JP2001086053A - Artificial satellite operating method and orbit determining method for artificial satellite - Google Patents

Abstract

PROBLEM TO BE SOLVED: To quantitatively determine which artificial satellite is to be selected from among other artificial satellites that are not used at present when the use of an artificial satellite which is presently used becomes difficult due to fault, etc. SOLUTION: When communication is performed between the artificial satellites and a ground terminal or broadcast is performed to the ground terminal by using plural artificial satellites orbiting around the earth, for example, when a second artificial satellite 712 is out of order and the use of it becomes impossible at time 13:00, operation of the second artificial satellite 712 is stopped. And a third artificial satellite 713 with the highest angle of elevation at 13:00 among the artificial satellites in the service area is selected and operated.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a method for operating an artificial satellite used for data communication or broadcasting and a method for determining an orbit of an artificial satellite.

[0002]

2. Description of the Related Art Artificial satellites are roughly classified into geostationary satellites and orbiting satellites orbiting the earth. Conventionally, as a backup of an artificial satellite, only a backup for a geostationary satellite has been proposed, and a backup for an orbiting satellite has not been considered. For example, in a geostationary satellite, two geostationary satellites have been launched in the same orbital position, and when an abnormality occurs in one geostationary satellite, a method is used in which the other geostationary satellite is switched to the other geostationary satellite. . A representative example of this is the broadcasting satellite BS series of the Japan Broadcasting Corporation.

[0003] Also, it is known that if the sun is present in the vicinity of the viewing direction when viewing the artificial satellite from the ground terminal, noise (sun noise) is generated by the sun, which hinders communication and broadcasting. I have. As a conventional countermeasure against such solar noise, an organization operating a satellite predicts the time at which solar noise will occur in a certain service area from orbit data of the satellite and obtains such information. (Time) was only notified to the user, and no special measures were taken. This is because the conventional communication or broadcasting service using artificial satellites is based on the assumption that a geostationary satellite is used, so that no effective measures can be taken against the deterioration of the line quality due to solar noise.

[0004]

As described above, according to the prior art, in a satellite system having a plurality of artificial satellites, when one artificial satellite fails and becomes unusable, another artificial satellite cannot be used. There is no quantitative method for determining which satellite to select from among them, and the accessibility between the same ground terminal and the satellite (ease of data exchange in communications etc.) will be greatly impaired. There is a problem.

[0005] Further, in the prior art, effective measures for improving the line quality could not be taken against the deterioration of the line quality due to the increase of the solar noise. In communication, errors occur in transmitted data, and in broadcasting, received image quality and sound are disturbed.

Further, in the prior art, in a satellite system having a plurality of satellites, when a backup is taken into consideration, a method of determining a satellite orbit that does not degrade the function or performance of the system after the backup is completely considered. Had not been.

[0007] An object of the present invention is to quantitatively determine which satellite to select from other satellites when it becomes difficult to use the satellite currently in use due to a failure, solar noise, or the like. It is to provide a satellite operation method that is possible.

Another object of the present invention is to provide a method for determining an artificial satellite orbit in a satellite system that takes backup into account without deteriorating the function or performance of the system even after backup.

[0009]

SUMMARY OF THE INVENTION In order to achieve the above object, a method of operating an artificial satellite according to the present invention uses a plurality of artificial satellites orbiting the earth to communicate with a ground terminal, or When broadcasting to terrestrial terminals, if one of the satellites in a certain service area out of satellites in the same satellite system fails and becomes unusable, the operation of the satellite And the satellite having the highest elevation angle is selected and operated from among the satellites in the service area excluding the artificial satellite whose operation has been stopped.

According to the above operation method, when the satellite becomes unavailable due to a failure of the satellite, etc.
Since the artificial satellite with the highest elevation angle is selected in the service area, the artificial satellite can be selected quantitatively. Then, the satellite is switched to the selected artificial satellite, and communication or broadcasting is performed.

