Abstract:
An information providing method and apparatus in which a radio wave signal is transmitted from plural artificial satellites and received at plural receiving equipment located within a pre-defined service area, and receiving within a pre-defined range of operational elevation angle by a receiving antenna. The plural artificial satellites are viewable in turn within a pre-defined range of operational elevation angle at an arbitrary location within the pre-defined service area, and an orbit is arranged so that at least one artificial satellite is viewable. Each orbit of the plural artificial satellites is arranged on an individual orbital plane according to a designated service hour. The individual orbit is located in a range where an orbital inclination angle is 63.4 degrees or closer to Japan, and a right ascension of a north-bound node of the individual artificial satellite is separated approximately at regular intervals.

Description:
CROSS REFERENCE TO RELATED APPLICATION  
       [0001]     This is a continuation of application Ser. No. 11/057,025, filed Feb. 14, 2005, which is a continuation of U.S. application Ser. No. 10/410,174, filed Apr. 10, 2003, which is a continuation U.S. application Ser. No. 09/471,503, filed Dec. 23, 1999, now U.S. Pat. No. 6,695,259, which is a continuation-in-part of U.S. application Ser. No. 09/081,551, filed May 20, 1998, now U.S. Pat. No. 6,352,222, the subject matter of which is incorporated by reference herein.  
     
