by the same symbols as those in FIGS. 8, 9 and 11 are the same as those in FIGS. Point L ′ indicates the near point of the trajectory, point D ″ indicates the far point of the trajectory, point B ″ indicates the trajectory position advanced by angle a clockwise from point D ″, and point M ′ is counterclockwise from point D ″. The trajectory position advanced by an angle a around is shown. The point B ′ is a point on the second orbit where the line OB ′ becomes a direction meeting line with the first orbit, and the point M is a direction meeting line with the next or third orbit. Is a point on the second orbit. The angle Δω is an angle obtained by subtracting the near point argument of the second trajectory from 270 degrees, and FIG. 12 shows a case of a positive value.
As already mentioned, in order for three satellites to successively meet at the direction meeting line, the satellite flight time from the direction meeting line to the next direction meeting line is 1/3 of the orbital period for each satellite. I need it. This indicates that the satellite flight time from point B ′ to point M is one third of the orbital period. The following formula is established according to the case where the perimeter argument already described is 270 degrees.
f ₁ is the near point separation angle of point B ″, f ₂ is the near point separation angle of point B ′, f ₃ is the near point separation angle of point M ′, and f ₄ is the near point separation angle of point M ′. Then, f ₁ = π−a from the geometric relationship shown in FIG.
f ₂ = π−a + Δω
f ₃ = π + a
f ₄ = π + a + Δω
_{g 2, g 4, m 2} , m 4 each _f 2, eccentricity anomaly of _{f 4,} as the average anomaly of _f 2, f _4, satellite flight time from point B 'to point M Because it is one third of the orbital period,
m ₄ −m ₂ = (2 × π) / 3
From the relationship between the near point separation angle, the eccentric point separation point, and the average near point separation angle,
_{tan (f 2/2) =} √ ((1 + e) / (1-e)) × tan (g 2/2)
m ₂ = g ₂ −e × sing _{2
tan (f 4/2) =} √ ((1 + e) / (1-e)) × tan (g 4/2)
m ₄ = g ₄ −e × sing ₄
Here, e is an eccentricity to be obtained. As can be seen from these equations, the equation for obtaining e from the relational expression of f and m is a transcendental equation and cannot be obtained directly analytically but by numerical calculation such as an iterative method.
As a specific example, when i = 45 degrees, since a = 50.768848 degrees, when the perimeter argument is 260 degrees (Δω is 10 degrees), e = 0.100904 and the perimeter argument is 265 degrees (Δω Is 5 degrees), e = 0.099644, and when the perimeter argument is 275 degrees (Δω is −5 degrees), e = 0.099644, and the perimeter argument is 280 degrees (Δω is −10 degrees). E = 0.100904. In addition, when the perimeter argument is 270 degrees, it is understood that e = 0.099231, which is the same as the case described in FIG.

  Thus, the quasi-zenith satellite, which is an example of the non-geostationary satellite that is the target of the information processing apparatus according to the present embodiment, has an ascending intersection of different longitudes of 120 degrees, that is, an orbital plane rotated by 120 degrees. Yes, in three satellite orbits of equal orbit inclination angle (45 degrees), isobaric argument (between about 240 degrees and about 300 degrees) and equal period (almost stationary period, 23 hours 56 minutes, synchronous orbit) Eccentricity common to the orbits of the three satellites is set so that the satellites flying in the orbits of the orbits meet in the direction meeting line of each orbit.