[0011] The artificial satellite operation method of the present invention is used when communicating with a ground terminal or broadcasting to a ground terminal using a plurality of artificial satellites orbiting the earth. If the sun is present in the vicinity of the viewing direction when viewing the currently used artificial satellite from the ground terminal, the operation of the artificial satellite is stopped, and within the service area where the artificial satellite that has stopped operating exists. It is characterized by selecting and operating another artificial satellite.

According to the above operation method, when the sun is in the vicinity of the viewing direction when viewing the artificial satellite and this affects the communication or broadcast line quality, the operation of the artificial satellite is stopped. In addition, it is possible to eliminate the influence of solar noise and suppress the deterioration of the line quality. When another artificial satellite is selected, the artificial satellite having the highest elevation angle at that time in the service area is selected except for the artificial satellite whose operation has been stopped.

Further, the method for determining the orbit of an artificial satellite according to the present invention can be used for determining the orbit of a plurality of artificial satellites which communicate with a ground terminal while circling the earth or broadcast to the ground terminal. In advance, an elevation range measured from the zenith direction, in which transmission, reception or transmission and reception of radio waves can be performed between the ground terminal and each of the artificial satellites, is determined in advance. The orbit of each artificial satellite is determined so that the artificial satellite always exists within the elevation angle range.

According to the above-described determination method, in an artificial satellite system in consideration of backup, an artificial satellite orbit can be determined so as not to deteriorate the function or performance of the system even after backup.

[0015]

Embodiments of the present invention will be described below with reference to the drawings. (Embodiment 1) FIG. 1 is a diagram for explaining Embodiment 1 of the present invention, showing an example of the elevation angle of each satellite viewed from a service area. In FIG. 1, the vertical axis indicates the elevation angle of the satellite viewed from the service area, and the horizontal axis indicates the time. In the figure, a curve 711 indicates a time series change of the elevation angle of the first satellite, and a curve 71
Numeral 2 represents a time series change of the elevation angle of the second satellite, and a curve 713 represents a time series change of the elevation angle of the third satellite.

It is assumed that the current time is 13:00 and the second satellite is operating. The elevation angle of the second satellite is 84.9 degrees as can be seen from FIG. At this time, if the second satellite becomes inoperable due to a failure or the like, the operation of the second satellite is stopped. Here, among the satellites other than the second satellite, since the third satellite has an elevation angle of 66.0 degrees and the largest elevation angle in the service area, the third satellite is selected and operated.

(Embodiment 2) FIG. 2 is a diagram for explaining Embodiment 2 of the present invention. The flow of a method for determining the orbit of an artificial satellite that does not deteriorate the functional performance as a satellite system even when a satellite fails. Is shown.

(1) Setting the major axis of the orbit (hereinafter, for example, about 42178 km of an orbit having a period of about 24 hours is used).

(2) Setting of a perigee argument (hereinafter, the service area is assumed to be the northern hemisphere, and the perigee argument which makes the perigee come over the southern hemisphere and the apogee come over the northern hemisphere is used).

(3) Setting of the eccentricity, the orbital inclination angle, the ascending intersection red diameter, and the near-point departure angle. (A) Setting of the half-vertical angle The allowable range of the direction in which the satellite can be seen from the zenith direction is determined by the half-vertical angle centered on the zenith. For example, the half apex angle during normal operation is 20 degrees, and the half apex angle at off-nominal time when one satellite cannot be operated is 30 degrees. (B) Setting of the number of satellites, the ascending intersection red diameter of each satellite, and the near-angle separation In order to provide continuous service from the satellites to the service area, the satellites are arranged one by one on the orbit plane, and Are desirably identical in shape. The ascending red point of each orbit may be separated by an angle obtained by dividing 360 degrees by the number of satellites. That is, assuming that the number of satellites is five, the satellites interlock on five orbit planes separated by an ascending intersection red diameter of 72 degrees. Also, when setting the near-point separation, when one satellite is at perigee in its orbit, the right-point separation of other satellites is calculated by dividing the period by the number of satellites constituting the system. What is necessary is just to separate by a corresponding angle. For example, in the case of five satellites, it is sufficient that the satellites on adjacent orbits are separated from each other in order of 1/5 of the orbital period. With the setting of the ascending intersection red diameter and the right-neighbor angle separation, the ground trajectories of each orbit almost coincide with each other, and the satellites can appear over the same area in order.