    
     BACKGROUND OF THE INVENTION  
       [0002]     The present invention relates to a communication system, and a communication sending and receiving device and a communication terminal in the system, and particularly to a communication system suitable for a satellite, a satellite orbit and a satellite orbit configuration algorithm usable in the field of communication and broadcast such as satellite communication, satellite broadcast, satellite mobile communication and in the field of observation with regard to a celestial body of the satellite traveling around, and suitable for a communication and broadcast system to which the satellite is applied, and a satellite communication sending and receiving device and a communication terminal in the system.  
         [0000]     (a) The prior art with regard to setting of an orbital element (argument of perigee) of an artificial satellite  
         [0003]     In a case where an artificial satellite is traveling around the earth as the center, the orbit of the artificial satellite always fluctuates under the influence of the nonuniformity of the earth&#39;s gravitational field, the attractive forces of the moon and the sun, the atmospheric drag and the sun&#39;s light pressure. From this viewpoint, the orbit of the artificial satellite traveling around the earth as the center can not be a circular orbit but is an elliptical orbit in a broad sense having a little eccentricity.  
         [0004]     Accordingly, the “elliptical orbit” in the present specification is defined as an “orbit having an eccentricity larger than zero and smaller than 1, and the eccentricity is not intended to become zero in a process of setting an orbital element of the orbit onto which the artificial satellite is injected in order to attain the purpose”.  
         [0005]     As an example of practically used artificial satellites having an elliptical orbit, there is the Molniya satellite (about a 12-hour orbital period) which has been used by Russia since the era of the former USSR. As artificial satellites having an elliptical orbit, a communication satellite called as Archimedes (about an 8-hour orbital period) is proposed in Europe. Further, a tundra orbit of about a 24-hour orbital period is proposed though it is not practically used yet. A common point in these satellites is that all the satellites has an orbital inclination angle of approximately 63.4 degrees.  
         [0006]     In general, the perigee of an orbit on which an artificial satellite is traveling will rotate on the orbital plane under the influence of the nonuniformity of the earth&#39;s gravitational field (oblateness of the earth&#39;s shape). However, by setting an orbital inclination angle to 63.4 degrees in a mathematical model for calculating a time-varying rate of the argument of perigee, a multiplicative term becomes zero to make the time-varying rate zero. Therefore, it is considered that the rotation is stopped.  
         [0000]     (b) The prior art with regard to orbit configuration methods for a plurality of artificial satellites  
         [0007]     Communication systems using a plurality of artificial satellites traveling on elliptical orbits have been in practical use and studied. Although arranging of a plurality of artificial satellites in the above-mentioned Molniya satellite and Archimedes is described, no detailed method of arranging the orbits is described.  
         [0008]     Although in resent years a communication system using a plurality of artificial satellites is proposed, no detailed technique with regard to orbit configuration method is disclosed. Therefore, a detailed orbit configuration technique is required.  
         [0009]     On the other hand, “an artificial satellite on an orbit having a long stretch of time staying in the zenith direction, a method of controlling the orbit and a communication system using the artificial satellite and the method” is proposed in Japanese Patent Application Laid-Open No. 11-34996.  
         [0000]     (c) The prior art with regard to mobile communication and broadcast to a mobile object  
         [0010]     In the past, when a television broadcast was tried to receive in a mobile object such as a vehicle, there were problems in that the picture came out badly in an area far from a broadcasting facility of a television station, that screen flicker occurred even at a place near the broadcasting facility of the television station, and that receivable channel varied by moving. When a television broadcast from a communication and broadcast satellite on a stationary orbit was received on the mobile object, it was difficult to comfortably enjoy watching television on the mobile object because the electromagnetic wave was frequently shielded by artificial structures such as buildings, trees and natural geographical features.  
         [0011]     Transmission of a large volume of data such as images from a mobile object such as an ambulance can not be performed by an existing ground communication infrastructure and an existing communication satellite.  
         [0012]     In order to solve the above-mentioned problems,. a method of setting orbital elements of an artificial satellite for transmitting a large volume of data to a mobile object such as a vehicle is proposed and in addition an orbital element of the artificial satellite is also proposed in Japanese Patent Application Laid-Open No. 11-34996.  
         [0013]     The problems in the prior art will be described below, corresponding to each of the above items (a), (b) and (c) based on examples known in the art.  
         [0000]     (a) Problems with regard to setting of an orbital element (argument of perigee) of an artificial satellite, and an object of the present invention  
         [0014]     In the above mentioned Molniya satellite, Archimedes and Tundra orbit, all the orbital inclination angles of them are fixed to approximately 63.4. It seems that the main object is to suppress the rotation of the perigee on the orbit plane. On the other hand, there is an advantage in using an orbital inclination angle as large as approximately 63.4 degrees because the area using the above-mentioned artificial satellites is a higher latitude area such as Europe and Russia.  
         [0015]     As for the location of Japan, the territory spreads from a middle latitude to a low latitude as Etorofu island in the northernmost end is situated in latitude approximately 45 degrees north and Okino-torishima in the southernmost end is situated in latitude approximately 20 degrees north. Therefor, when the orbital inclination angle of 63.4 degrees is employed as described above, the artificial satellite system becomes difficult to be used from the territory of Japan unless an altitude of the orbit is sufficiently high. Accordingly, when an orbit of an artificial satellite matching the location of the territory of Japan is taken into consideration, the orbital inclination angle can not help employing a value other than approximately 63.4 degrees, and consequently the perigee of orbit rotates.  
         [0016]     In order to control the rotation of the perigee, propellent for controlling the rotation needs to be mount on the artificial satellite. An analytical simulation was performed on a case where the orbital inclination angle is 40 degrees and the eccentricity is 0.24 among the orbital elements proposed in, for example, Japanese Patent Application Laid-Open No. 11-34996. As a result, it was found that an amount of the propellent for controlling the argument of perigee becomes dominant to the total amount of poropellent depending on a condition of setting the orbital elements because there occurs a case where the orbit control propellent of approximately 75% must be used at maximum for controlling the argument of perigee to nearly 270 degrees. Thereby, devices mountable on the artificial satellite may be reduced, or on-orbit lifetime of the artificial satellite may be shortened.  
       SUMMARY OF THE INVENTION  
       [0017]     An object of the present invention is to set an argument of perigee of one of the six-orbital elements at a setting stage of the orbital elements of an orbit on which the artificial satellite travels in order to solve the above-mentioned problem.  
         [0000]     (b) Problems with regard to orbit configuration methods for a plurality of artificial satellites, and an object of the present invention  
         [0018]     Many communication systems which use an elliptical orbit of about a 270-degree argument of perigee and about an 8-hour orbital period are proposed. Apogees of these systems appear above three areas of Europe, North America and Japan, and an object of these systems is to provide communication services using three or six artificial satellites. It can be supposed that the number of the satellites, that is, three or six is intuitively or naturally determined so that three of the artificial satellites come around in the sky above the three areas at a time, respectively. The references do not describe any case where different number of satellites is employed. Further, as to methods of setting the orbital elements, most of the references do not describe any specific numerical values except for a semimajor axis of 20,270 km which is mathematically derived from an orbital period of 8 hours and an orbital inclination angle of 63.4 degrees which is considered to be stable in orbit kinetics. In addition, there is no description on the method of deriving the values.  
         [0019]     When an orbit of an artificial satellite is mapped on the ground, what can be expressed geometrically and visually are only four orbital elements of an orbital semimajor axis or orbital period, an eccentricity, an orbital inclination angle and an argument of perigee. Therefore, in a stage of preliminary conceptual design, it is sufficient to set these four elements. This is considered one of the reason why the orbital elements are not clearly described and the deriving method is not described.  
         [0020]     In Japanese Patent Application Laid-Open No. 11-34996, a method of setting orbital elements of an artificial satellite having a satellite orbital period of approximately 12 hours or 24 hours is proposed. Further, there is description on values of orbital elements of a satellite and number of satellite in a case of service target area of. Japan and an elevation angle above 70 degrees. However, there is no proposal on numerical values of orbital elements and number of satellites for orbital periods other than the above-mentioned orbital periods. What is described in the above-mentioned reference is a method of setting orbital elements of an artificial satellite on an elliptical orbit having a long stretch of time staying in the zenith direction in a specified area, and the method can not be applied to all the cases of setting orbital elements of an artificial satellite.  
         [0021]     Further, description will be made below on systems using a plurality of artificial satellites which are proposed now or have been developed now. In a mobile communication satellite system, the service target is global, and the satellite travels on a circular orbit (zero eccentricity) having a constant semimajor axis and a constant orbital inclination angle, but the other orbital elements and the method of deriving them are not disclosed. The reason why this is not disclosed may be that they think this belongs to the know-how of the inventors proposed the system using the artificial satellites. An earth survey satellite system developed is a combination of satellites traveling on a sun synchronous semi-tropical orbit so as to survey all over the world. On the other hand, in a case where communication service or surveillance is concentratively and continuously performed to a specified area, stationary satellites are used.  
         [0022]     In a case where communication service or broadcast service is intended to be concentratively and continuously performed using an arbitrary number of artificial satellites to a specified area on a celestial body of the satellites traveling around, or in a case where a specified area on a celestial body of the satellites traveling around or a weather condition of the area is concentratively and continuously observed, an object of the present invention is to provide a method capable of being generally applied to setting of orbital elements of the artificial satellite, particularly, a method of setting an orbital semimajor axis, an eccentricity, an orbital inclination angle, an argument of perigee, right ascension of north-bound node and true anomaly of the arbitrary number of artificial satellites, and to provide detailed numerical values of the orbital elements obtained from orbit design according to the method.  
         [0000]     (c) Problems with regard to mobile communication and broadcast to a mobile object, and an object of the present invention  
         [0023]     It is clear that the existing communication infrastructures such as common line telephones, cellular phones and personal handy phones can not cope with large volume communication to mobile objects.  
         [0024]     The stationary satellite communication system likely to cause communication interruption by artificial structures and natural geographical features can not cope with large volume communication to mobile objects.  
         [0025]     It is clear that a satellite communication system using a low-to-middle altitude orbit such as Iridium currently under development can not cope with large volume communication to mobile objects because the duration of time while the satellite comes and stays visible in a high elevation angle is as short as several minutes.  
         [0026]     The various kinds of communication systems described above can not sufficiently cope with the communication to the mobile-objects, but their applicability to digital television broadcast and digital voice broadcast to the mobile objects is negative.  
         [0027]     In the broadcast service to the mobile objects using an artificial satellite, the artificial satellite must stably stay visible in a high elevation angle for a long duration of time within a service target area.  
         [0028]     The words “an artificial satellite is visible” in the present specification is determined that “an artificial satellite stays within a spatial area, under the condition of which communication between an artificial satellite tracking and controlling ground station, various kind of satellite communication send and receive facilities and an artificial satellite can be performed with electromagnetic wave of light”.  
         [0029]     In order to realize the above, it is generally thought that an elliptical orbit of which the apogee stays in the sky above a service target area is preferable, but appropriate methods and algorithms of setting the orbital elements are not firmly proposed except for in Japanese Patent Application Laid-Open No. 11-34996.  
         [0030]     In the orbital elements proposed in Japanese Patent Application Laid-Open No. 11-34996, the minimum value of eccentricity is 0.24. Even if the value of eccentricity is employed, the distance from the ground to the satellite is generally larger than the distance from that position to a stationary satellite. Therefore, there are the following problems to be solved.  
         [0031]     (1) Free spatial loss on electromagnetic wave transmission becomes large, and accordingly the communication/broadcasting devices mounted on the artificial satellite are required to have higher sending and receiving capability. In more detail, in the artificial satellite side, a larger antenna or a sender having a larger output power and a receiver having a higher receiving capability are necessary. In the ground side, a send and receive facility for satellite communication similarly requires a larger antenna or a sending unit having a larger output power and a receiving unit having a higher receiving capability.  
         [0000]     (2) Communication delay becomes larger because the distance from the ground larger.  
         [0032]     Further, the distance to the artificial satellite in service becomes different between one end part of a service target area and the other end part in the opposite side because the eccentricity is somewhat large. Thereby, when the artificial satellite in service is switched, breakdown time may occur in broadcasting.  
         [0033]     In order to solve the above problems, the present invention improves the orbital elements proposed in Japanese Patent Application Laid-Open No. 11-34996 from the viewpoint of “communication with ground”. An object of the present invention is to set more effective ranges of orbital elements within the specified service area of Japanese territory.  
         [0034]     Although an object of the present invention is to individually solve the problems described the above items (a), (b) and (c), the object is also to solve combinations of the items (a), (b) and (c), or all of the items (a), (b) and (c) at a time. Another object of the present invention is to provide a method of deriving orbital elements of artificial satellites capable of making mobile communication and mobile broadcast easy to the specified area of Japanese territory using a plurality of artificial satellites by solving the items (a), (b) and (c) together, and at the same time to express the orbital elements suitable for the Japanese territory by limiting the ranges.  
         [0035]     Furthermore, a further object of the present invention is to construct various kinds of systems utilizing a plurality of artificial satellites after solving the -problems described in the above items (a), (b) and (c).  
         [0036]     Description will be made below, corresponding to each of the above items (a), (b) and (c).  
         [0000]     (a) With regard to setting of one of orbital elements of an argument of perigee  
         [0037]     In a plan using an artificial satellite, a duration to be operable of the artificial satellite is general defined as a mission lifetime. It is necessary to accurately estimate over the period of the mission lifetime using a computer how mach an argument of perigee changes from an initial value of the argument of perigee just after the artificial satellite is injected onto an orbit the satellite should be injected in order to attain its purpose under the condition that the argument of perigee is not controlled at all.  
         [0038]     In order to attain the above objects, it is assumed in the present invention that the initial value of the argument of perigee of an orbital element of an orbit on which the artificial satellite travels is set using the above-mentioned estimated value.  
         [0039]     In order to attain the above-mentioned objects, the present invention uses an artificial satellite comprising an attitude sensor for detecting its own attitude, a computer for processing detected attitude data, an actuator or gas jet unit for maintaining or changing the attitude using the computer, a gas jet unit for changing its own orbit, and a communication unit for establishing a communication line between the artificial satellite and a control station with electromagnetic wave. In a case where a center celestial body of the artificial satellite traveling around is the earth, the artificial satellite may comprise a unit for receiving electromagnetic wave from a GPS satellite and calculate its own position and velocity. Here, the GPS satellite is a generic name for Navster satellite composing Global Positioning System (GPS) of the USA, Glonass satellite of Russia for navigation, transportation multipurpose satellite of Japan and so on.  
         [0000]     (b) With regard to orbit arranging method for a plurality of satellites  
         [0040]     In order to attain the above-mentioned objects, in the present invention, an orbit on which an artificial satellite travels is defined by six orbital elements obtained under input conditions of a specified area supplied with service using the artificial satellites, number of the artificial satellites, frequency of service by the artificial satellites to the service target area, a duration time of service by one of the artificial satellites to the service target area and reference time defining the orbital elements.  
         [0041]     In more detail, at defining the orbital elements, the six orbital elements are determined by a process of defining number of artificial satellites, a process of defining an orbit semi-major axis, a process of setting an eccentricity, an orbit inclination angle and an argument of perigee, a process of setting a right ascension of north-bound node, a process of setting a true anomaly, and repeating of all the processes from the process of defining number of artificial satellites to the a process of setting a true anomaly.  
         [0042]     In order to attain the above-described object, the artificial satellites similar to the above item (a) may be employed in the present invention.  
         [0000]     (c) With regard to mobile communication and broadcast to a mobile object  
         [0043]     In order to attain the above-described object, the artificial satellites similar to the above item (a) may be employed in the present invention.  
         [0044]     Further, in order to attain the above-described object, the present invention employs a group of artificial satellites composed of three or four artificial satellites traveling on three or four elliptical orbits with an orbital period of 24 hours, wherein each of the orbits is formed so that an orbital inclination angle is larger than 37 degrees and smaller than 44 degrees and an eccentricity is not larger than 0.24, or so that an orbital inclination angle is larger than 40 degrees and smaller than 44 degrees and an eccentricity is larger than 0.24 and smaller than 0.35. Therein, each of the artificial satellites is arranged on each of the orbits. Here, with regard to the orbital period of 24 hours, the words, 24 hours, in this specification are defined as a time duration including an error of ±10 minutes to 23 hours 56 minutes.  
         [0045]     The followings are means commonly used in the above items (a), (b) and (c).  
         [0046]     In order to attain the above-described object, in the present invention, the artificial satellites traveling on the orbits in accordance with the present invention are used in systems using various kinds of satellites such as an orbit control system for controlling the orbit of the artificial satellite, a satellite communication system for performing satellite communication through the artificial satellite, an earth survey system using the artificial satellite mounting an earth survey unit and the like.  
         [0047]     A satellite communication send and receive apparatus in the satellite communication system is a unit comprising a send and receive means for performing signal sending and receiving with the artificial satellite when the artificial satellite in accordance with the present invention is used within a service target area, and may be mounted on a mobile object moving within the service target area. Further, the send and receive apparatus may be equipped with a GPS means for measuring at least its own position by receiving an electromagnetic wave from a GPS satellite composing a global positioning system, or a measuring means for measuring a consumed amount of substances relating to charges for public services such as electricity, city gas and city water.  
         [0048]     Furthermore, in order to attain the above object, in the present invention, the satellite communication system for performing satellite communication through an artificial satellite comprises at least the artificial satellite, a satellite communication send and receive apparatus for performing satellite communication through the artificial satellite and a base station for performing communication with the satellite send and receive apparatus through the artificial satellite. The artificial satellite is an artificial satellite traveling on an elliptical orbit. The satellite communication send and receive apparatus is mountable on a mobile object, and comprises a send and receive means for performing sending and receiving of signals with the artificial satellite when the artificial satellite is used within a specified target service area.  
         [0049]     Still further, in order to attain the above object, in the present invention, the satellite communication system for performing satellite communication through an artificial satellite comprises at least the artificial satellite, a plurality of satellite communication send and receive apparatuses for performing satellite communication through the artificial satellite. The artificial satellite is an artificial satellite traveling on an elliptical orbit. Each of the plurality of satellite communication send and receive apparatuses comprises a send and receive means for performing sending and receiving of signals with the other satellite communication send and receive apparatuses through the artificial satellite. At least one of the plurality of satellite communication send and receive apparatuses is located within the above-described target service area though the others are located outside the target service area, and located at a position capable of performing satellite communication through the artificial satellite. Depending on an elevation angle of the artificial satellite when the artificial satellite is seen from the target service area, a relaying mode is selected among a relaying mode between the satellite send and receive apparatuses positioned within the target service area; a relaying mode between the satellite send and receive apparatus positioned within the target service area and the satellite send and receive apparatus positioned within the other area; and a relaying mode between the satellite send and receive apparatuses positioned at positions outside the target service area. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0050]      FIG. 1  is a diagram showing the long-term change of an artificial satellite orbit on which an artificial satellite of 24-hour orbital period travels without performing orbit control, the orbit being projected on a map of the world (isometric projection with respect to latitude and longitudinal measures).  