FIG. 2 shows the trajectory on the ground surface at the point directly below the satellite. Here, a case where the perimeter argument is 270 degrees is shown.
The trajectory has the shape of an asymmetrical 8-character, and the intersection of the 8 characters indicates the trajectory of the direction meeting line in FIG. In FIG. 1, three orbits are shown. By setting these orbital elements as described above, the surface trajectory at the point directly below the satellite becomes one figure of eight. When each satellite is located at a far point or near point, the longitude of the point immediately below it is about 135 degrees east longitude, and when the satellite goes from the direction meeting line to the next direction meeting line through the far point, the locus of the point just below is a figure of 8 Starting at the intersection, it returns to the intersection of figure 8 again at 45 degrees north latitude, through a point just below the satellite at about 135 degrees east longitude. During this period (about 8 hours), the satellite stays above Japan, the elevation angle when looking at the satellite from Japan is about 70 degrees or more, and the possibility of entering the shadow of a mountain or building is lower than that of a normal geostationary satellite. Become.
When two satellites directionally meet on the direction meeting line, for example, when a satellite in the third orbit and a satellite in the first orbit directionally meet on the line OC (the satellite to be communicated by the ground station is in the third orbit When switching from a satellite to a satellite in the first orbit, the signal transmitted to the satellite is sent from the antenna of the ground station toward the satellite in the third orbit, and the signal transmitted to the satellite here is the first. Transmission to the satellite in the first orbit is started at a frequency different from that for the satellite in the third orbit so that the satellite is also sent to the satellite in the orbit. When the switching process at the ground station is completed, transmission to the satellite in the third orbit is stopped, and instead, the satellite in the first orbit continues to take charge of communication with the ground.
Here, when the satellite in the third orbit and the satellite in the first orbit are direction-meeting, the ground station simultaneously signals the two satellites with the same antenna for the satellite in the third orbit and the satellite in the first orbit. It is desirable that each direction is close enough to send out
For this reason, in the information processing apparatus according to the present embodiment, when the two satellites are direction-meeting, their directions approach to the extent that signals can be sent to the two satellites simultaneously using the same antenna of the ground station. Thus, the positional relationship of each satellite is determined.

  FIG. 3 is a block diagram illustrating a configuration example of the information processing apparatus 40 according to the present embodiment.

The input unit 401 inputs various instructions from the operator.
The display unit 402 displays the processing result of the information processing apparatus 40 on the display display.

The first holding range setting unit 403 sets the first holding range around the mass that flies in the reference orbit of the first artificial satellite. The first holding range is a range in which the first satellite should stay, and is a rectangle composed of two sides of the orbital traveling direction centered on the mass point flying in the reference orbit and the orbital plane vertical direction.
The orbit of an artificial satellite gradually deviates from the reference orbit due to the effects of gravity deflection due to earth flatness, lunar and solar attraction, and solar radiation pressure, but the holding range defines the limit of deviation from the reference orbit. In the case of deviating, the orbit of the satellite is corrected by injecting an orbit control thruster, that is, a gas jet provided in the satellite.
By always keeping the satellite within this holding range, the communication antenna can be accurately pointed to the specified ground direction, and radio wave interference and physical collision with other satellites can be avoided. become.

  The second holding range setting unit 404 sets the second holding range around the mass point flying in the reference orbit of the second artificial satellite. The second holding range is a range in which the second satellite should stay, and is a rectangle composed of two sides of the trajectory traveling direction centering on the mass point flying in the reference trajectory and the trajectory plane vertical direction.

  The third holding range setting unit 405 sets the third holding range around the mass point that flies in the reference orbit of the third artificial satellite. The third holding range is a range in which the third satellite is to stay, and is a rectangle centering on a mass point that flies in a reference orbit consisting of two sides of the orbit traveling direction and the orbit plane perpendicular direction.