(C) Setting Initial Value of Orbit Inclination Angle Analysis is performed using an orbit inclination angle of the same angle as the latitude near the center of the area of the service area as an initial value. (D) Calculation of Visible Time from Representative Point in Service Area The time during which a satellite can be seen inside a cone formed by a half-vertical angle in the zenith direction from the representative point in the service area is obtained for each satellite.
This is performed for both the normal operation and the off-nominal time shown in FIG. This can be obtained by geometric calculation and Kepler's law. By changing the eccentricity from 0.0 to 1.0, a range of the eccentricity in which the visible time length from the representative point is substantially the same is obtained. Next, the orbit inclination angle is slightly changed from the initial value, and the eccentricity is similarly changed to obtain a range of the eccentricity in which the visible time from the above-mentioned representative point is almost the same time length.

(E) Setting of Combination of Orbit Inclination Angle and Eccentricity By repeating the analysis shown in the above (d), a combination of the orbit inclination angle and the eccentricity value is obtained. At this time, it is also possible to return to (b) and reset the number of satellites if necessary. (F) Reset of the ascending intersection red diameter and the near-point departure Lastly, the reference time is set in accordance with the launch time of the satellite, etc. Set.

Table 1 shows the orbital parameters of the satellite system that enable services throughout Japan to be provided with an elevation angle of 70 degrees or more during normal operation and at an elevation angle of 60 degrees or more during off-nominal operation when one satellite fails.

[0024]

[Table 1]

It is needless to say that the above-mentioned satellite orbit determination method is applicable even if the service area and the operation elevation angle are different.

Third Embodiment Next, a third embodiment of the present invention will be described with reference to FIG. As shown in FIG. 3, a terrestrial terminal 500 and a satellite-operated earth station 7
00 is provided. In the satellite orbit 300, a satellite 410 and a satellite 420 orbit. Sun 200
Is generally known from observations at an observatory or the like, and a position at an arbitrary time can be predicted.

The positions of the satellites 410 and 420 are calculated by the following equations. Note that all quantities with _ are vector quantities. Az = arctan [−E_ · P _ / (V_ · P_)] El = arcsin [Z_ · P_] P = (S_−L _) / | S_−L_ | S _ = (sinφ _L cos (λ _L + θ _G ), sin φ _L sin (λ _L + θ _G ), co
_{sφ L) E _ = (-} sin (λ L + θ G), cos (λ L + θ G), 0) Z _ = (cosφ L cos (λ L + θ G), cosφ L sin (λ L + θ G), si
_{nφ L) V_ = a (cosE} -e) P_ + a√ (1-e 2) sinEQ_ L _ = ((n s + h) cosφ L cos (λ L + θ G), (n s + h) cosφ
_{L sin (λ L + θ G} ), {n s (1-e E 2) + h} sinφ L) P _ = (cosωcosΩ-sinωsinΩcosi, cosωsinΩ + sin
ωcosΩcosi, sinωsini) Q _ = (-sinωcosΩ-cosωsinΩcosi, -sinωsinΩ + cos
_{ωcosΩcosi, cosωsini) n s = R} E / √ (1-e E 2 sin 2 φ L) R E = 6378.136km ( average equatorial radius of the ^{_{earth) e E 2 = f E (}} 2-f E) (e E Is the eccentricity of an ellipsoid representing the shape of the earth) f _E = 1 / 298.257 θ _G = Θ ₀ + E _q + ΔUT 1 (θ _G is the rotation angle of the earth) where Θ ₀ : Greenwich mean stellar time E _q : point difference (Difference in rotation angle of the earth caused by nutation of the rotation axis of the earth) ΔUT1: Difference caused due to uneven rotation speed of the earth.