         [0051]      FIG. 2  is a flowchart showing a method of setting six orbital elements in accordance with the present invention.  
         [0052]      FIG. 3  is an explanatory diagram showing flow of information for controlling an orbit of an artificial satellite to the six orbital elements set by an algorithm in accordance with the present invention.  
         [0053]      FIG. 4  is an explanatory diagram showing flows of work and information performed in an artificial satellite tracking and controlling facility to control the orbit of the artificial satellite.  
         [0054]      FIG. 5  is an explanatory diagram showing flows of processing and information performed in an artificial satellite to control the orbit of the artificial satellite.  
         [0055]      FIG. 6  is an explanatory diagram showing orbits around the earth with respect to an example of orbit configuration using three artificial satellites obtained by an algorithm in accordance with the present invention.  
         [0056]      FIG. 7  is an explanatory diagram showing orbits around the earth with respect to Example 2 of orbit configuration obtained by an algorithm in accordance with the present invention.  
         [0057]      FIG. 8  is an explanatory view showing an example of a satellite broadcast system to which the present invention is applied.  
         [0058]      FIG. 9  is an explanatory diagram showing an example of a base station of a satellite broadcast system to which the present invention is applied.  
         [0059]      FIG. 10  is an explanatory diagram showing an example of a satellite broadcast terminal of a satellite broadcast system to which the present invention is applied.  
         [0060]      FIG. 11  is an explanatory view showing an example of a satellite broadcast system capable of also receiving ground broadcast to which the present invention is applied.  
         [0061]      FIG. 12  is an explanatory diagram showing an example of a ground broadcast station of a satellite broadcast system to which the present invention is applied.  
         [0062]      FIG. 13  is an explanatory diagram showing another example of a satellite broadcast terminal of a satellite broadcast system to which the present invention is applied.  
         [0063]      FIG. 14  is an explanatory view showing another example of a satellite broadcast system to which the present invention is applied.  
         [0064]      FIG. 15  is an explanatory diagram showing another example of a base station of a satellite broadcast system to which the present invention is applied.  
         [0065]      FIG. 16  is an explanatory diagram showing another example of a satellite broadcast terminal of a satellite broadcast system to which the present invention is applied.  
         [0066]      FIG. 17  is an explanatory view showing an example of a satellite communication system to which the present invention is applied.  
         [0067]      FIG. 18  is an explanatory diagram showing an example of a satellite communication system to which the present invention is applied.  
         [0068]      FIG. 19  is an explanatory diagram showing an example of a satellite communication system to which the present invention is applied.  
         [0069]      FIG. 20  is an explanatory view showing an example of a satellite communication system to which the present invention is applied.  
         [0070]      FIG. 21  is an explanatory diagram showing an example of a satellite communication system to which the present invention is applied.  
         [0071]      FIG. 22  is an explanatory view showing an example of a satellite communication system to which the present invention is applied.  
         [0072]      FIG. 23  is an explanatory diagram showing an example of a satellite communication system to which the present invention is applied.  
         [0073]      FIG. 24  is an explanatory view showing an example of a satellite communication system to which the present invention is applied.  
         [0074]      FIG. 25  is an explanatory diagram showing an example of a satellite communication system to which the present invention is applied.  
         [0075]      FIG. 26  is an explanatory view showing an example of a satellite communication system to which the present invention is applied.  
         [0076]      FIG. 27  is an explanatory diagram showing an example of a satellite communication system to which the present invention is applied.  
         [0077]      FIG. 28  is an explanatory view showing an example of a satellite communication system to which the present invention is applied.  
         [0078]      FIG. 29  is an explanatory diagram showing an example of a satellite communication system to which the present invention is applied.  
         [0079]      FIG. 30  is an explanatory view showing an example of a satellite communication system to which the present invention is applied.  
         [0080]      FIG. 31  is an explanatory diagram showing an example of a satellite communication system to which the present invention is applied.  
         [0081]      FIG. 32  is an explanatory diagram showing an example of a satellite communication system to which the present invention is applied.  
         [0082]      FIG. 33  is an explanatory view showing an example of a satellite-to-satellite communication system to which the present invention is applied.  
         [0083]      FIG. 34  is an explanatory view showing an example of an earth survey system to which the present invention is applied.  
         [0084]      FIG. 35  is a characteristic graph showing maximum coincidentally visible duration time for each combination of orbital inclination angle and eccentricity with respect to duration time in which an artificial satellite traveling on an orbit of 24-hour orbital period is visible above 70 degrees of elevation angle coincidentally from Nemuro, Sapporo, Sendai, Niigata, Tokyo, Nagoya, Kanazawa, Osaka, Hiroshima, Kochi, Fukuoka, Kagoshima and Naha.  
         [0085]      FIG. 36  is a detailed characteristic graph showing a part of  FIG. 35  where the maximum coincidentally visible duration time is above 6 hours and 45 minutes.  
         [0086]      FIG. 37  is a graph showing a computer simulation result of change with time of elevation angle at which an artificial satellite traveling on an orbit is visible, the computer simulation result being performed on a case where duration time in which the artificial satellite traveling on the orbit of 24-hour orbital period having orbital elements of a 42.5 degree orbital inclination angle and a 0.21 eccentricity is visible above 70 degrees of elevation angle coincidentally from Nemuro, Sapporo, Sendai, Niigata, Tokyo, Nagoya, Kanazawa, Osaka, Hiroshima, Kochi, Fukuoka, Kagoshima and Naha becomes maximum.  
         [0087]      FIG. 38  is a sky map showing a computer simulation result of visible direction at Naha of an artificial satellite traveling on an orbit, the computer simulation result being performed on a case where duration time in which the artificial satellite traveling on the orbit of 24-hour orbital period having orbital elements of a 42.5 degree orbital inclination angle and a 0.21 eccentricity is visible above 70 degrees of elevation angle coincidentally from Nemuro, Sapporo, Sendai, Niigata, Tokyo, Nagoya, Kanazawa, Osaka, Hiroshima, Kochi, Fukuoka, Kagoshima and Naha becomes maximum.  
         [0088]      FIG. 39  is a graph showing change of maximum coincidentally visible duration time depending on combination of eccentricity and orbital inclination angle when argument of perigee is changed under various combinations of the eccentricity giving the longest maximum coincidentally visible duration time in each orbital inclination angle and the orbital inclination angle in  FIG. 35  and  FIG. 36 .  
         [0089]      FIG. 40  is a diagram showing an artificial satellite orbit projected on a map of the world on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 210 degrees, the orbit map of the world is of isometric projection with respect to latitude and longitudinal measures.  
         [0090]      FIG. 41  is a diagram showing an artificial satellite orbit projected on a map of the world on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 230 degrees,. the orbit map of the world is of isometric projection with respect to latitude and longitudinal measures.  
         [0091]      FIG. 42  is a diagram showing an artificial satellite orbit projected on a map of the world on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 250 degrees, the orbit map of the world is of isometric projection with respect to latitude and longitudinal measures.  
         [0092]      FIG. 43  is a diagram showing an artificial satellite orbit projected on a map of the world on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 270 degrees, the orbit map of the world is of isometric projection with respect to latitude and longitudinal measures.  
         [0093]      FIG. 44  is a diagram showing an artificial satellite orbit projected on a map of the world on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 290 degrees, the orbit map of the world is of isometric projection with respect to latitude and longitudinal measures.  
         [0094]      FIG. 45  is a diagram showing an artificial satellite orbit projected on a map of the world on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 310 degrees, the orbit map of the world is of isometric projection with respect to latitude and longitudinal measures.  
         [0095]      FIG. 46  is a graph showing argument of perigee in a case where an elliptical orbit of a 24-hour orbital period intersects with a stationary orbit for each eccentricity of the elliptical orbit.  
         [0096]      FIG. 47  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 220 degrees in the case where service is provided 24 hours per day using three artificial satellites in the combination (42.5 degrees, 0.21).  
         [0097]      FIG. 48  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 230 degrees in the case where service is provided 24 hours per day using three artificial satellites in the combination (42.5 degrees, 0.21).  
         [0098]      FIG. 49  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 250 degrees in the case where service is provided 24 hours per day using three artificial satellites in the combination (42.5 degrees, 0.21).  
         [0099]      FIG. 50  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 270 degrees in the case where service is provided 24 hours per day using three artificial satellites in the combination (42.5 degrees, 0.21).  
         [0100]      FIG. 51  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 290 degrees in the case where service is provided 24 hours per day using three artificial satellites in the combination (42.5 degrees, 0.21).  
         [0101]      FIG. 52  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 310 degrees in the case where service is provided 24 hours per day using three artificial satellites in the combination (42.5 degrees, 0.21).  
         [0102]      FIG. 53  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 220 degrees in the case where service is provided 24 hours per day using four artificial satellites in the combination (42.5 degrees, 0.21).  
         [0103]      FIG. 54  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 230 degrees in the case where service is provided 24 hours per day using four artificial satellites in the combination (42.5 degrees, 0.21).  
         [0104]      FIG. 55  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 250 degrees in the case where service is provided 24 hours per day using four artificial satellites in the combination (42.5 degrees, 0.21).  
         [0105]      FIG. 56  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 270 degrees in the case where service is provided 24 hours per day using four artificial satellites in the combination (42.5 degrees, 0.21).  
         [0106]      FIG. 57  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 290 degrees in the case where service is provided 24 hours per day using four artificial satellites in the combination (42.5 degrees, 0.21).  
         [0107]      FIG. 58  is a chart showing a simulation result expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 310 degrees in the case where service is provided 24 hours per day using four artificial satellites in the combination (42.5 degrees, 0.21).  
         [0108]      FIG. 59  is a graph showing change in eccentricity obtained by simulation of long-term orbit prediction over three years when orbit control is completely performed in a case of the combination (42.5 degrees, 0.21) taking the time 0:00:000 (UTC) on Oct. 1, 2001 as the reference time.  
         [0109]      FIG. 60  is a graph showing change in orbit inclination angle obtained by simulation of long-term orbit prediction over three years when orbit control is completely performed in a case of the combination (42.5 degrees, 0.21) taking the time 0:00:000 (UTC) on Oct. 1, 2001 as the reference time.  
         [0110]      FIG. 61  is a graph showing change in argument of perigee obtained by simulation of long-term orbit prediction over three years when orbit control is completely performed in a case of the combination (42.5 degrees, 0.21) taking the time 0:00:000 (UTC) on Oct. 1, 2001 as the reference time.  
         [0111]      FIG. 62  is a graph showing change in right ascension of north-bound node obtained by simulation of long-term orbit prediction over three years when orbit control is completely performed in a case of the combination (42.5 degrees, 0.21) taking the time 0:00:000 (UTC) on Oct. 1, 2001 as the reference time.  
         [0112]      FIG. 63  is a simulation map illustrating service with artificial satellites in the United States. 
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS  
       [0113]     The following embodiments in accordance with the present invention will be described below.  
         [0000]     Method of setting the orbital elements (algorithm)  
         [0000]     Examples of set values of the orbital elements and orbit configuration by the algorithm  
         [0000]     Method of realizing and controlling the set orbital elements  
         [0000]     Systems employing an artificial satellite traveling on the orbit according to the present invention  
         [0000]     (1) Method of setting the orbital elements (algorithm)  
         [0114]     Corresponding to each of the above items (a), (b) and (c), description will be made on the methods with regard to the items (a) and (b) below. With regard to the problem (c), setting of an argument of perigee is solved by the item (a) and a method of arranging a plurality of artificial satellites is solved by the item (b). Therefore, initially, numerical ranges of an orbital elements suitable for mobile broadcast and communication using the group of artificial satellites in Japan will be described here.  
         [0115]     Combinations of orbital elements for an orbit of 24 hour orbital period is investigated by taking an evaluation index of time duration that an artificial satellite traveling on the orbit can be seen in an elevation angle above 70 degrees from 13 cities of Nemuro, Sapporo, Sendai, Tokyo, Nagoya, Kanazawa, Osaka, Hiroshima, Kochi,. Fukuoka, Kagoshima and Naha at the same time.  
         [0116]     An orbit of an artificial satellite can be uniquely determined by giving a position and a velocity of the artificial satellite at a certain time. Therefore, by giving six orbital elements at a certain time, an orbit of an artificial satellite can be uniquely determined. Here, Kepler orbital elements are used for the method of describing the orbital elements. The Kepler orbital elements are composed of six orbital elements, that is, of a semi-major axis expressing size of an ellipse, an eccentricity expressing oblateness of an ellipse, an orbit inclination angle expressing inclination of an orbital plane, a right ascension of north-bound node indicating an angle between a straight line connecting between an ascending node where the orbit passes an equatorial plane from south to north and the vernal equinoctial direction, an argument of perigee indicating an angle from an ascending node to the perigee, and a true anomaly of an angle measuring a position on the orbit of the artificial satellite at a certain time with respect to a geocentric position as a center. An average anomaly or an eccentric anomaly may be used instead of the true anomaly.  
         [0117]     Since the orbital period is 24 hours, the orbital semi-major axis of one of the orbital elements is given by the orbital period. Next, the argument of perigee of one of the orbital elements is foxed to 270 degrees. In an orbit having the semi-major axis and the argument of perigee, a combination of two orbital elements of an orbit inclining angle and an eccentricity is considered. By doing so, a shape of the orbit projected on the ground is uniquely determined. If the other remaining orbital elements of aright ascension of north-bound node and a true anomaly are determined to a reference time, a position in the longitudinal direction of the orbit projected on the ground is determined. At that time, a duration time in which the artificial satellite is visible above an elevation angle of 70 degrees from the above-described 13 cities at a time can be calculated. Therefore, by successively changing the combination of two orbital elements of an orbit inclining angle and an eccentricity, the orbit projected on the ground is shifted in the longitudinal direction. By doing so, a duration time in which the artificial satellite is visible above an elevation angle of 70 degrees from the above-described 13 cities to a combination of two orbital elements of an orbit inclining angle and an eccentricity can be successively calculated. By comparing these results, a maximum value of the duration time in which the artificial satellite is visible above an elevation angle of 70 degrees from the above-described 13 cities to a combination of two orbital elements of an orbit inclining angle and an eccentricity can be finally calculated.  
         [0118]      FIG. 35  shows maximum coincidentally visible duration time given by a combination of an orbit inclination angle and an eccentricity through the above-described procedure when the orbit inclination angle is changed within a range of from 35 degrees to 44 degrees and the eccentricity is changed within a range of from 0.0 to 0.35. It can be understood from  FIG. 35  that, for example, when the orbit inclination angle is 35 degrees and the eccentricity is 0.2, the maximum coincidentally visible duration time becomes approximately 6 hours.  
         [0119]      FIG. 36  shows a portion of the above figure in detail where the maximum coincidentally visible duration time is above 6 hours and 45 minutes and the eccentricity range of from 0.09 to 0.25.  
         [0120]     It can be understood from  FIG. 36  that the maximum value among the maximum coincidentally visible duration time is given by the combination of the orbit inclination angle of 42.5 and the eccentricity of 0.21, and that the maximum coincidentally visible duration time in that condition is longer than 8 hours.  FIG. 37  is a graph showing change with time of elevation angle when the artificial satellite is seen from the above-described 13 cities. Since the time when the elevation angle becomes above 70 degrees is latest at Sapporo and the time when the elevation angle becomes below 70 degrees is earliest at Nemuro, the time difference between the two time points is larger than 8 hours.  
         [0121]     As shown in  FIG. 37 , the elevation angle at Naha changes in such a manner that the elevation angle increases, then decrease, after that again increases and finally decreases.  FIG. 38  is a sky map showing a computer simulation result of visible direction at Naha of the artificial satellite in this case.  FIG. 38  may be seen in a same manner as a star chart, and the center of the concentric circles is the zenith, the top directs to the north, the right hand side directs to the west, the bottom direct to the south and the left hand side directs to the east. The concentric circles are elevation angle and drawn at intervals of 20 degrees. The plotted dots express positions of the artificial satellite in the sky at intervals of 1 hour, and the line connecting the plotted dots is the orbit in the sky.  
         [0122]     It can be understood from  FIG. 38  that the visible direction of the artificial satellite is shifted from the zenith direction toward the horizon direction in the north, and from this the change with time of the elevation angle in  FIG. 37  can be explained. The changes with time of the elevation angle at the other cities in  FIG. 37  can be explained using similar star charts. The reason why the maximum coincidentally visible duration time is shortened at the orbit inclination angle of 44 degrees in  FIG. 35  and  FIG. 36  is that the visible direction of the artificial satellite at Naha is shifted excessively toward the horizon direction so that the elevation angle once becomes below 70 degrees and then is returned toward the direction of the elevation angle above 70 degrees. Therefore, the coincidentally visible duration time is shortened.  
         [0123]     In a case where 24-hour service is performed using three artificial satellites, the combination of the orbit inclination angle and the eccentricity which makes the maximum coincidentally visible duration time above 8 hours is preferable. In a case where 24-hour service is performed using four artificial satellites, the combination of the orbit inclination angle and the eccentricity which makes the maximum coincidentally visible duration time above 6 hours is preferable. The combination of the orbit inclination angle and the eccentricity may be referred to  FIG. 35  and  FIG. 36 .  
         [0124]     In a case where broadcast service is performed using a plurality of artificial satellites traveling on elliptical orbits, the artificial satellite to perform the broadcast service needs to be successively switched. In a case where performing mobile broadcast and communication is performed as described above, it is important that the artificial satellite becomes visible at a high elevation angle from a plurality of cities at a time.  
         [0000]     (a) Method of setting an argument of perigee  
         [0125]     Here, it is assumed that the other orbital elements of the semi-major axis, the eccentricity and the orbit inclination angle have been set by a method (b) to be described later.  
         [0126]     In a case where a specified area such as the territory of Japan is set to a service area, when a right ascension of north-bound node at a certain reference time is set, the argument of perigee is given with a certain range in order that the artificial satellite travels above the service area.  
         [0127]     When an argument of perigee is set, the argument of perigee is set by computer simulation using the orbital element and the reference time as the input condition.  
         [0128]     As a known example, in a case where the orbit inclination angle is 40 degrees and the eccentricity is 0.24 in a 24-hour orbital period, a result of computer simulation of change in the orbital elements over 10 years is shown in Table 1. The duration of 10 years is determined by summing the above-described mission lifetime is 10 years. The simulation result is based on a condition that the orbit correction control is not performed for the 10 years, in detail, the gas jet unit mounted on the artificial satellite is not jetted at all for the 10 years.  
                                                                                     TABLE 1                                               Right                                   Orbit Incl.   Ascension of   Argument of               Time   Semimajor       Angle   north-bound   perigee   True Anomaly       Time   Date   (UTC)   axis (km)   Eccentricity   (degrees)   node (degrees)   (degrees)   (degrees)                                initial value   Feb. 1, 2000   0:00:00   42,164   0.240   40.0   357.5   270.0   0.0        365 days after   Jan. 31, 2001   0:00:00   42,183   0.239   39.8   352.1   278.0   271.5        730 days after   Jan. 31, 2002   0:00:00   42,158   0.237   39.6   346.7   286.0   236.5       1095 days after   Jan. 31, 2003   0:00:00   42,149   0.235   39.3   341.5   294.5   310.0       1460 days after   Jan. 31, 2004   0:00:00   42,175   0.234   38.9   336.3   303.0   316.3       1825 days after   Jan. 30, 2005   0:00:00   42,174   0.233   38.5   331.1   311.7   231.0       2190 days after   Jan. 30, 2006   0:00:00   42,147   0.233   38.0   326.0   320.7   246.2       2555 days after   Jan. 30, 2007   0:00:00   42,162   0.234   37.5   320.8   329.8   313.2       2920 days after   Jan. 30, 2008   0:00:00   42,183   0.236   36.9   315.5   339.0   247.2       3285 days after   Jan. 29, 2009   0:00:00   42,157   0.239   36.3   310.2   348.2   216.2       3650 days after   Jan. 29, 2010   0:00:00   42,157   0.242   35.5   304.6   357.4   278.6                  
 