The orbital difference amount setting unit 406 includes a distance from the earth center of the first artificial satellite and a distance from the earth center of the second artificial satellite at the time when the communication is shifted from the first artificial satellite to the second artificial satellite. And the difference between the distance of the second artificial satellite from the earth center and the distance of the third artificial satellite from the earth center at the time of transfer of communication from the second artificial satellite to the third artificial satellite And a difference amount between the distance from the earth center of the third artificial satellite and the distance from the earth center of the first artificial satellite at the time when the communication is transferred from the third artificial satellite to the first artificial satellite. To do. This difference amount is the same for the three artificial satellites, and the three-orbit near point argument is shifted by the same value from 270 degrees as follows. FIG. 12 shows a state in which the near point argument is deviated from 270 degrees, the point B ′ is a point on the second trajectory where the line OB ′ is a direction meeting line with the first trajectory, and the point M is the line OM. It is a point on the second orbit that becomes a direction meeting line with the next or third orbit. Since this geometric relationship holds in the same way in other orbits if the value of the near point argument is the same, the difference in the distance from the earth center of the artificial satellite is the difference between the line segment OB ′ and the line segment OM itself. . Incidentally, since the deviation angle Δω of the near point argument in FIG. 12 is a positive value, the line segment OB ′ is larger than the line segment OM. In addition, when the near point argument is 270 degrees, that is, when the angle Δω is zero, the difference amount is zero, and it can be seen that the absolute value of the difference amount monotonously increases as the near point argument deviates from 270 degrees.
The amount of difference in the distance from the earth center of the artificial satellite is determined from the condition for transition of communication between satellites and the condition for avoiding physical collision. In general, when the amount of difference increases, a time lag of communication that occurs at the time of transition of communication between satellites increases, and communication performance deteriorates. On the other hand, when the difference amount is small, the possibility of physical collision of satellites arises. Therefore, it can be seen that a difference amount satisfying both requirements is necessarily selected.
If the difference amount is selected, the near point argument is shifted from 270 degrees, and the near point argument realizing the specified difference amount is selected. The value of the selected near point argument is used as the near point argument setting value in the reference trajectory setting unit described later.
The following is a specific numerical example of the near point argument obtained from the difference amount. The absolute value of the difference amount is 639 Km, and the absolute value of the angle Δω obtained by subtracting the near point argument from 270 degrees is 5 degrees. Similarly, the absolute value is 4 degrees for 511 Km, 3 degrees for 383 Km, and 255 Km. On the other hand, it is 2 degrees, 1 degree for 128 km, 0.7 degree for 90 km, 0.5 degree for 63 km, 0.3 degree for 40 km, and 0.1 degree for 13 km.

The first reference orbit setting unit 407 sets a reference orbit for the first artificial satellite. The reference orbit refers to the ideal orbit that an artificial satellite that satisfies the mission performed by the payload mounted on the satellite should fly, and it is ideal for the artificial satellite to fly along the mass point that flies in the ideal orbit. State. As mentioned in the holding range setting section above, artificial satellites gradually deviate from the reference orbit due to gravity deflection due to earth flatness, lunar / sun attraction, and solar radiation pressure. Is provided with a holding range, and orbit control is performed so that the satellite does not deviate from the holding range.
A method for setting the reference trajectory will be described later.

The second reference orbit setting unit 408 sets a reference orbit for the second artificial satellite. The reference orbit refers to the ideal orbit that an artificial satellite that satisfies the mission performed by the payload mounted on the satellite should fly, and it is ideal for the artificial satellite to fly along the mass point that flies in the ideal orbit. State. As mentioned in the holding range setting section above, artificial satellites gradually deviate from the reference orbit due to gravity deflection due to earth flatness, lunar / sun attraction, and solar radiation pressure. Is provided with a holding range, and orbit control is performed so that the satellite does not deviate from the holding range.
A method for setting the reference trajectory will be described later.

The third reference orbit setting unit 409 sets a reference orbit for the third artificial satellite. The reference orbit refers to the ideal orbit that an artificial satellite that satisfies the mission performed by the payload mounted on the satellite should fly, and it is ideal for the artificial satellite to fly along the mass point that flies in the ideal orbit. State. As mentioned in the holding range setting section above, artificial satellites gradually deviate from the reference orbit due to gravity deflection due to earth flatness, lunar / sun attraction, and solar radiation pressure. Is provided with a holding range, and orbit control is performed so that the satellite does not deviate from the holding range.
The method for setting the reference trajectory will be described next.

Next, a reference trajectory setting method in the information processing apparatus 40 according to the present embodiment will be described.
As described above, the eccentricity common to the three satellite orbits was derived so that the three satellites would successively meet in the direction meeting line based on the near point argument common to the three satellite orbits. In accordance with this derivation method, the eccentricity at which the three satellites sequentially meet in the direction meeting line with respect to the near point argument set by the orbital difference amount setting unit 406 is calculated and obtained. In this way, the eccentricity and the near point argument as the orbital elements of the three satellites are set. The values set in advance are used as is for the orbital radius, the orbital inclination angle, the ascending intersection longitude, and the average near point separation angle. Note that the average near-point separation angle differs by 120 degrees for every three satellites, and the average motion (orbital average rate) increases in proportion to time with the average motion (orbital average rate) so that the three satellites sequentially meet in the direction meeting line. It is.
The reference orbit setting method is to set the orbital length radius, the eccentricity, the orbital inclination angle, the ascending intersection red longitude, the near-point argument, and the average near-point separation angle that increases with time in this way for each of the three satellites. .