Az: an azimuth angle indicating an angle from north to east from true north at 0 °, El: an elevation angle at 0 ° in the horizontal direction, and E: eccentric near-point departure. Then, the eccentric near point departure angle E is obtained from the true near point departure angle θ and the eccentricity e by the following equation. E = 2arctan {{((1-e) / (1 + e)) tan (θ / 2)} The six orbital elements used for defining the above-mentioned orbital are defined as values at a certain reference time.

That is, orbit major radius a: major radius of ellipse eccentricity e: flatness of ellipse (0 ≦ e <1) orbit inclination angle i: inclination of orbit surface from equatorial plane (0 degrees ≦ i ≦
180 °) Ascending intersection red diameter Ω: Angle measured eastward from the vernal equinox direction at the point where the orbit crosses the equatorial plane from the southern hemisphere to the northern hemisphere (0 ° ≦ Ω ≦ 180 °) Perigee argument ω: Perigee position on the orbital plane From the ascending intersection point (0 degrees ≤ ω ≤ 360 degrees) True near point departure angle θ: The angle formed by the line segment connecting the perigee and the focus of the ellipse and the line segment connecting the position of the satellite in orbit and the focus of the ellipse (0 degree ≤
θ ≦ 360 degrees).

The viewing direction of the sun 200 obtained by the above method is compared with the viewing direction of the satellite 410 or the satellite 420,
When the viewing directions of the sun and the satellite fall below a certain threshold, a command is transmitted from the satellite operating earth station 700 to the satellite,
Stop operation. FIG. 3 shows a case where the operation of the satellite 410 is stopped.

In this embodiment, the number of satellites is two, but the same applies to the case where the number of satellites is three or more.

Further, in the present embodiment, the number of satellite orbits as viewed from the service area on the earth is one, but the same is true even if there are a plurality of satellite orbits. The satellite orbit is a geosynchronous orbit,
Any of a low orbit, a medium orbit, and an elliptical orbit may be used. Here, the low orbit is about 100-200 km, and the middle orbit is about 20 km.
It is about 0 to 20000 km.

Further, in the present embodiment, the ground terminal may be a mobile terminal or a fixed terminal. A plurality of ground terminals may exist.

(Embodiment 4) FIG. 4 is a diagram for explaining Embodiment 4 of the present invention, showing an example of the elevation angle of each satellite viewed from a service area. In FIG. 4, the vertical axis indicates the elevation angle of the satellite viewed from the service area, and the horizontal axis indicates the time. In the figure, a curve 731 represents a time series change of the elevation angle of the first satellite, a curve 732 represents a time series change of the elevation angle of the second satellite, and a curve 733 represents a time series change of the elevation angle of the third satellite. I have.

It is assumed that the current time is 14:00 and the second satellite is operating. The elevation angle of the second satellite is 81.3 degrees as can be seen from the figure. At this time, when the difference between the viewing direction of the second satellite and the viewing direction of the sun falls below a certain threshold, the operation of the second satellite is stopped. Then, the satellite is switched to a satellite other than the second satellite to operate the satellite. Here, among the satellites other than the second satellite, the third satellite is selected and operated because the third satellite has an elevation angle of 72.8 degrees and has the largest elevation angle in the service area.

From the position and velocity of the satellite observed from the earth, the above-mentioned six orbital elements are derived by the following equations. Eccentricity e = | e_ | where e _ = {(v ² −μ / r) r _− (r_ · v_) v _} / μ orbital major radius a = (1−e ² ) | h_ | ² / μ h_ = r_ × v_ (cross product of r_ and v_) orbit inclination angle i = arccos [(K_ · h _) / | h_ |] (i is always positive) Ascending intersection red diameter Ω = arccos [I_ · (K_ × h _) / | K_ × h_
|] However, if I_ · (K_ × h _)> 0, Ω <180 ° perigee argument ω = arccos [(K_ × h_) · e _ / (| K_ × h_
|| e_ |)] However, if (K_ × h_) · e_> 0, ω <180 °, near-point separation θ = arccos [(e_ · r0 _) / | e_ · r0_], where (e_ · r0_) If> 0, θ <180 ° Note that all the quantities with _ are vector quantities, and the base point of r_ is the center of the earth.