         [0129]      FIG. 1  shows the orbits projected on the ground when the orbital elements shown in Table 1 are employed. Each of the orbits projected on the ground over one day is calculated based on the orbital elements of the initial values, the values on 1095 days after, the values on 2190 days after, and the values on 3650 days after in Table 1.  
         [0130]     The followings can be understood from the orbits projected on the ground of the initial value to the value on 3650 days after. In the figure, the reference character 1 is the orbit projected on the ground over one day from the reference time, the reference character 2 is the orbit projected on the ground over one day from 1095 day after to 1096 days after, the reference character 3 is the orbit projected on the ground over one day from 2190 day after to 2191 days after, the reference character 4 is the orbit projected on the ground over one day from 3650 day after to 3651 days after. 
        The orbit projected on the ground is shifting toward the west side with time until 2190 days after and then shifting toward the east side with time until 3650 days after.     The orbit projected on the ground is slanting with time until 2190 days after and the north side end becomes sharp.        
 
         [0133]     The arrival range of the orbit projected on the ground in the latitudinal direction is narrowed with time.  
         [0134]     With regard to the shifting toward the west side, the orbit projected on the ground can be moved to the sky over Japan by moving back the orbit projected on the ground toward the east side by performing orbit correction to adjust the true anomaly so as to correspond to the right ascension of north-bound node. This orbit correction is the same method as the longitude control of the orbit of a stationary satellite. Therefore, this control can be performed by jetting the gas jet of the artificial satellite at three positions of the perigee, the apogee and the perigee in the last. At that time, the semi-major axis and the eccentricity as well as the true anomaly may be corrected. By the control of the correction of the eccentricity, the sharpened north side end of the orbit projected on the ground described in the above second problem can be corrected to be returned to the original form.  
         [0135]     The shortening of the arrival range in the latitudinal direction described in the above third problem is caused by reduction in the orbit inclination angle, and can be corrected by the method similar to the latitude control of the orbit of a stationary satellite. In detail, the propellant is jetted by the gas jet unit in the direction normal to the orbital plane when the artificial satellite passes through the equatorial plane on the orbit.  
         [0136]     The above two kinds of control may be periodically, for example, every 30 days or ever 60 days, performed. By the periodical control, the semi-major axis, the eccentricity and the orbit inclination angle can be controlled to the nominal values or the values near the values at setting the orbital element.  
         [0137]     A problem is the phenomenon described in the second that the orbit projected on the ground is slanting with time. This phenomenon is caused by change in the argument of perigee,. and not observed in the orbit of a stationary satellite.  
         [0138]     As shown in Table 1, according to the result of the computer simulation described above, the argument of perigee changes by approximately 90 degrees from the initial value of 270 degrees to the value nearly 10 years after of 357.4 degrees. In a case where the change in the argument of perigee is not allowed and the gas jet unit for controlling the argument of perigee is jetted similarly to the orbit control described above, the control is performed by jetting the gas jet unit in the direction parallel to the orbital plane at the time when the artificial satellite passes just before the ascension node or just after the descending node through the equatorial plane on the orbit, but a large amount of the propellant is consumed. Table 2 shows a simulation result on an amount of acceleration for the orbit control required when the periodic orbit control id performed in a case where four artificial satellites are arranged on four orbits each separated by 90 degrees in the right ascension of north-bound node. This is an amount of acceleration required in one year.  
                                                                   TABLE 2                           (unit: m/s)                Satel-   Satel-   Satel-   Satel-           Kind of Control   lite 1   lite 2   lite 3   lite 4   Remarks                    Control of True   94.11   25.25   75.76   42.12   Control       Anomaly,                   Frequency = 60       Semimajor Axis,                   days       Eccentricity       Control of   74.59   146.56   1.03   105.82   Control       argument of                   Frequency = 60       perigee                   days       Control of Orbit   5.88   21.10   17.98   34.24   Control       Inclination                   Frequency = 60       angle                   days       Total   174.58   192.91   94.77   182.18                  
 