Next, a procedure for setting a reference trajectory of the information processing apparatus 40 according to the present embodiment will be described with reference to the flowchart of FIG.
First, in step S701, a first artificial satellite, a second artificial satellite, and a third artificial satellite are determined. Input unit for determining which of the first orbit satellite, the second orbit satellite, and the third orbit satellite shown in FIG. 1 is the first artificial satellite, the second artificial satellite, and the third artificial satellite. 401 inputs an instruction from the operator.

  Next, in step S702, a first holding range is set. The input unit 401 inputs the value of the first holding range angle from the operator, and the first holding range setting unit 403 sets and stores the input value as the first holding range angle. The first holding range is a range in which the first satellite should stay, and is a rectangle composed of two sides of the orbital traveling direction centered on the mass point flying in the reference orbit and the orbital plane vertical direction. The first holding range angle refers to the angle of each side of the two-sided rectangle.

  Next, in step S703, a second holding range is set. The input unit 401 inputs the value of the second holding range angle from the operator, and the second holding range setting unit 404 sets and stores the input value as the second holding range angle. The second holding range is a range in which the second satellite should stay, and is a rectangle composed of two sides of the trajectory traveling direction centering on the mass point flying in the reference trajectory and the trajectory plane vertical direction. The second holding range angle refers to the angle of each side of the rectangle composed of these two sides.

  Next, in step S704, a third holding range is set. The input unit 401 inputs the value of the third holding range angle from the operator, and the third holding range setting unit 405 sets and stores the input value as the third holding range angle. The third holding range is a range in which the third satellite should stay, and is a rectangle composed of two sides of the trajectory traveling direction centering on the mass point flying in the reference trajectory and the trajectory plane vertical direction. The third holding range angle refers to the angle of each side of the two-sided rectangle.

 Next, in step S705, the orbital difference amount that satisfies both the orbital difference amount resulting from the communication transition condition between the satellites and the orbital difference amount resulting from the condition for avoiding the physical collision of the satellites is calculated. Set and determine a near point argument common to the first, second, and third satellites. The input unit 401 inputs the difference amount between the two types of orbits and the orbital elements of the first artificial satellite, the second artificial satellite, and the third artificial satellite from the operator, and the value input by the orbital difference amount setting unit 406 Based on the above, the setting of the orbital difference common to the three satellites and the near point argument are determined and stored.

  Next, in step S706, the reference orbit of the first artificial satellite is set. The near point argument common to the three satellites determined by the orbital difference amount setting unit 406 is input to the first reference orbit setting unit 407, and the reference orbit of the first artificial satellite is calculated. For the average near point separation angle in the reference orbit of the first artificial satellite, the operator inputs a predetermined time (date time defined by the calendar date and time) and the average near point separation angle at that time in the input unit 401, This is the initial value of the reference trajectory. The mass of the reference trajectory consists of the trajectory length radius, eccentricity, trajectory tilt angle, ascending intersection, the ascending intersection, the near-point argument, and the average near-point declination calculated up to the time to determine the position of the mass point. According to the two-body problem calculation method.

  Next, in step S707, the reference orbit of the second artificial satellite is set. The near point argument common to the three satellites determined by the orbital difference amount setting unit 406 is input to the second reference orbit setting unit 408, and the reference orbit of the second artificial satellite is calculated. The average near point separation angle in the reference orbit of the second satellite is the average near point delayed by one third of the orbital period common to the three satellites from the average near point separation angle in the reference orbit of the first satellite. The point separation angle. The mass of the reference trajectory consists of the trajectory length radius, eccentricity, trajectory tilt angle, ascending intersection, the ascending intersection, the near-point argument, and the average near-point declination calculated up to the time to determine the position of the mass point. According to the two-body problem calculation method.

  Next, in step S708, the reference orbit of the third artificial satellite is set. The near point argument common to the three satellites determined by the orbital difference amount setting unit 406 is input to the third reference orbit setting unit 409, and the reference orbit of the third artificial satellite is calculated. The average near-point separation angle in the reference orbit of the third satellite is the average near-point delayed by two-thirds of the orbital period common to the three satellites from the average near-point separation angle in the reference orbit of the first satellite. The point separation angle. The mass of the reference trajectory consists of the trajectory length radius, eccentricity, trajectory tilt angle, ascending intersection, the ascending intersection, the near-point argument, and the average near-point declination calculated up to the time to determine the position of the mass point. According to the two-body problem calculation method.