Here, μ: the gravitational constant of the earth (= 3.98)
6 × 10 ⁵ km ³ / sec ² ) r: Position of satellite (observed amount) r0: Position of satellite in the first phase of orbit v: Speed of satellite (observed amount) I_: X-axis direction (Earth equinox from the center of the earth Unit vector K_: unit vector in the Z-axis direction (positive in the north pole direction from the earth center).

[0038]

As described above, according to the present invention,
When one artificial satellite in the service area breaks down and cannot be used, it is possible to quantitatively determine which artificial satellite to select from other artificial satellites. Then, since the satellite with the highest elevation angle is selected, it is possible to minimize the loss of accessibility from the ground terminal to the artificial satellite.

When the sun is present in the vicinity of the viewing direction of the artificial satellite, the operation of the artificial satellite is stopped, so that the influence of the noise due to the sun is eliminated and the deterioration of the line quality is prevented. Can be.

Further, in the artificial satellite system in consideration of the backup, since the artificial satellite orbit is determined so as not to deteriorate the function or performance of the system even after the backup, the reliability of the system can be improved.

[Brief description of the drawings]

FIG. 1 is an explanatory diagram of an artificial satellite operation method according to a first embodiment of the present invention.

FIG. 2 is a flowchart of an artificial satellite orbit determination method according to Embodiment 2 of the present invention.

FIG. 3 is an explanatory diagram of an artificial satellite operation method according to a third embodiment of the present invention.

FIG. 4 is an explanatory diagram of an artificial satellite operation method according to a fourth embodiment of the present invention.

[Explanation of symbols]

 100 Earth 200 Sun 300 Satellite orbit 410 Satellite 420 Satellite 500 Ground terminal 700 Satellite operation earth station 711 Time series change of elevation angle of first satellite 712 Time series change of elevation angle of second satellite 713 Time of elevation of third satellite Series change 731 Time series change of elevation angle of the first satellite 732 Time series change of elevation angle of the second satellite 733 Time series change of elevation angle of the third satellite

Claims (4)

[Claims] When communicating with a terrestrial terminal using a plurality of satellites orbiting the earth or broadcasting to a terrestrial terminal, a satellite in the same satellite system is used. When one satellite in a certain service area fails and becomes unusable, the operation of the satellite is stopped, and the artificial satellite in the service area excluding the satellite whose operation has been stopped is removed. A satellite operation method characterized by selecting and operating the satellite with the highest elevation angle at that time among the satellites. 2. When communicating with a ground terminal using a plurality of satellites orbiting the earth or broadcasting to a ground terminal, the satellite currently in use is transmitted to the ground terminal. If the sun is present in the vicinity of the viewing direction when viewed from the terminal, the operation of the artificial satellite is stopped, and another artificial satellite is selected and operated in the service area where the artificial satellite stopped operating. A satellite operation method. 3. The artificial satellite operating method according to claim 2, wherein the artificial satellite having the highest elevation angle at that time in the service area is selected as the other artificial satellite except for the artificial satellite whose operation has been stopped. A satellite operation method characterized by the above-mentioned. 4. When communicating with a ground terminal while circling the earth or determining orbits of a plurality of artificial satellites broadcasting to the ground terminal, the ground terminal and each of the artificial satellites Between the transmission of radio waves,
Reception or transmission / reception can be performed, an elevation angle range measured from the zenith direction is determined in advance, and a plurality of satellites among the respective artificial satellites are always present in the elevation angle range. An orbit determination method for an artificial satellite, which determines an orbit.