         [0139]     It can be understood from Table 2 that the acceleration amount required for control of the argument of perigee becomes approximately ¾ of the total acceleration amount for the orbit control at the maximum though it depends on positions of the orbit arrangement. Since an amount of propellant consumed by the gas jet unit to jet for the orbit control is inevitably increased, devices mountable on the artificial satellite must be reduced, or on-orbit lifetime of the artificial satellite must be shortened, in the worst case.  
         [0140]     Therefore, in the present invention, change in the argument of perigee is allowed in advance, and the value is set with a certain allowance in the setting process of the argument of perigee.  
         [0141]     If change of 90 degrees can be allowed in the above example, and when the initial value of the argument of perigee is set to, for example, 235 degrees, it can be predicted that the argument of perigee at the mission lifetime of 10 years after will become nearly 325 degrees. If the nominal value at that time is assumed to be 270 degrees, the argument of perigee can be maintained to a value in a range of the nominal value ±45 degrees without control, and the orbit projecting on the ground and accordingly the visibility of the artificial satellite can be maintained.  
         [0142]     By setting the orbital elements on the premise that the argument of perigee is not controlled, the amount of propellant to be mounted on the artificial satellite can be substantially reduced.  
         [0143]     In a case where change of 90 degrees can not be allowed, there is a method that when the mission lifetime is 10 years, the argument of perigee is controlled at the time, for example, 5 years after starting of the mission to return the argument of perigee to the initial value. In the example described in Table 1, the argument of perigee about 5 years after is nearly 312 degrees, and changes by nearly 42 degrees from the initial value of 270 degrees. At that time, if the initial value is set to 249 degrees, the argument of perigee is estimated to be changed to nearly 291 degrees at the time 5 years after. If the nominal value at that time is assumed to be 270 degrees, the argument of perigee can be maintained to a value in a range of the nominal value ±21 degrees without control, and the orbit projecting on the ground and accordingly the visibility of the artificial satellite can be maintained. Further, it can be also considered that the mission lifetime can be lengthened by again returning the argument of perigee to the initial value 5 years after.  
         [0144]     Although the time correcting the argument of perigee is 5 years after starting the mission in the above example, the correction time may be set to the time 1 year after, 3 years after and so on depending on the allowance of the argument of perigee.  
         [0145]     Further, there is a control method that in a case of the mission lifetime of 10 years, an initial value is set depending on an allowable width of the argument of perigee, and the argument of the perigee is returned to the initial value at the time when it reaches the limit value of the allowable width of the argument of perigee. For example, in the example of Table 1, since the argument of perigee at about 6 years after is nearly 321 degrees and changes by nearly 51 degrees from the initial value, there can be considered a control method that the initial value of the argument of perigee is set to 255 degrees, and the argument of perigee is returned to the initial value 6 years after in order to maintain the argument of perigee within a range of the nominal value of 270 degrees ±25 degrees. At that time, during the remaining mission lifetime of 4 years, the argument of perigee changes by a value corresponding to 4 years from the initial value.  
         [0146]     A detailed example will be described below.  
         [0147]     Description has been made in the beginning of the item (1) referring to  FIG. 35  and  FIG. 36  on the preferable orbital elements for a case where broadcast and communication service to a mobile object in Japan is performed using a plurality of artificial satellites. From  FIG. 36 , a combination of an orbit inclination angle and an eccentricity giving maximum coincidentally visible duration time from the 13 cities can be obtained for each orbit inclination angle. In detail, when the orbit inclination angle is 40 degrees, the maximum coincidentally visible duration time becomes longest at the eccentricity of 0.16.  
         [0148]     When such a combination is expressed by (40 degrees, 0.16), combinations (40 degrees, 0.18), (42 degrees, 0.20), (42.5 degrees, 0.21), (43 degrees, 0.22) and (43.5 degrees, 0.24) are obtained. When the argument of perigee is changed to these combinations, the artificial satellite becomes visible with the elevation angle larger than 70 degrees coincidentally from the above-mentioned 13 cities.  FIG. 39  shows how the coincidentally visible duration time changes.  
         [0149]     In the case of the combination of, for instance, (42.5 degrees, 0.21), the maximum coincidentally visible duration time is longer than 8 hours when the argument of perigee is within a range from nearly 223 degrees to 270 degrees, and the maximum coincidentally visible duration time monotonously decreases when the argument of perigee is larger than 270 degrees. It can be understood that good service can be provided for long time also in the cases of the other combinations when the lower limit of the argument of perigee is set to a value from 220 degrees to 230 degrees and the upper limit of the argument of perigee is set to a value of 270 degrees. However, in a case where 24-hour service is provided using three artificial satellites, the setting width is selected from a range of argument of perigee in which the maximum coincidentally visible duration time is longer than 8 hours. In a case where 24-hour service is provided using four artificial satellites, the setting width is selected from a range of argument of perigee in which the maximum coincidentally visible duration time is longer than 6 hours. Therefore, in this case, there is no need to set the lower limit of the argument of perigee to a value from 220 degrees to 230 degrees.  
         [0150]     As references, in the case of the combination(42.5 degrees, 0.21),  FIG. 40 ,  FIG. 41 ,  FIG. 42 ,  FIG. 43 ,  FIG. 44  and  FIG. 45  show orbits projected on the ground giving the maximum coincidentally visible duration time when the argument of perigee is 210 degrees, 230 degrees, 250 degrees, 270 degrees, 290 degrees and 310 degrees, respectively.  
         [0151]     What should be noted here is that an intersect point of the orbit of the artificial satellite with a stationary satellite orbit may appear at a certain argument of perigee because the orbit of the artificial satellite rotates within the orbital plane due to change in the argument of perigee.  FIG. 46  is a graph showing argument of perigee in a case where the orbit of a 24-hour orbital period intersects with a stationary orbit for each eccentricity. As shown in  FIG. 46 , when the perigee exists in the sky of the southern hemisphere, there are two cases where the orbit in the ascending node side intersects with the stationary satellite orbit and where the orbit in the descending node side intersects with the stationary satellite orbit. Similarly, when the perigee exists in the sky of the northern hemisphere, there are two cases. In the case where the combination of argument of perigee is (42.5 degrees, 0.21), the orbit intersects with the stationary satellite orbit at the argument of perigees of 257.9 degrees and 282.1 degrees since the eccentricity is 0.21, as shown in  FIG. 46 . In actual operation of the artificial satellite, the argument of perigee is changed by performing orbit control just before intersecting with the stationary satellite orbit so as to prevent the orbit of the artificial satellite from intersecting with the stationary satellite orbit.  
         [0152]     As references, in the case where service is provided 24 hours per day using three artificial satellites in the combination (42.5 degrees, 0.21),  FIG. 47  to  FIG. 52  show examples of simulation results expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 220 degrees, 230 degrees, 250 degrees, 270 degrees, 290 degrees and 310 degrees, respectively. Similarly, in the case where service is provided 24 hours per day using four artificial satellites in the combination (42.5 degrees, 0.21),  FIG. 53  to  FIG. 58  show examples of simulation results expressed by contour lines of time ratio (%) in which any one of the artificial satellites is visible at an elevation angle of above 70 degrees when the argument of perigee is 210 degrees, 230 degrees, 250 degrees, 270 degrees, 290 degrees and 310 degrees, respectively. The area surrounded by the line “100” indicates an area where the time ratio is 100%, and the area surrounded by the line “90” indicates an area where the time ratio is 90%. By setting the argument of perigee to a value with the allowance as described above, it is possible to moderate the requirement of the orbit control of the artificial satellite as described above, and to provide satellite communication and broadcast service with a high elevation angle to the almost area in the territory of Japan.  
         [0153]      FIG. 59  to  FIG. 62  show a simulation result of long-term orbit prediction over three years when orbit control is completely performed in a case of the combination (42.5 degrees, 0.21) taking the time 0:00:000 (UTC) on Oct. 1, 2001 as the reference time. The initial values of the orbital elements are an orbital period of 24 hours, an orbit inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 270 degrees.  FIG. 59  shows the long-term change in the eccentricity.  FIG. 60  shows the long-term change in the orbit inclination angle.  FIG. 61  shows the long-term change in the argument of perigee.  FIG. 62  shows the long-term change in the right ascension of north-bound node. The abscissa in the figures expresses initial value of right ascension of north-bound node. It can be understood that the changes in the orbital elements depend on the initial value of the right ascension of north-bound node. Since the change in the argument of perigee depends on the initial value of the right ascension of north-bound node, an initial value of the argument of perigee may be set in taking the change in the initial value of the right ascension of north-bound node into consideration. For example, when the initial value of the right ascension of north-bound node is around 190 degrees, change in the argument of perigee is hardly observed over a long time. Therefore, by setting the initial value of the right ascension of north-bound node one of the artificial satellites to 190 degrees, the orbit control can be simplified.  
         [0000]     (b) Method of arranging orbits for a plurality of artificial satellites  
         [0154]     Here, a method of setting the orbital elements in accordance with the present invention will be successively described.  FIG. 2  is a flow chart showing the setting method.  
         [0155]     In a case where service is performed concentratingly and continuously to a specified area using a plurality of artificial satellites, orbits projecting on the ground just below the individual satellites must agree with one another. In order to satisfy this condition, the semi-major axis, the eccentricity, the orbit inclination angle and the argument of perigee out of the orbital elements are nearly equal among the orbits. Therefore, in the following processes, the semi-major axis, the eccentricity, the orbit inclination angle and the argument of perigee are commonly set as the orbital elements for all of the artificial satellites, and the right ascension of north-bound node at a certain reference time and the argument of perigee are individually set for each artificial satellites.  
         [0000]     (i) Setting of a reference time (reference character 5)  
         [0156]     A reference time (epoch) for defining the six orbital element of the artificial satellite is set.  
         [0000]     (ii) Setting of number n of the artificial satellites (reference character 6)  
         [0157]     Number n (n is a positive integer) of the artificial satellites is set.  
         [0000]     (iii) Setting of a temporary value of the argument of perigee ω (reference character 7)  
         [0158]     An arbitrary argument of perigee ω is given as a temporary value.  
         [0159]     In a case where the service target area is a specified area in the northern hemisphere, if it is preferable for performing communication and broadcast service that the apogee of the artificial satellite orbit is placed in the sky above the specified area, it is often advantageous that the nominal value of the argument of perigee is set to 270 degrees. If it is preferable for observing a central celestial body of the artificial satellite traveling around that the perigee of the artificial satellite orbit is placed in the sky above the specified area, it is often advantageous that the nominal value of the argument of perigee is set to 90 degrees. In a case where the service target area is a specified area in the southern hemisphere, on the contrary, in the former case, it is often advantageous that the nominal value of the argument of perigee is set to 90 degrees. In the latter case, it is often advantageous that the nominal value of the argument of perigee is set to 270 degrees.  
         [0160]     Further, as having been described in the item (a), the set value of the argument of perigee may be set with the allowance.  
         [0000]     (iv) Setting of a temporary value of the semi-major axis a (reference character 8)  
         [0161]     In a case where service is performed concentratingly and continuously to a specified area using a plurality of artificial satellites, orbits projecting on the ground just-below the individual satellites must agree with one another. Further, each of the artificial satellites must travel in the sky above the same point on the ground every day. That is, the orbit projected on the ground must be fixed for a long time irrespectively of elapsing of time. In order to satisfy this requirement, the artificial satellite must travel around a central celestial body integer times while the central celestial body of the artificial satellite traveling around rotates one turn. When the central celestial body is the earth, a range of number m of traveling-around times satisfies the relation 1≦m≦16 (m is an integer), and the present invention can be applied to a system using a plurality of artificial satellites traveling on an orbit having an orbital period shown in Table 3.  
         [0162]     By the number m of traveling-around times, in a case where the central celestial body is the earth, an orbital period p (unit: hour) of the artificial satellite can be calculated by the following equation. 
 
 p= 23.93/m 
 
         [0163]     By the orbital period p, an orbit semi-major axis a of the artificial satellite can be determined as shown in Table 3. In setting the orbital elements to be described below, a candidate is selected from the semi-major axes described in Table 3. However, at the selection, in a case where the central celestial body is the earth, service can not be continuously supplied to the specified area over 24 (hours/day) unless at least the following relation is satisfied. 
 
 p×n≧ 23.93 
 
         [0164]                                                              TABLE 3                                       Number of                       Traveling-       Semi-major               around times   Orbital Period   Axis           No.   per day   (hour)   (km)                                        1   1   23.93   42,164           2   2   11.97   26,562           3   3   7.98   20,270           4   4   5.98   16,733           5   5   4.79   14,420           6   6   3.99   12,770           7   7   3.42   11,522           8   8   2.99   10,541           9   9   2.66   9,745           10   10   2.39   9,084           11   11   2.18   8,525           12   12   1.99   8,044           13   13   1.84   7,626           14   14   1.71   7,259           15   15   1.60   6,932           16   16   1.50   6,640                        
 (v) Setting of a temporary value of the eccentricity e (reference character 9) 
 
         [0165]     An arbitrary eccentricity e is given as a temporary value.  
         [0166]     For example, a service duration time required for one of artificial satellites in the specified area is let be Ts (unit: second). Letting time length required the artificial satellite travels from the perigee to a point starting the service of the orbit on which the artificial satellite travels be Ti, an eccentric anomaly at the point starting the service be Ei (unit: radian), and the true anomaly be θi (unit: radian), the following equations can be obtained. 
 
 Ti= ( Ei−e ×sin  Ei )× p /(2×π) 
 
cos  Ei= ( e +cos θ i )/1 +e ×cos θ i ) 
 
 Therefor, assuming that the point stopping the service on the orbit is placed at a point symmetric to the point starting the service with respect to the major axis of the orbit, the following relation can be obtained. 
 
 Ts≦p−Ti× 2 
 
∴( p−Ti× 2)− Ts≧ 0 
 
 Since it is necessary to take rotation of the celestial body of the artificial satellite traveling around into consideration, an eccentricity satisfying the above relation is given as an initial value in taking a combination of the eccentric anomaly Ei at the point starting the service and the eccentricity e into consideration. At considering the combination, it is convenience to take it into consideration that the minimum value of the eccentric anomaly is π/2 and the maximum value is π. 
 