  By setting the reference orbit of the first satellite, the reference orbit of the second satellite, and the reference orbit of the third satellite as described above, the direction meets at the direction meeting line formed by the intersection of adjacent orbital surfaces. Two satellites will be located in the same direction as seen from the earth. As a result, the communication between the satellite and the earth can be smoothly transferred to the communication between another satellite and the earth without interrupting the communication at the time of the direction meeting.

  In addition, since the beam width of the 45 cm diameter parabolic antenna using Ku band in the ground device is 3 degrees (pp), when the center of the parabolic antenna is pointing to one satellite at the time of communication handover, It is possible to cover two satellites with the same antenna if the satellite is at an angle of 1.5 degrees or less. For this reason, in order to cover two satellites simultaneously with one user antenna at the time of handover, it is considered appropriate to set the upper limit of the angle of separation of the two satellites to 1.5 degrees. If the holding range of the satellite is 0.1 degree, the maximum value of the separation angle of the two satellites at the time of direction meeting is twice as much as √2 times of about 0.1 degree, that is, 0.28 degree. It is well within 5 degrees.

  In the above, the artificial satellite orbiting the satellite orbit shown in FIG. 1, that is, the ascending intersection is different by 120 degrees, that is, the orbital plane is rotated by 120 degrees, and the orbital inclination angle (45 degrees). , Orbits around three satellite orbits with isochonical point argument (between about 260 degrees and about 280 degrees) and equal period (almost stationary period, 23 hours 56 minutes, synchronous orbit) and common eccentricity As an example, the three quasi-zenith satellites have been described. However, any satellite having an orbit that intersects with the satellite orbits of other satellites can be used in this embodiment. The reference trajectory can be determined by the information processing apparatus according to the embodiment.

  As described above, according to the present embodiment, the first artificial satellite and the second artificial satellite overlap the first holding range and the second holding range, and the orbital planes of these two satellites are overlapped. The second and third satellites overlap the second holding range and the third holding range so that the orbital planes of these two satellites overlap. The third satellite and the first satellite can overlap the third holding range and the first holding range so that the orbital planes of these two satellites overlap each other. Handover of communication during direction meeting to set the reference orbit of the first satellite, the reference orbit of the second satellite, and the reference orbit of the third satellite so that they can pass through the intersection line or direction meeting line Sometimes the first satellite, the second satellite, the second satellite, Possible of the satellite and the and the third satellite is possible to close the first satellite. For this reason, the distance between each satellite at the time of handover as seen from the ground can be reduced, and only one terrestrial signal transmission antenna is required, and the delay time control of the transmission signal at the time of handover is facilitated.

Embodiment 2. FIG.
In the present embodiment, an artificial satellite that orbits a satellite orbit based on the set reference orbit shown in the first embodiment and an orbit control device that performs orbit control of the artificial satellite will be described.

As mentioned above, the orbit of the satellite gradually shifts due to the effects of gravity deflection due to the flattening of the earth, lunar and solar attraction, and solar radiation pressure. The orbit control device can predict the amount of deviation in advance by measuring the distance to the satellite by the round trip time of the radio wave and estimating the position of the satellite. When the deviation becomes large and the satellite is likely to deviate from the holding range, an instruction for orbit control (orbit control instruction) is issued from the orbit control device to the satellite.
In the satellite, an orbit control thruster, that is, a gas jet mounted on the satellite is jetted to control so as to stay within the holding range. The orbit control thruster injects a fixed thrust (a combination of about 22N) on and off, and linear thrust cannot be generated. When the satellite is about to deviate from the holding range, the orbit control device makes an injection that stays in the holding range for a long time. This is an intermittent control in which the direction and the injection amount are obtained by calculation, and this information is sent to the satellite as an orbit control instruction to perform the thruster injection.
As described above, by performing the orbit control of the artificial satellite by the orbit control device, the artificial satellite can be circulated based on the set reference orbit shown in the first embodiment. When the two satellites that are directionally meeting are the first satellite and the second satellite, the first satellite and the second satellite have the first holding range and the second holding range. Can be passed through the direction meeting line. The first holding range and the second holding range are as described in the first embodiment, and are the ranges where the first artificial satellite and the second artificial satellite should stay.