 (vi) Setting of a temporary value of the orbit inclination angle (reference character 10) 
 
         [0167]     An arbitrary orbit inclination angle i is given as a temporary value.  
         [0168]     In that time, it is convenience that initially the maximum value and the minimum value of longitudes of the specified service target area are obtained, and then the average value is given as the initial value of the orbit inclination angle.  
         [0000]     (vii) Setting of a temporary value of the true anomaly  
         [0169]     When the satellite having satellite number  1  is at the perigee, the true anomaly θk (unit: radian) for satellite number k(1≦k≦n, k is an integer) is given by the following equation. Since the eccentric anomaly corresponds to the true anomaly one-to-one, the eccentricity anomaly is initially calculated and then the true anomaly is calculated. 
 
−23.93×( k− 1)/ n× 3600=( Ek−e× sin  Ek )× P /(2×π)cos θ I= ( e− cos  Ei )/( e× cos  Ei− 1) 
 
 By the combination of the right ascension of north-bound node Ωk and the true anomaly θk described above, the orbits projected on the ground of the plurality of artificial satellites agree with one another, and the n number of satellites from satellite number  1  to satellite number n sequentially draw the single orbit projected on the ground. That is, the n number of satellites from satellite number  1  to satellite number n repetitively travel in the sky above the specified area. 
 
         [0170]     (viii) Setting of temporary values of the right ascension of north-bound node Ω1 and the true anomaly θ1 of the artificial satellite of satellite number 1 (reference character 11)  
         [0171]     In order to perform service to the specified area, the individual artificial satellites must travel in the sky above the specified area. Therefore, the right ascension of north-bound node Ω1 and the true anomaly θ1 of the artificial satellite of satellite number  1  at the reference time are set so as to travel in the sky above the specified area. At that time, the right ascension of north-bound node Ω1 can be easily set by performing simulation using a computer by setting the true anomaly θ1=0 (degree).  
         [0000]     (ix) Setting of temporary values of the right ascension of north-bound node Ωk and the true anomaly θk of the artificial satellite of satellite number k (reference character 12)  
         [0172]     By the right ascension of north-bound node Ω1 and the true anomaly θ1 set for the artificial satellite of satellite number  1 , the right ascension of north-bound node Ωk and the true anomaly θk of the artificial satellite of satellite number k at the reference time is sequentially calculated from k=2 to k=n. In detail, when the right ascension of north-bound node Ω1 (unit: radian) of the orbit of the artificial satellite of satellite number  1  is taken as the reference, the right ascension of north-bound node Ωk (unit: radian) of the orbit of the artificial satellite of satellite number k (1≦k≦n) is given by the following equation. 
 
Ω k=Ω 1+( k− 1)/ n× 360 
 
         [0173]     This equation is for arranging the plurality of artificial satellites on the orbital plane spaced in an equal angle.  
         [0000]     (x) Evaluation (reference character 14)  
         [0174]     It is evaluated whether or not requirements are satisfied by the group of artificial satellites traveling on the orbit determined by the orbital elements given above. For example, the requirements with regard to contents of the service in the specified area are as follows. 
        Time duration in which one artificial satellite can continuously perform the service to the specified area.     Time duration in which one artificial satellite is visible in the sky in a high elevation angle when it is seen from the specified area.     Distance between the artificial satellite and the specified area, and change with time of the distance.     Delay time of the electromagnetic wave propagation.     Doppler shift of the electromagnetic wave, and line design.     Spatial resolution when the specific area is observed. The requirements with regard to control of the orbit of artificial satellite are as follows.     Long-term change of each of the orbital elements of each orbit.     Control amount of each of the orbital elements of each orbit, and an amount of propellant necessary for the control.        
 
         [0183]     The evaluation can be easily performed by performing computer simulation using the above-set orbital elements as input values. Items and requirement necessary for the evaluation are set usually before examining the orbit arrangement.  
         [0000]     (xi) Repeating of process of setting the orbital elements  
         [0184]     When the requirements are not satisfied in the above evaluation, the processes from the step (iii) to the step (x) are repeated. The processes from the step (i) to the step (x) are repeated, if necessary. In a case where the temporary value needs not to be reviewed, the corresponding item needs not to be reviewed. Therein, the order of the above-described processes may be arbitrarily changed depending on necessity.  
         [0185]     In the case where the orbital elements are set through the above-described method, the artificial satellites from the artificial satellite of satellite number  1  to the artificial satellite of satellite number n sequentially appear in the sky above the specified target area to be supplied with the service by the group of artificial satellites. Further, if the right ascension of north-bound node set in the step (ix) is determined using the following equation 
 
Ω k=Ω 1−( k− 1)/ n× 360(1 ≦k≦n, k  is an integer), 
 
 and the eccentric anomaly set in the step (vii) is determined using the following equation 
 
23.93×( k− 1)/ n× 3600=( Ek−e ×sin  Ek )× P /(2×π), 
 
 the similar result can be obtained. In this case, the artificial satellites from the artificial satellite of satellite number n to the artificial satellite of satellite number  1  sequentially appear in the sky above the specified target area to be supplied with the service by the group of artificial satellites. 
 
 (xii) In a case where the requirements are satisfied in the above evaluation, the final orbital elements of each of the artificial satellites at the reference time are obtained. 
 
 (2) Examples of the values of the orbital elements and the orbit arrangement set by the above-described algorithm 
 
         [0186]     Examples of the values of the orbital elements and the orbit arrangement set by the above-described algorithm will be described below.  
         [0187]     As to be described later, an orbit of an artificial satellite is always changed by the effects of the gravitational field of the earth and the attractive forces of the moon and the sun, and generally orbit-controlled with an allowable range to a certain degree. Therefore, each value of the orbital elements indicates a target nominal value after orbit-controlled.  
         [0188]     In the following tables, Ω1 and θ1 are the right ascension of north-bound node and the true anomaly of the artificial satellite of satellite number  1  set corresponding to the reference time.  
         [0189]     Table 4 and Table 5 show examples of the orbital elements and the orbit arrangement for a satellite communication and broadcast network by three artificial satellites traveling on an orbit of 24-hour orbital period. The eccentricity and the orbit inclination angle may be a combination within the range shown in  FIG. 26  and  FIG. 27 . The argument of perigee may be smaller than 180 degrees.  
         [0190]     The example of the orbit arrangement covers the whole territory of Japan as the target service area.  
                                                                                     TABLE 4                                   Item   Value                                        Satellite No.   1   2   3                Orbital Period (hour)   24           Semimajor Axis (km)   approx. 42,164           Eccentricity   not larger than 0.24           Orbit Inclination   larger than 37 degrees and           Angle (degrees)   smaller than 44 degrees           Argument of Perigee   larger than 180 degrees and           (degrees)   smaller than 360 degrees                Right Ascension   Ω1   Ω1 + 120   Ω1 + 240           of North-Bound           node (degrees)           True Anomaly (degrees)   θ1   an angle ⅓   an angle ⅓                   of orbital   of orbital                   period   period                   behind θ1   ahead of θ1                      
 
         [0191]    
       
         
               
               
               
             
               
               
               
               
               
             
               
               
               
             
               
               
               
               
               
             
           
               
                   
                 TABLE 5 
               
               
                   
                   
               
               
                   
                   
               
               
                   
                 Item 
                 Value 
               
               
                   
                   
               
             
             
               
                   
               
             
          
           
               
                   
                 Satellite No. 
                 1 
                 2 
                 3 
               
             
          
           
               
                   
                 Orbital Period (hour) 
                 24 
               
               
                   
                 Semimajor Axis (km) 
                 approx. 42,164 
               
               
                   
                 Eccentricity 
                 not larger than 0.24 
               
               
                   
                 Orbit Inclination 
                 larger than 40 degrees and 
               
               
                   
                 Angle (degrees) 
                 smaller than 44 degrees 
               
               
                   
                 Argument of Perigee 
                 larger than 180 degrees and 
               
               
                   
                 (degrees) 
                 smaller than 360 degrees 
               
             
          
           
               
                   
                 Right Ascension 
                 Ω1 
                 Ω1 + 120 
                 Ω1 + 240 
               
               
                   
                 of North-Bound 
               
               
                   
                 node (degrees) 
               
               
                   
                 True Anomaly (degrees) 
                 θ1 
                 an angle ⅓ 
                 an angle ⅓ 
               
               
                   
                   
                   
                 of orbital 
                 of orbital 
               
               
                   
                   
                   
                 period 
                 period 
               
               
                   
                   
                   
                 behind θ1 
                 ahead of θ1 
               
               
                   
                   
               
             
          
         
       
     
         [0192]     The artificial satellite using either of the above orbital elements becomes visible in an elevation angle above 70 degrees coincidentally from all the cities from Nemuro to Naha over 8 hours at the maximum to 6 hours at the minimum. The longest duration time of 8 hours can be obtained when the orbit inclination angle is approximately 42.5 degrees and the eccentricity is approximately 0.21. Therefore, by the group of artificial satellites using the above-mentioned orbital element, it can be realized that at least one or more satellites become visible in a high elevation angle coincidentally from all the cities from Nemuro to Naha over 24 hours per day.  
         [0193]     In this orbit arrangement example, there are three orbital planes as shown in  FIG. 6 , and each of the artificial satellite  60 , the artificial satellite  61  and the artificial satellite  62  is arranged one on each of the orbits. The artificial satellite  60 , the artificial satellite  61  and the artificial satellite  62  travel with one turn in approximately 24 hours on the orbit  63 , the orbit  64  and the orbit  65 , respectively. Each of the artificial satellite  60 , the artificial satellite  61  and the artificial satellite  62  travels on each of the orbits with an orbital period of 24 hours, and each of the orbits is formed so that the argument of perigee is within the range of larger than 180 degrees and smaller than 360 degrees; and the eccentricity is not larger than 0.24 and the orbital inclination angle is larger than 37 degrees and smaller than 44 degrees, or the eccentricity is larger than 0.24 and smaller than 0.35 and the orbital inclination angle is larger than 40 degrees and smaller than 44 degrees. The right ascensions of north-bound node of the three artificial satellites are separated by 120 degrees as shown in  FIG. 6 , and set so that the apogee of each of the orbits appears an appropriate position in the sky above the territory of Japan. As the positional relationship of each of the artificial satellites in each of the orbits, when the artificial satellite  60  is at the perigee on the corresponding orbit  63 , the artificial satellite  61  is at a position having a true anomaly lagging behind by one-third of the orbital period on the corresponding orbit  64 ; and the artificial satellite  62  is at a position having a true anomaly leading ahead by one-third of the orbital period on the corresponding orbit  65 . This orbit arrangement is obtained by the algorithm shown the outline in  FIG. 2  and the algorithm of the setting method of the argument of perigee, and is realized by the control method shown in  FIG. 3 ,  FIG. 4  and  FIG. 5 .  
         [0194]     By this orbit arrangement, any one of the artificial satellite  60 , the artificial satellite  61  and the artificial satellite  62  is always visible in an elevation angle above 70 degrees from the territory of Japan from Hokkaido to Okinawa. Since each of the artificial satellite  60 , the artificial satellite  61  and the artificial satellite  62  has the period of nearly 24 hours, the time when the artificial satellite becomes visible in the elevation angle above 70 degrees and the time when the artificial satellite becomes invisible are periodical and regular. In this case, in the territory of Japan, the artificial satellite  60 , the artificial satellite  61  and the artificial satellite  62  alternatively appear in the elevation angle above 70 degrees with one cycle per day, and each of the artificial satellites stays and is visible in a direction in the elevation angle above 70 degrees for 8 hours at the maximum and 6 hours at the minimum. This cycle is repeated every day with 24 hour cycle.  
         [0195]     Therefore, by using the artificial satellites typically shown by the artificial satellite  90  in  FIG. 8  to  FIG. 34  showing examples of systems employing the orbit arrangement for satellite communication or satellite broadcast, it is possible to construct a satellite communication system or a satellite broadcast system which seldom cause communication interruption by shielding objects or interfering objects.  
         [0196]     Table 6 and Table 7 show examples of the orbital elements and the orbit arrangement for a satellite communication and broadcast network by four artificial satellites traveling on an orbit of 24-hour orbital period. The eccentricity and the orbit inclination angle may be a combination within the range shown in  FIG. 35  and  FIG. 36 . The argument of perigee may be smaller than 180 degrees.  
                                                                         TABLE 6                       Item   Value                                Satellite No.   1   2   3   4            Orbital Period (hour)   24       Semimajor Axis (km)   approx. 42,164       Eccentricity   not larger than 0.24       Orbit Inclination   larger than 37 degrees and smaller       Angle (degrees)   than 44 degrees       Argument of Perigee   larger than 180 degrees and smaller       (degrees)   than 360 degrees            Right Ascension   Ω1   Ω1 + 90   Ω1 + 180   Ω1 + 270       of North-Bound       node (degrees)       True Anomaly   θ1   an angle ¼ of   θ1 + 180   an angle ¼ of       (degrees)       orbital period       orbital period               behind θ1       ahead of θ1                  
 
         [0197]    
       
         
               
               
             
               
               
               
               
               
             
               
               
             
               
               
               
               
               
             
           
               
                 TABLE 7 
               
               
                   
               
               
                   
               
               
                 Item 
                 Value 
               
               
                   
               
             
             
               
                   
               
             
          
           
               
                 Satellite No. 
                 1 
                 2 
                 3 
                 4 
               
             
          
           
               
                 Orbital Period (hour) 
                 24 
               
               
                 Semimajor Axis (km) 
                 approx. 42,164 
               
               
                 Eccentricity 
                 larger than 0.24 and smaller then 0.35 
               
               
                 Orbit Inclination 
                 larger than 40 degrees and smaller 
               
               
                 Angle (degrees) 
                 than 44 degrees 
               
               
                 Argument of Perigee 
                 larger than 180 degrees and smaller 
               
               
                 (degrees) 
                 than 360 degrees 
               
             
          
           
               
                 Right Ascension 
                 Ω1 
                 Ω1 + 90 
                 Ω1 + 180 
                 Ω1 + 270 
               
               
                 of North-Bound 
               
               
                 node (degrees) 
               
               
                 True Anomaly 
                 θ1 
                 an angle ¼ of 
                 θ1 + 180 
                 an angle ¼ of 
               
               
                 (degrees) 
                   
                 orbital period 
                   
                 orbital period 
               
               
                   
                   
                 behind θ1 
                   
                 ahead of θ1 
               
               
                   
               
             
          
         
       
     
         [0198]     The artificial satellite using either of the above orbital elements becomes visible in an elevation angle above 70 degrees coincidentally from all the cities from Nemuro to Naha over 8 hours at the maximum to 6 hours at the minimum. The longest duration time of 8 hours can be obtained when the orbit inclination angle is approximately 42.5 degrees and the eccentricity is approximately 0.21. Therefore, by the group of artificial satellites using the above-mentioned orbital element, it can be realized that at least one or more satellites become visible in a high elevation angle coincidentally from all the cities from Nemuro to Naha over 24 hours per day.  
         [0199]     In this orbit arrangement example, there are four orbital planes as shown in  FIG. 7 , and each of the artificial satellite  70   a,  the artificial satellite  70   b , the artificial satellite  70   c  and the artificial satellite  70   d  is arranged one on each of the orbit planes. The artificial satellite  70   a  travels on the orbit  71   a  with one turn in approximately 24 hours, the artificial satellite  70   b  on the orbit  71   b,  the artificial satellite  70   c  on the orbit  71   c  and the artificial satellite  70   d  on the orbit  71   d.  Each of the artificial satellite  70   a,  the artificial satellite  70   b,  the artificial satellite  70   c  and the artificial satellite  70   d  travels on each of the orbits with an orbital period of 24 hours, and each of the orbits is formed so that the argument of perigee is within the range of larger than 180 degrees and smaller than 360 degrees; and the eccentricity is not larger than 0.24 and the orbital inclination angle is larger than 37 degrees and smaller than 44 degrees, or the eccentricity is larger than 0.24 and smaller than 0.35 and the orbital inclination angle is larger than 40 degrees and smaller than 44 degrees. The right ascensions of north-bound node of the four artificial satellites are separated by 90 degrees as shown in  FIG. 7 , and set so that the apogee of each of the orbits appears an appropriate position in the sky above the territory of Japan. As the positional relationship of each of the artificial satellites in each of the orbits, when the artificial satellite  70   a  is at the perigee on the corresponding orbit  71   a,  the artificial satellite  70   b  is at a position having a true anomaly lagging behind by one-fourth of the orbital period on the corresponding orbit  71   b;  the artificial satellite  70   c  is at an apogee on the corresponding orbit  71   c;  and the artificial satellite  70   d  is at a position having a true anomaly leading ahead by one-fourth of the orbital period on the corresponding orbit  71   d.    
         [0200]     By this orbit arrangement, any one of the artificial satellite  70   a,  the artificial satellite  70   b,  the artificial satellite  70   c  and the artificial satellite  70   d  is always visible in an elevation angle above 70 degrees from the territory of Japan from Hokkaido to Okinawa. Since each of the artificial satellite  70   a , the artificial satellite  70   b,  the artificial satellite  70   c  and the artificial satellite  70   d  has the period of nearly 24 hours, the time when the artificial satellite becomes visible in the elevation angle above 70 degrees and the time when the artificial satellite becomes invisible are periodical and regular. This orbit arrangement is obtained by the algorithm shown the outline in  FIG. 2  and the algorithm of the setting method of the argument of perigee, and is realized by the control method shown in  FIG. 3 ,  FIG. 4  and  FIG. 5 .  
         [0201]     In this case, in the four islands of Hokkaido, Honshuu, Shikoku and Kyuushuu and Okinawa, the artificial satellite  70   a , the artificial satellite  70   b , the artificial satellite  70   c  and the artificial satellite  70   d  alternatively appear in the elevation angle above 70 degrees once a day, and each of the artificial satellites stays and is visible in a zenith direction above 70 degrees for 8 hours at the maximum and 6 hours at the minimum. In addition, there is some time when a plurality of the artificial satellites are visible in a zenith direction above 70 degrees at a time. This cycle is repeated every day with 24 hour cycle.  
         [0202]     Therefore, by using the artificial satellites typically shown by the artificial satellite  90  in  FIG. 8  to  FIG. 34  showing examples of systems employing the orbit arrangement for satellite communication or satellite broadcast, it is possible to construct a satellite communication system or a satellite broadcast system which seldom cause communication interruption by shielding objects or interfering objects.  
         [0000]     (3) Means for realizing and controlling the set orbital elements and the set orbit arrangement  
         [0203]     The orbits of the artificial satellites having the orbital elements set such a manner are controlled and realized as follows.  
         [0204]     As shown in  FIG. 3 , at launching the artificial satellite  20 , information of the previously set six orbital elements  17  at the reference time is input to a launch vehicle tracking and control facility  21  from which information of target injection orbit elements  22  is transmitted to the launch vehicle. The launch vehicle  23  is injected into a target orbit automatically based on the information or by control from the launch vehicle tracking and control facility  21 .  
         [0205]     After the artificial satellite is injected into the orbit, the information of the six orbital elements at reference time is periodically input to an artificial satellite tracking and control facility  18  to transmit information of a control command  19  to the artificial satellite  20 , and the artificial satellite  20  is controlled to the six orbital elements at target orbital time by a control system mounted on the artificial satellite  20 .  
         [0206]     This method of orbit control is based on a commonly used method, and the detail is as follows.  
         [0207]     The six orbital elements  17  (the semi-major axis  11 , the argument of perigee  12 , the eccentricity  13 , the orbit inclination angle  14 , the right ascension of north-bound node  15  and the true anomaly  16 ) at reference time obtained by the above-mentioned algorithm are input to the launch vehicle tracking and control facility  21  as the target injection orbit elements, as shown in  FIG. 3 . This information is transmitted from the launch vehicle tracking and control facility  21  to the launch vehicle  23  to inject the artificial satellite  20  into the target orbital elements. When the launch vehicle  23  mounting the artificial satellite  20  is about to deviate from the target orbit in the stage of launching, the launch vehicle  23  may automatically correct the orbit, or the launch vehicle tracking and control facility  21  may transmit an orbit correction command to the launch vehicle  23  to guide the launch vehicle.  
         [0208]     Even after reaching the target injection orbital elements  22  by the launch vehicle  23 , the orbital elements are affected with perturbation by the effects of the gravitation field of the earth, the gravitational forces of the sun and the moon and the solar window, and the orbital elements are always changed in short period and long period as time is elapsing. In this case, the artificial satellite  20  is required to be controlled.  
         [0209]     As shown in  FIG. 4 , in general, the six orbital elements  31  of the orbit on which the artificial satellite  20  is now traveling are determined by that a send and receive system  24  of the artificial satellite tracking and control facility  18  receives telemetry and ranging signals  27  sent by the artificial satellite  20 , and extracts the ranging signal  28  to send it to a range measurement system  25 , and then a computer system  26  executes calculation from the measured range and the rate of change of the range  29  as final inputs using an orbit determining program in the computer system  26 . The computer system  26  calculates a necessary attitude control variable and an orbital control variable  33  by comparing the obtained six orbital elements  31  with the target six orbital elements at reference time using an orbit control program  32  in the computer system  26 . Thereby, it is determined when and how long and which thruster of the gas jet unit mounted on the artificial satellite should be jetted. The results are converted into a control command  35  using a command generating program  34  in the computer system  26  to be sent to the artificial satellite  20  through the send and receive system  24  of the artificial satellite tracking and control facility  18 .  
         [0210]     As shown in  FIG. 5 , the control command transmitted to the artificial satellite  20  is received by a communication system  37  mounted on the artificial satellite  20 , and then transmitted to a data processing system  38 , and the transmitted command is deciphered there. Information of an attitude control variable and an orbit control variable  41  is appropriately processed in an attitude and orbit control system  39  mounted on the artificial satellite from the deciphered command, and finally the artificial satellite  20  is injected and controlled to the orbit shown by the above-mentioned six orbital elements at reference time by being changed in the attitude by operating an attitude control actuator drive  42  depending on necessary, and further by jetting the gas jet unit  40  of the artificial satellite mounting propelling system according to the command. In a case where the artificial satellite  20  mounts a GPS satellite receiver, the artificial satellite  20  may be constructed so that the artificial satellite  20  itself stores the six orbital elements at reference time  17  preferable at that time point in advance, and autonomously controls the orbit using the stored six orbital elements at reference time  17 .  
         [0211]     As described above, the orbit elements  17  determined by the aforementioned algorithm are controlled and realized.  
         [0212]     Further, in the case where the plurality of artificial satellites are arranged on the orbits, it is necessary to appropriately control the individual orbits of the artificial satellites so that each of the orbits of the artificial satellites maintains a preferable harmonious relationship of the orbit arrangement.  
         [0213]     Description will be made below on systems to which a group of artificial satellites traveling on the orbits obtained by the aforementioned algorithm in accordance with the present invention is applied.  
         [0000]     (4) Systems to which a group of artificial satellites traveling on the orbits in accordance with the present invention is applied  
         [0000]     (4-1) System Example 1  
         [0214]     An example of system  1  is a satellite broadcast system.  
         [0215]      FIG. 8  shows the embodiment of the satellite broadcast system in accordance with the present invention.  
         [0216]     As shown in  FIG. 8 , the satellite broadcast system is composed of a group of artificial satellites  90  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; a base station  91  for sending satellite broadcast through the group of artificial satellites  90 ; and a satellite broadcast terminal  92  for receiving the satellite broadcast through the group of satellites  90 .  
         [0217]     As shown in  FIG. 9 , the base station  91  is composed of an antenna  91   a , an antenna homing system  91   b , a large electric power amplifier  91   c,  a frequency converter  91   d,  a modulator  91   e , an error correction encoder  91   f,  an encipherer  91   g,  a multiplexer  91   h,  an encoder  91   i.    
         [0218]     Image information  91   k  and voice information  91   l  are highly efficiently encoded by the encoder  91   i,  and other image information and other voice information and data  91   j  are maltiplexed by the multiplexer  91   h.  Further, they are enciphered by the encipher  91   g,  added with error correction code by the error correction encoder  91   f,  further modulated by the modulator  91   e  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  91   d,  amplified by the large electric power amplifier  91   c,  and then sent from the antenna  91   a  homing the group of artificial satellites  90  using the antenna homing system  91   b.    
         [0219]     On the other hand, as shown in  FIG. 10 , the satellite broadcast terminal  92  is composed of an antenna  92   a,  a low noise amplifier  92   b,  a frequency converter  92   c,  a demodulator  92   d,  an error corrector  92   e,  a decipherer  92   f,  a demultiplexer  92   g,  a decoder  92   h,  a frame memory  91   i.    
         [0220]     An electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  92   a,  amplified by the low noise amplifier  92   b,  converted to an intermediate frequency by the frequency converter  92   c,  and demodulated to a digital signal by the demodulator  92   d.  Further, the digital signal is corrected by the error corrector  92   e  if there is an error, the ciphered information is deciphered by the decipherer  92   f,  and a requested broadcast is selected by the demultiplexer  92   g.  Further, the signal is returned to image information  92   k  and voice information  92   l  by the decoder  92   h.  The decoder  92   h  has a frame memory  92   i  and can complement lack of data thereby.  
         [0221]     According to the present invention, even when the satellite broadcast terminal  92  uses a directional antenna, there is an advantage in that it is sufficient to simply direct it to the zenith direction and it is completely unnecessary for a user to search a direction (north, south, east or west direction) of the group of satellites.  
         [0222]     Further, in the case of broadcast from a stationary satellite, the satellite broadcast terminal antenna for a mobile object needs to be symmetric in direction and sensitive in 45-degree direction. However, in the case of the present invention, since it is sufficient that the antenna has a directionality only in the zenith direction, there is an advantage in that the antenna is easily manufactured and gain of the antenna can be increased twice or more. By making use of this advantage, the output power from the satellite may be reduced ½, or may be transmit twice as much as information (broadcast) if the output power is kept as it is.  
         [0223]     Furthermore, since the group of artificial satellites  90  are always located in a high elevation angle, the electromagnetic wave from the group of artificial satellites  90  can be directly received regardless of an environmental condition of the mobile object such as a place opened only in the zenith direction in a street lined with large buildings. Therefore, it is possible to provide a high-quality receiving environment without reflected waves from the buildings, and accordingly there is an advantage in that it is possible to transmit more information (broadcast) than in a case of broadcast from a stationary satellite even if the same frequency band is used. The effects described above can be said to the systems to be described below.  
         [0224]     Referring to  FIG. 9  and  FIG. 10 , description will be made below on an example where an accounting system is added, and broadcast is provided to limited customers with fee.  
         [0225]     As shown in  FIG. 9 , an accounting system  91   m  of the base station  91  is composed of a limiting receiving unit  91   n,  a customer management unit  91   o,  an accounting management system  91   p  and a limiting receiving module issuing unit  91   q.    
         [0226]     The satellite broadcast terminal  92  comprises a limiting receiving module  92   m,  as shown in  FIG. 10 .  
         [0227]     Customer information (receiving status of viewing fee, viewing request information, address, name and so on) is managed by the customer management unit  91   o,  and the accounting management system  91   p  controls cipher for each customer by controlling the encipherer  91   g  through the limiting receiving unit  91   n  according to the customer information. Further, the accounting management system  91   p  issues a limiting receiving module (an IC card, as an example) using the limiting receiving module issuing unit  91   q  according to the customer information in the customer management system  91   p.  Although it is not described here, customer information on receipt of fee from a financial institution is input to the accounting management system  91   p  to update the customer information using the customer management unit  91   o.    
         [0228]     A user can view a requested broadcast by inserting the above-mentioned limiting receiving module  92   m  obtained as a consideration for the payment into the decipherer  92   f  of the satellite broadcast terminal  92 .  
         [0229]     Thereby, by using a means for enciphering a broadcast program in the base station  91  and by using a means for deciphering it in the satellite broadcast terminal  92 , it is possible to broadcast only to the limited satellite broadcast terminal and to charge for the service.  
         [0230]      FIG. 11  shows another embodiment of satellite broadcast system.  
         [0231]     As shown in  FIG. 11 , the satellite broadcast system is composed of a group of artificial satellites  90  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; a base station  91  for sending satellite broadcast through the group of artificial satellites  90 ; a ground broadcast station  93 ; and a satellite broadcast terminal  94  having a means for receiving the satellite broadcast through the group of satellites  90  and a means for receiving ground broadcast.  
         [0232]     As shown in  FIG. 12 , the ground broadcast station  93  is composed of an antenna  93   a,  a large electric power amplifier  93   b,  a frequency converter  93   c,  a modulator  93   d,  an error correction encoder  93   e,  an encipherer  93   f,  a maltiplexer  93   g,  an encoder  93   h.    
         [0233]     Image information  93   j  and voice information  93   k  are highly efficiently encoded by the encoder  93   h,  and other image information and other voice information and data  93   i  are maltiplexed by the maltiplexer  93   g.  Further, they are enciphered by the encipher  93   f,  added with error correction code by the error correction encoder  93   e,  further modulated by the modulator  93   d  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  93   c,  amplified by the large electric power amplifier  93   b,  and then sent from the antenna  93   a.    
         [0234]     On the other hand, as shown in  FIG. 13 , the satellite broadcast terminal  94  is composed of an antenna  94   a   1  for receiving ground broadcast and a low noise amplifier  94   a   2 , an antenna  94   b   1  for receiving a electromagnetic wave from the group of artificial satellites  90  and a low noise amplifier  94   b   2 , a frequency converter  94   c  for broadcast from the ground station and the satellite, a demodulator  94   d,  an error corrector  94   e,  a decipherer  94   f,  a demultiplexer  94   g,  a decoder  94   h,  a frame memory  94   i.    
         [0235]     An electromagnetic wave sent from the ground broadcast station  93  is received by the antenna  94   a   1 , amplified by the low noise amplifier  94   b   1 , and on the other hand, an electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  94   a   2 , amplified by the low noise amplifier  94   b   2 , and each of the signals are converted to an intermediate frequency by the frequency converter  94   c,  and demodulated to a digital signal by the demodulator  94   d.  Further, the digital signal is corrected by the error corrector  94   e  if there is an error, the ciphered information is deciphered by the decipherer  94   f,  and a requested broadcast is selected by the demultiplexer  94   g.  Further, the signal is returned to image information  94   k  and voice information  94   l  by the decoder  94   h.  The decoder  94   h  has a frame memory  94   i  and can complement lack of data thereby.  
         [0236]     Since the satellite broadcast terminal  94  can receive the ground broadcast as well as the satellite broadcast, the satellite broadcast terminal  94  has an advantage in that a user can select a desired broadcast program at will. In addition, there is an advantage in that if the user has the satellite broadcast terminal  94 , he needs not to possess both of a satellite broadcast terminal and a ground broadcast terminal. Further, an antenna may be used both as the antennas  94   a   1  and  94   a   2 .  
         [0237]     Referring to  FIG. 12  and  FIG. 13 , description will be made below on an example where an accounting system is added, and broadcast is provided to limited customers with fee.  
         [0238]     As shown in  FIG. 12 , an accounting system  93   m  of the base station  93  is composed of a limiting receiving unit  93   n,  a customer management unit  93   o,  an accounting management system  93   p  and a limiting receiving module issuing unit  93   q.    
         [0239]     The satellite broadcast terminal  94  comprises a limiting receiving module  92   m,  as shown in  FIG. 13 .  
         [0240]     Customer information (receiving status of viewing fee, viewing request information, address, name and so on) is managed by the customer management unit  93   o,  and the accounting management system  93   p  controls cipher for each customer by controlling the encipherer  93   f  through the limiting receiving unit  93   n  according to the customer information. Further, the accounting management system  93   p  issues a limiting receiving module (an IC card, as an example) using the limiting receiving module issuing unit  93   q  according to the customer information in the customer management system  93   o.  Although it is not described here, customer information on receipt of fee from a financial institution is input to the accounting management system  93   p  to update the customer information using the customer management unit  93   o.    
         [0241]     A user can view a requested broadcast by inserting the above-mentioned limiting receiving module  94   m  obtained as a consideration for the payment into the decipherer  94   f  of the satellite broadcast terminal  94 .  
         [0242]     Thereby, by using a means for enciphering a broadcast program in the base station  93  and by using a means for deciphering it in the satellite broadcast terminal  94 , it is possible to broadcast only to the limited satellite broadcast terminal and to charge for the service.  
         [0243]      FIG. 14  shows another embodiment of satellite broadcast system.  
         [0244]     As shown in  FIG. 14 , the satellite broadcast system is composed of a group of artificial satellites  90  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; a ground communication means  95  such as a public network, a cellular phone and the like; a base station  96  having a means for sending satellite broadcast through the group of artificial satellites  90  and the above-described ground communication means; and a satellite broadcast terminal  97  having a means for receiving the satellite broadcast through the group of satellites  90  and the above-described ground communication means.  
         [0245]     As shown in  FIG. 15 , the base station  96  is composed of an antenna  96   a , an antenna homing system  96   b , a large electric power amplifier  96   c , a frequency converter  96   d , a modulator  96   e,  an error correction encoder  96   f,  an encipherer  96   g,  a multiplexer  96   h,  an encoder  96   i,  a data selector  96   n,  a data memory  96   o  and a ground communication network  95 .  
         [0246]     Request information  96   m  through the ground communication network  95  is input to the data selector  96   n , and image information  96   k  and voice information  96   l  and data  96   j  are cited from the data memory  96   o , if necessary. Further, the image information  96   k  and the voice information  96   i  are highly efficiently encoded, other image information and other voice information and data  96   j  are maltiplexed by the maltiplexer  96   h.  Further, they are enciphered by the encipher  96   g , added with error correction code by the error correction encoder  96   f,  further modulated by the modulator  96   e  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  96   d , amplified by the large electric power amplifier  96   c , and then sent from the antenna  96   a  homing the group of artificial satellites  90  using the antenna homing system  96   b.    
         [0247]     On the other hand, as shown in  FIG. 16 , the satellite broadcast terminal  92  is composed of an antenna  97   a , a low noise amplifier  97   b , a frequency converter  97   c , a demodulator  97   d , an error corrector  97   e , a decipherer  97   f , a demultiplexer  97   g,  a decoder  97   h , a frame memory  97   i , a request sender  97   n  and the ground communication network  95 .  
         [0248]     A request  97   m  is sent to the base station  96  of  FIG. 15  by the request sender  97   n  such as a PHS, a cellular phone or the like through the ground communication network  95 . The base station  96  sends the requested information to the group of artificial satellites  90 , and an electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  97   a , amplified by the low noise amplifier  97   b , converted to an intermediate frequency by the frequency converter  97   c , and demodulated to a digital signal by the demodulator  97   d . Further, the digital signal is corrected by the error corrector  97   e  if there is an error, the ciphered information is deciphered by the decipherer  97   f , and a requested broadcast is selected by the demultiplexer  97   g.  Further, the signal is returned to image information  97   k  and voice information  97   l  by the decoder  97   h . The decoder  97   h  has a frame memory  97   i  and can complement lack of data thereby.  
         [0249]     Thereby, the satellite broadcast terminal  97  has an advantage in that an user can send a request to the base station  91  to broadcast desired information.  
         [0250]     The satellite broadcast terminals  92 ,  94 ,  97  in the present invention may be mounted on a mobile objects such as a car, a train, a ship, an aircraft or the like, and further may be carried by a walker, a climber and so on. Furthermore, the satellite broadcast terminals may be used at a place not movable such as a home.  
         [0251]     The contents of program to be broadcast are not limited in the present invention.  
         [0252]     The programs to be broadcast are not only TV broadcast and voice broadcast, but also digital information.  
         [0253]     There are various kinds of programs to be broadcast such as weather information, fishing information (water level, water temperature and so on), ITS information (traffic amount information, traffic speed information, traffic congestion place information, traffic congesting time information, driving environment information, stricken area information,. traffic restriction information, optimum route information, information on required time in congesting time, parking lot status information, parking lot reservation information, destination information (weather, travel, sightseeing, meals, recreation information), various kinds of reservation information (public transportation, hotels, amusement facilities)), map information (map information, updated information and so on), car navigation information (car navigation information, updated information and so on), software program information (programs for car navigation, programs for game, OS and so on), voice data (including compressed data by MP3 or the like), and amusement information.  
         [0254]     Further, as the programs to be broadcast, there are maltimedia information such as the Internet and the like, and differential GPS information.  
         [0255]     Further, as the programs to be broadcast, there is information limited to an area or on an area where a mobile object is moving such as time service information of a department store or a supermarket, exhibition information of an art gallery and a museum, information on presentation contents of a movie house or a show house, information on a criminal or a lingering person.  
         [0256]     Referring to  FIG. 15  and  FIG. 16 , description will be made below on an example where an accounting system is added, and broadcast is provided to limited customers with fee.  
         [0257]     As shown in  FIG. 15 , an accounting system  96   p  of the base station  96  is composed of a limiting receiving unit  96   q,  a customer management unit  96   r,  an accounting management system  96   s  and a limiting receiving module issuing unit  96   t.    