  FIG. 5 is a block diagram illustrating a configuration example of the trajectory control device 50 according to the present embodiment.

  The satellite communication unit 501 communicates with an artificial satellite orbiting the satellite orbit. The satellite communication unit 501 transmits / receives radio waves and estimates orbit control instructions for estimating the satellite position.

  The satellite position estimation unit 502 estimates the satellite position by measuring the distance to the satellite based on the round-trip time of the radio wave.

The trajectory information storage unit 503 stores the same information as the information set by the first reference trajectory setting unit, the second reference trajectory setting unit, and the third reference trajectory setting unit of the information processing apparatus 40 shown in the first embodiment. For example, for the three quasi-zenith satellites shown in FIG. 1, reference orbit information and holding range setting information of each satellite are stored.
The reference orbit information is information indicating the mass point position of the reference orbit that each satellite should orbit, and the holding range setting information is information indicating the holding range at each mass point position of the reference orbit of the satellite.
The reference orbit information and the holding range setting information stored in the orbit information storage unit 503 correspond to the set reference orbit shown in the first embodiment, and the first artificial satellite, the second artificial satellite, However, the first holding range and the second holding range can be overlapped to pass the direction meeting line.

  The control instruction unit 504 may cause the satellite to deviate from the holding range based on the satellite position estimated by the satellite position estimating unit 502 and the reference orbit information and holding range setting information stored in the orbit information storage unit 503. If there is a risk of deviation, the injection direction and the injection amount of the orbit control thruster are obtained by calculation, and an orbit control instruction indicating the calculated injection direction and injection amount is generated, and the satellite communication unit It transmits to the satellite via 501.

FIG. 6 shows the form of an artificial satellite for performing orbit control.
A plurality of orbit control thrusters are installed in the artificial satellite body to generate thrust in the direction perpendicular to the orbit plane (in both + and − directions) and in the traveling direction of the orbit (in both + and − directions). The satellite body performs three-axis stability control so that the antenna is directed toward the earth and the orbit control thruster can generate thrust in a predetermined direction. By injecting the orbit control thruster in accordance with the orbit control instruction transmitted from the orbit control device 50, the satellite can always remain within the holding range.

  As described above, according to the present embodiment, the orbit control device estimates the position of the artificial satellite, and when the satellite is likely to deviate from the holding range, the orbit control instruction is sent, and the orbit control is performed according to the orbit control instruction by the artificial satellite. It becomes possible to keep the artificial satellite in the holding range by performing the above, and as a result, the first artificial satellite and the second artificial satellite overlap the first holding range and the second holding range in the direction. It becomes possible to pass the meeting line, and the first artificial satellite and the second artificial satellite can be brought close to each other at the time of handover. Therefore, the distance between the satellites at the time of handover as seen from the ground can be reduced, and only one ground signal transmission antenna is required, and the delay time control of the transmission signal at the time of handover is facilitated.

Embodiment 3 FIG.
In the present embodiment, a terrestrial communication device that communicates with the artificial satellite described in the first and second embodiments will be described.

FIG. 7 is a diagram for explaining the communication device according to the present embodiment.
In the figure, a satellite receiving antenna 701 is an antenna mounted on a vehicle, and receives content information, positioning information, and correction information distributed from the quasi-zenith satellite 60.
The content reproduction device 702 and the car navigation device 703 are communication devices mounted on a vehicle. The content reproduction apparatus 702 receives content information via the satellite reception antenna 701 and reproduces the content information. The car navigation device 703 receives the positioning information and the correction information via the satellite reception antenna 701, measures the location of the vehicle using the positioning information and the correction information, and performs car navigation. The correction information is information for correcting an error that occurs during positioning based on the positioning information, and is information created based on a D-GPS (Differential-Global Positioning System) technique or the like.
The mobile terminal 801 receives content information, positioning information, and correction information distributed from the quasi-zenith satellite 60, reproduces the content information, and uses the positioning information and correction information to determine the location of the user of the mobile terminal 801. Perform positioning.
The content playback device 702, the car navigation device 703, and the portable terminal 801 correspond to examples of communication devices.

The quasi-zenith satellite 60 with which the content reproduction device 702, the car navigation device 703, and the portable terminal 801 communicate is, for example, three satellites that orbit the three satellite orbits shown in FIG. Of the two quasi-zenith satellites, two quasi-zenith satellites meet in order at the direction meeting line in order every eight hours, and the content reproduction device 702, the car navigation device 703, and the portable terminal 801 are two quasi-zenith satellites. When a direction meeting is made on the direction meeting line, the satellite to be communicated is switched (handover is performed).
This switching may be performed by the content reproduction device 702, the car navigation device 703, and the portable terminal 801 in accordance with an instruction from the quasi-zenith satellite. When the content playback device 702, the car navigation device 703, and the portable terminal 801 receive signals from two of the three quasi-zenith satellites, the content playback device 702, the car navigation device 703, The portable terminal 801 may automatically switch.
Each quasi-zenith satellite 60 orbits the satellite orbit based on the set reference orbit shown in the first embodiment. When the content communication device 702, the car navigation device 703, and the mobile terminal 801 communicate with each other before the handover is the first artificial satellite and the satellite with which the communication is performed after the handover is the second artificial satellite, The artificial satellite and the second artificial satellite can pass the direction meeting line by overlapping the first holding range and the second holding range. That is, the first artificial satellite and the second artificial satellite communicate with two satellites (the first artificial satellite and the second artificial satellite) at the same time on the direction meeting line. Meeting direction as close as possible.
Since the quasi-zenith satellite 60 approaches the direction close to each other in this way, the content reproduction device 702, the car navigation device 703, and the portable terminal 801 each have the first and second artificial satellites with one antenna at the time of handover. And delay time control at the time of handover can be easily performed.

  In the above, mobile communication devices such as a content reproduction device, a car navigation device, and a portable terminal are examples of communication devices on the ground, but a fixed communication device may be used.

  In the above description, the content reproduction device, the car navigation device, and the mobile terminal receive content information, positioning information, and correction information. However, other information may be received.

  In the above description, the quasi-zenith satellite has been described as an example. However, a satellite other than the quasi-zenith satellite may be used as long as it is a non-stationary satellite in which satellites communicating with the ground communication device are sequentially switched.

  Thus, according to the present embodiment, the communication device includes a first artificial satellite that passes through the direction meeting line with the first holding range and the second holding range overlapping each other and approaching each other. Since communication is performed with the second artificial satellite, the first artificial satellite and the second artificial satellite can be captured by one antenna at the time of handover, and delay time control at the time of handover can be easily performed.

The figure which shows the example of the satellite orbit of a quasi-zenith satellite. The figure which shows the example of the satellite orbit of a quasi-zenith satellite. 2 is a diagram illustrating a configuration example of an information processing device according to Embodiment 1. FIG. The flowchart figure which shows the setting procedure of a reference | standard trajectory. FIG. 4 is a diagram illustrating a configuration example of a trajectory control device according to a second embodiment. FIG. 6 shows a configuration example of an artificial satellite according to a second embodiment. FIG. 6 shows an example of a communication apparatus according to Embodiment 3. The figure explaining the derivation | leading-out process of the eccentricity from a perimeter argument. The figure explaining the derivation | leading-out process of the eccentricity from a perimeter argument. The figure explaining the derivation | leading-out process of the eccentricity from a perimeter argument. The figure explaining the derivation | leading-out process of the eccentricity from a perimeter argument. The figure explaining the derivation | leading-out process of the eccentricity from a perimeter argument.

Explanation of symbols

  10 first orbit, 20 second orbit, 30 third orbit, 40 information processing device, 50 orbit control device, 60 quasi-zenith satellite, 401 input unit, 402 display unit, 403 first holding range setting unit, 404 second Holding range setting unit, 405 third holding range setting unit, 406 trajectory difference amount setting unit, 407 first reference trajectory setting unit, 408 second reference trajectory setting unit, 409 third reference trajectory setting unit, 501 Satellite communication unit, 502 satellite position estimation unit, 503 orbit information storage unit, 504 control instruction unit, 701 satellite reception antenna, 702 content reproduction device, 703 car navigation device, 801 portable terminal.

Claims (10)

Information processing for determining a first satellite orbit that the first artificial satellite orbits, and a second satellite orbit of the second artificial satellite that intersects the first satellite orbit and the satellite direction viewed from the earth center A device,
Orbit for setting the difference between the distance of the first artificial satellite from the earth center and the distance of the second artificial satellite from the earth center at the time of transfer of communication from the first artificial satellite to the second artificial satellite Difference amount setting section,
The orbit holding range of the first artificial satellite and the orbit holding range of the second artificial satellite overlap each other when viewed from the center of the earth based on the near point argument determined based on the difference amount set by the orbit difference amount setting unit. And a reference trajectory setting unit for determining a passing point. An information processing apparatus according to claim 1 is provided.
The ascending intersection red longitude is different by 120 degrees, the orbit tilt angle is equal to 45 degrees, the near point argument is equal between 260 and 280 degrees, and the orbit period is set to be approximately equal to the geostationary satellite orbit period. Information processing device. It is determined from the amount of difference between the distance from the earth center of the first artificial satellite and the distance from the earth center of the second artificial satellite at the time when the communication is transferred from the first artificial satellite to the second artificial satellite. Orbit control apparatus for controlling the orbit of the first artificial satellite so that the holding range of the first artificial satellite and the holding range of the second artificial satellite overlap each other when viewed from the center of the earth. The trajectory control device according to claim 3 is:
4. A satellite orbit in which the ascending intersection is different by 120 degrees, the orbit inclination angle is equal to 45 degrees, the near point argument is between 260 and 280 degrees and equal, and the orbit period is substantially equal to the geostationary satellite orbit period. Orbit control device according to 1. An orbiting satellite orbiting a second satellite orbit intersecting the first orbit of the other satellite orbiting as seen from the center of the earth,
A holding range is provided when the other artificial satellite and the own artificial satellite orbit each satellite orbit, and the near-point argument determined by the difference in distance from the earth center of the other artificial satellite and the own artificial satellite is used. Based on the second satellite orbit in which the holding range of the other artificial satellite and the holding range of the own artificial satellite overlap each other as viewed from the earth center. The artificial satellite according to claim 5 is:
An artificial satellite in which the ascending intersection is different by 120 degrees, the orbit inclination angle is equal to 45 degrees, the near point argument is equal between 260 degrees and 280 degrees, and the orbit period is set to be approximately equal to the geostationary satellite orbit period. Item 6. The artificial satellite according to item 5. The first artificial satellite orbiting the first satellite orbit and the second artificial satellite orbiting the second satellite orbit intersecting the first satellite orbit and the satellite direction seen from the earth center are communicable. A communication device that switches a satellite to be communicated from a first satellite to a second satellite at a predetermined satellite switching time;
A holding range of the first artificial satellite and a holding range of the second artificial satellite are provided, and based on the near-point argument determined by the orbital difference between the first artificial satellite and the second artificial satellite, A communication device that switches communication from the first artificial satellite to the second artificial satellite when the holding range and the holding range of the second artificial satellite overlap when viewed from the center of the earth. The communication device according to claim 7 is:
8. The ascending intersection red longitude is different by 120 degrees, the orbit tilt angle is equal to 45 degrees, the near point argument is equal between 260 degrees and 280 degrees, and the orbit period is set substantially equal to the geostationary satellite orbit period. Communication equipment. Communicating with the first artificial satellite orbiting the first satellite orbit and the second artificial satellite orbiting the second satellite orbit where the first satellite orbit intersects the satellite direction seen from the earth center, A communication method for switching an artificial satellite to be communicated from a first artificial satellite to a second artificial satellite at a predetermined satellite switching time,
A holding range is provided when the first artificial satellite and the second artificial satellite orbit each satellite orbit, and based on the near-point argument determined from the orbital difference between the first artificial satellite and the second artificial satellite, A communication method for switching communication when a holding range of a first artificial satellite and a holding range of a second artificial satellite overlap when viewed from the earth center. The communication method according to claim 9 is:
The ascending intersection red longitude is different by 120 degrees, the orbit tilt angle is equal to 45 degrees, the near point argument is equal between 260 degrees and 280 degrees, and the orbit period is set to be approximately equal to the geostationary satellite orbit period. Communication method.

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