         [0258]     The satellite broadcast terminal  97  comprises a limiting receiving module  97   o,  as shown in  FIG. 16 .  
         [0259]     Customer information (receiving status of viewing fee, viewing request information, address, name and so on) is managed by the customer management unit  96   r,  and the accounting management system  96   s  controls cipher for each customer by controlling the encipherer  96   g  through the limiting receiving unit  96   q  according to the customer information. Further, the accounting management system  96   s  issues a limiting receiving module (an IC card, as an example) using the limiting receiving module issuing unit  96   t  according to the customer information in the customer management system  96   r . Although it is not described here, customer information on receipt of fee from a financial institution is input to the accounting management system  96   s  to update the customer information using the customer management unit  96   r.    
         [0260]     A user can view a requested broadcast by inserting the above-mentioned limiting receiving module  97   o  obtained as a consideration for the payment into the decipherer  97   f  of the satellite broadcast terminal  97 .  
         [0261]     Thereby, by using a means for enciphering a broadcast program in the base station  96  and by using a means for deciphering it in the satellite broadcast terminal  97 , it is possible to broadcast only to the limited satellite broadcast terminal and to charge for the service.  
         [0000]     (4-2) System Example 2  
         [0262]     An example of system  2  is a satellite broadcast system.  
         [0263]      FIG. 17  shows the embodiment of the satellite communication system in accordance with the present invention.  
         [0264]     As shown in  FIG. 17 , the satellite communication system is composed of a group of artificial satellites  90  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; a base station  98  and a satellite communication send and receive unit  99  for performing satellite communication through the group of artificial satellites  90 .  
         [0265]     As shown in  FIG. 18 , the base station  98  is composed of an antenna  98   a , an antenna homing system  98   b , a large electric power amplifier  98   c , a frequency converter  98   d , a modulator  98   e,  an encoder  98   f , a low nose amplifier  98   h , a frequency converter  98   i,  a demodulator  98   j,  and a decoder  98   k.    
         [0266]     Sent data  98   g  is encoded and enciphered and added with error correction code by the encoder  98   f , and further modulated by the modulator  98   e  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  98   d , amplified by the large electric power amplifier  98   c , and then sent from the antenna  98   a  homing the group of artificial satellites  90  using the antenna homing system  98   b . On the other hand, an electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  98   a , amplified by the low noise amplifier  98   h , converted to an intermediate frequency by the frequency converter  98   h , and decoded to a digital signal by the demodulator  98   j.  Further, by being error-corrected and deciphered and decoded by the decoder  98   k,  received data  98   l  can be obtained.  
         [0267]     On the other hand, as shown in  FIG. 19 , the satellite communication send and receive unit  99  is composed of an antenna  99   a,  a large electric power amplifier  99   b,  a frequency converter  99   c , a modulator  99   d , an encoder  99   e , a low noise amplifier  99   g,  a frequency converter  99   h,  a demodulator  99   i,  and a decoder  99   j.    
         [0268]     Sent data  99   f  is encoded and enciphered and added with error correction code by the encoder  99   e , and further modulated by the modulator  99   d  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  99   c , amplified by the large electric power amplifier  99   b,  and then sent from the antenna  99   a.  On the other hand, an electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  99   a,  amplified by the low noise amplifier  99   g,  converted to an intermediate frequency by the frequency converter  99   h,  and decoded to a digital signal by the demodulator  99   i.  Further, by being error-corrected and deciphered and decoded by the decoder  99   j,  received data  99   k  can be obtained.  
         [0269]     According to the present embodiment, since at least one artificial satellite of the group of artificial satellites is visible in a position near the zenith, the communication line can be easily maintained for a long time by using the satellite communication system even in an area where there are shielding objects shielding the field of view such as artificial buildings, trees, mountains and so on.  
         [0270]     For example, by installing the base station  98  at a gateway communication station of public lines and by making the satellite communication send and receive unit  99  carried with a person, the satellite communication system can be used as a cellular phone.  
         [0271]     For example, by installing the base station  98  at a hospital and by installing the satellite communication send and receive unit  99  in an ambulance, a patient taken to the hospital in the ambulance can be appropriately treated since appropriate first aid can be transmitted from a medical specialist in the hospital by sending image data with regard to the patient from the ambulance to the hospital. Thereby, it becomes possible to save life in such a case where if a patient have been treated with appropriate first aid, his live might have been saved.  
         [0272]     For example, by installing the base station  98  at a broadcast station and by installing the satellite communication send and receive unit  99  in a broadcast car, the satellite communication system can be used for a TV program of mobile sport relay broadcasting such as a marathon race relay broadcasting or the like, and accordingly a high quality image can be transmitted in real time and a dynamic program can be provided.  
         [0273]     For example, by installing the base station  98  at a fire station and by installing the satellite communication send and receive unit  99  in a fire engine, since a high quality image of a situation of a site under fire-fighting can be transmitted to the fire station from a place between tall buildings or in a narrow path in teal time, appropriate judgment can be made.  
         [0274]     For example, by installing the base station  98  at a police station and by installing the satellite communication send and receive unit  99  in a squad car, since a high quality image of a criminal or a situation of a site under fire-fighting can be transmitted to the fire station from a place between tall buildings or in a narrow path in teal time, effective guard can be performed.  
         [0275]     For example, by installing the base station  98  at a hospital having a medical specialist and by installing the satellite communication send and receive unit  99  in a mobile object and moving the mobile object to a clinic requiring an advice or diagnosis of the medical specialist in a distant place to communicate information on a patient, regional difference in medical service can be solved by transmitting information in real time even in a mountainous region.  
         [0276]     For example, by installing the base station  98  at a stock center and by installing the satellite communication send and receive unit  99  in a vending machine, customer service can be improved since the inventory can be periodically or arbitrarily checked.  
         [0277]     Further the satellite communication system can be applied to train control information communication, train maintenance communication, train signal control communication, vehicle operating status communication, ship information operating status communication, data acquisition system (float, buoy or the like), personal computer communication (electronic mail, Internet, on-line shopping and so on), parking lot vacant information supply/reservation system and so on.  
         [0278]      FIG. 20  shows another embodiment of the satellite communication system in accordance with the present invention.  
         [0279]     As shown in  FIG. 20 , the satellite communication system is composed of a group of artificial satellites  90  in accordance with the present invention; a base station  98  for performing satellite communication through the group of artificial satellites  90 ; a group of artificial satellites  100  composing a global position measuring system; and a satellite communication send and receive unit  101  having a function capable of measuring its own position using a positioning signal from the group of artificial satellites composing the global positioning system and a function capable of performing communication through the group of artificial satellites  90 .  
         [0280]     As shown in  FIG. 21 , the satellite communication send and receive unit  101  is composed of an antenna  101   a , a large electric power amplifier  101   b , a frequency converter  101   c,  a modulator  101   d,  an encoder  101   e,  a maltiplexer  101   f,  a low noise amplifier  101   h,  a frequency converter  101   i,  a demodulator  101   k,  and a GPS receiver  101   m.    
         [0281]     Sent data  101   g  is maltiplexed with positioning information  101   n  output from the GPS receiver  101   m  by the multiplexer  101   f,  encoded and enciphered and added with error correction code by the encoder  101   e,  and further modulated by the modulator  101   d  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  101   c,  amplified by the large electric power amplifier  101   b,  and then sent from the antenna  101   a . On the other hand, an electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  101   a , amplified by the low noise amplifier  101   h,  converted to an intermediate frequency by the frequency converter  101   i,  and decoded to a digital signal by the demodulator  101   j.  Further, by being error-corrected and deciphered and decoded by the decoder  101   k,  received data  101   l  can be obtained.  
         [0282]     In this system, position information of the send and receive unit can be sent to the base station.  
         [0283]     Further, with regard to timing to send positioning information, there are a demand method (sending is made at the time when a user requests) and a polling method (by receiving sending request of the center station, the send and receive unit automatically sends the information).  
         [0284]     For example, by installing the base station  98  at a mountain search and rescue center such as a police station or a fire station and by making the satellite communication send and receive unit  101  carried with a mountaineer, rescue activity can be speedily and accurately performed if the mountaineer meets with a disaster because the center can be informed of a position of the mountaineer. Further, because of capability of bi-directional communication, the present invention has an advantage in that words of encouragement to the mountaineer met with a disaster and confirmation whether or not information is erroneous can be performed. Furthermore, the satellite communication send and receive unit may have only a send function in order to make light in weight and small in consuming electric power.  
         [0285]     For example, by installing the base station  98  at a search and rescue center for the perils of the sea such as a police station or a fire station and by installing the satellite communication send and receive unit  101  in a ship, it is possible to perform an error check, to speedily and accurately perform rescue activity because the center and the Maritime Safety Agency can be informed of a position of the wrecked ship if the ship meets with a shipwreck. Further, because of capability of bi-directional communication, the present invention has an advantage in that words of encouragement to the sailors met with the shipwreck and confirmation whether or not the information is erroneous can be performed. Furthermore, the satellite communication send and receive unit may have only a send function in order to make light in weight and small in consuming electric power.  
         [0286]     For example, by installing the base station  98  at a police station and by making the satellite communication send and receive unit  101  carried with a person, rescue activity can be speedily and accurately performed if the person lingers around or loses his way because a position of the person can be detected by the police station. Further, because of capability of bi-directional communication, the present invention has an advantage in that words of encouragement to the mountaineer met with a disaster and confirmation whether or not information is erroneous can be performed. Furthermore, the satellite communication send and receive unit may have only a send function in order to make light in weight and small in consuming electric power.  
         [0287]     For example, by installing the base station  98  at a police station and by installing the satellite communication send and receive unit  101  in a vehicle, search activity can be speedily and accurately performed if the vehicle is stolen because a position of the stolen vehicle can be detected by the police station.  
         [0288]     For example, by installing the base station  98  at a physical distribution center and by installing the satellite communication send and receive unit  101  in a mobile object (a track, a train, a taxicab, a bus, a container), physical distribution management and mobile object arranging management can be speedily and accurately performed because a position of the mobile object can be instantaneously detected by the center.  
         [0289]     Further, the satellite communication system can be applied to an optimum path guiding system, a request type navigation system, an animal behavior monitor (wild animal (behavior monitor), cattle (stray prevention, exercise amount detection), animals fed in a zoo (danger prevention) and so on).  
         [0290]      FIG. 22  shows another embodiment of the satellite communication system in accordance with the present invention.  
         [0291]     As shown in  FIG. 22 , the satellite communication system is composed of a group of artificial satellites  90  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; a base station  98  for performing satellite communication through the group of artificial satellites  90 ; and a satellite communication send and receive unit  102  having a function to measure a consumed amount of at least one of electricity, city gas and city water and a function capable of communication through the group of satellites  90 .  
         [0292]     In the case of measuring electricity consumed amount, the satellite communication send and receive unit  102  is composed of an antenna  102   a , a large electric power amplifier  102   b , a frequency converter  102   c , a modulator  102   d , an encoder  102   e,  a multiplexer  102   f,  a low noise amplifier  102   h,  a frequency converter  102   i,  a demodulator  102   j,  a decoder  102   k,  and an electric power meter  102   m,  as shown in  FIG. 23 .  
         [0293]     Sent data  102   g  is multiplexed with consumed amount information  102   n  output from the electric power meter  102   m  by the multiplexer  102   f , encoded and enciphered and added with error correction code by the encoder  102   e,  and further modulated by the modulator  102   d  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  102   c , amplified by the large electric power amplifier  102   b , and then sent from the antenna  102   a . On the other hand, an electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  102   a , amplified by the low noise amplifier  102   h,  converted to an intermediate frequency by the frequency converter  102   i,  and decoded to a digital signal by the demodulator  102   j.  Further, by being error-corrected and deciphered and decoded by the decoder  102   k,  received data  102   l  can be obtained.  
         [0294]     In this system, the amounts of consumed electricity measured by the satellite communication send and receive unit  102  can be totaled at the base station  98 . Further, charges for public service such as electricity, city gas and city water have been totaled by visiting from house to house. However, by using the satellite traveling on the orbit in accordance with the present invention, the charges for public service can be totaled through the satellite because the satellite communication line can be easily ensured only by setting the satellite communication send and receive unit  102  even at a house surrounded by tall buildings or at a place in a mountainous region without communication means. Therefore, personnel expenses required for totaling the consumed amounts can be substantially reduced. By the effect of reducing the personnel expenses, it can be expected that the charges for public service are further reduced.  
         [0295]      FIG. 24  shows another embodiment of the satellite communication system.  
         [0296]     As shown in  FIG. 24 , the satellite communication system is composed of a group of artificial satellites  90  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; a base station  98  for performing satellite communication through the group of artificial satellites  90 ; and a satellite communication send and receive unit  103  having a function for collecting and relaying information if an information network and capable of performing communication through the group of artificial satellites  90 .  
         [0297]     As shown in  FIG. 25 , the satellite communication send and receive unit  103  is composed of an antenna  103   a , a large electric power amplifier  103   b,  a frequency converter  103   c , a modulator  103   d,  an encoder  103   e,  a multiplexer  103   f,  a low noise amplifier  103   h,  a frequency converter  103   i,  a demodulator  103   j,  a decoder  103   k,  a demultiplexer  103   l  and a network  103   n.    
         [0298]     Sent data  103   g  is multiplexed with consumed network information  103   o   1  output from the network  103   n  by the multiplexer  103   f,  encoded and enciphered and added with error correction code by the encoder  103   e,  and further modulated by the modulator  103   d  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  103   c , amplified by the large electric power amplifier  103   b,  and then sent from the antenna  103   a . On the other hand, an electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  103   a , amplified by the low noise amplifier  103   h,  converted to an intermediate frequency by the frequency converter  103   i,  and decoded to a digital signal by the demodulator  103   j.  Further, by being error-corrected and deciphered and decoded by the decoder  103   k,  received data  103   m  and network information  103   o   2  to be input to the network can be obtained.  
         [0299]     In this system, network information, for example, in an office or in a home (security, utility status/use value and control of utility) can be communicated between the satellite communication send and receive unit  103  and the base station  98 . Further, by using the satellite traveling on the orbit in accordance with the present invention, the satellite communication line can be easily ensured by an antenna unit installed in a house surrounded with tall buildings. Furthermore, in a case of security information, there is an advantage in that even if a telephone wire is cut, communicating (reporting) means can be independently ensured through the satellite.  
         [0300]      FIG. 26  shows another embodiment of the satellite communication system.  
         [0301]     As shown in  FIG. 26 , the satellite communication system is composed of a group of artificial satellites  90  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; a base station  98  for performing satellite communication through the group of artificial satellites  90 ; and a satellite communication send and receive unit  104  having a function for monitoring an environment and capable of performing communication through the group of artificial satellites  90 .  
         [0302]     As shown in  FIG. 27 , the satellite communication send and receive unit  104  is composed of an antenna  104   a,  a large electric power amplifier  104   b,  a frequency converter  104   c,  a modulator  104   d,  an encoder  104   e,  a multiplexer  104   f,  a low noise amplifier  104   h,  a frequency converter  104   i,  a demodulator  104   j,  a decoder  104   k  and a detector  104   m.  Sent data  104   g  is multiplexed with measured information  104   n  output from the detector  104   m  by the multiplexer  104   f,  encoded and enciphered and added with error correction code by the encoder  104   e,  and further modulated by the modulator  104   d  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  104   c,  amplified by the large electric power amplifier  104   b,  and then sent from the antenna  104   a.  On the other hand, an electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  104   a,  amplified by the low noise amplifier  104   h,  converted to an intermediate frequency by the frequency converter  104   i,  and decoded to a digital signal by the demodulator  104   j.  Further, by being error-corrected and deciphered and decoded by the decoder  104   k,  received data  104   l  can be obtained.  
         [0303]     By using the satellite traveling on the orbit in accordance with the present invention, communication can be easily performed at a place surrounded with tall buildings or at a place in a mountainous without any communication means. Therefore, since environmental data (weather information, water level (river, lake and so on), earthquake, volcano, carbon monoxide, nitrogen oxide, sulfur dioxide, dioxin and so on) over a wide area can be easily collected , for example, by installing the base station  98  at an environment center and placing the satellite communication send and receive units  104  having the function of monitoring environment at various regions, a speedy and appropriate measure devised to deal with a problem can be performed to protect inhabitants and environment in the district. Further, since there is little limitation in an installation place of the satellite communication send and receive unit  104 , the expense necessary for collecting the environment data can be substantially reduced. Furthermore, with regard to timing to send environmental information, there are an urgent communication method (sending is made at the time when a value of environmental data exceeds a predetermined threshold value) and a polling method (by receiving sending request of the center station, the send and receive unit automatically sends the information).  
         [0304]      FIG. 28  shows another embodiment of the satellite communication system.  
         [0305]     As shown in  FIG. 28 , the satellite communication system is composed of a group of artificial satellites  90  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; a base station  98  for performing satellite communication through the group of artificial satellites  90 ; and a satellite communication send and receive unit  105  having a function for detecting and monitoring an abnormality and capable of performing communication through the group of artificial satellites  90 .  
         [0306]     As shown in  FIG. 29 , the satellite communication send and receive unit  105  is composed of an antenna  104   a,  a large electric power amplifier  105   b,  a frequency converter  105   c , a modulator  105   d , an encoder  105   e,  a multiplexer  105   f,  a low noise amplifier  105   h , a frequency converter  105   i,  a demodulator  105   j,  a decoder  105   k  and a detector  105   m.    
         [0307]     Sent data  105   g  is multiplexed with abnormality detection information  105   n  output from the detector  105   m  by the multiplexer  105   f,  encoded and enciphered and added with error correction code by the encoder  105   e,  and further modulated by the modulator  105   d  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  105   c , amplified by the large electric power amplifier  105   b,  and then sent from the antenna  105   a . On the other hand, an electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  105   a,  amplified by the low noise amplifier  105   h , converted to an intermediate frequency by the frequency converter  105   i,  and decoded to a digital signal by the demodulator  105   j.  Further, by being error-corrected and deciphered and decoded by the decoder  105   k,  received data  105   l  can be obtained.  
         [0308]     In this system, an emergency signal can be automatically generated when an abnormality occurs. For example, by installing the base station  98  in a fire station and placing the satellite communication send and receive unit  105  having a function for monitoring an abnormality of a vehicle such as a large impact or operation of an air bag in the vehicle, when a traffic accident occurs, the fire station and an insurance company can be automatically informed of occurrence of an abnormality and accordingly the rescue activity of the driver and the passengers can be speedy performed.  
         [0309]     For example, by installing the base station  98  at the Maritime Safety Agency and by installing the satellite communication send and receive unit  105  having a function for monitoring an abnormality of a vehicle such as filling of water or an excessive impact in a ship, the Maritime Safety Agency can be automatically informed of a marine accident when it occurs and accordingly the crew and the passengers can be safely rescued.  
         [0310]      FIG. 30  shows another embodiment of the satellite communication system.  
         [0311]     As shown in  FIG. 30 , the satellite communication system is composed of a group of artificial satellites  90  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; a base station  98  for performing satellite communication through the group of artificial satellites  90 ; a satellite communication send and receive unit  108  having a function for receiving traffic information  107  from a traffic information notice system  106  such as VICS and capable of performing communication through the group of artificial satellites  90 ; and a ground communication network  95  for transmitting a request from the satellite communication send and receive unit  108  through the base station  98  to the traffic information notice system  106 .  
         [0312]     As shown in  FIG. 30 , the base station  98  sends received data obtained from the satellite communication send and receive unit  108  to the traffic information notice system  106  through the ground communication network  95  as a user&#39;s request. Further, information from the traffic information notice system  106  is sent to the used as sent data.  
         [0313]     As shown in  FIG. 31 , the traffic information notice system  106  is composed of an antenna  106   a , a large electric power amplifier  106   b , a frequency converter  106   c , a modulator  106   d , an error correction encoder  106   e,  an encipherer  106   f,  a multiplexer  106   g , a data selector  106   i,  a data memory  106   j  and the ground communication network  95 .  
         [0314]     Request information  106   k  through the ground communication network  95  is input to the data selector  106   i,  and traffic information  106   h  is cited from the data memory  106   j,  if necessary. Further, the traffic information  106   h  is maltiplexed by the multiplexer  106   g , enciphered by the encipherer  106   f,  added with error correction code by the error correction encoder  106   e,  further modulated by the modulator  106   d  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  106   c , amplified by the large electric power amplifier  106   b , and then sent from the antenna  106   a.    
         [0315]     On the other hand, as shown in  FIG. 32 , the satellite communication send and receive unit  108  is composed of an antenna  108   a , a large electric power amplifier  108   b , a frequency converter  108   c , a demodulator  108   d , an encoder  108   e,  a low noise amplifier  108   g,  a frequency converter  108   h , a demodulator  108   i,  a decipherer  108   j,  a traffic information antenna  108   n,  a traffic information receiver  108   m  and a car navigator  108   l.    
         [0316]     A request  108   f  from the user is encoded and enciphered and added with error correction code by the encoder  108   e,  and further modulated by the modulator  108   d  so as to be suitable for wireless communication, converted into a carrier wave by the frequency converter  108   c , amplified by the large electric power amplifier  108   b , and then sent from the antenna  108   a . On the other hand, an electromagnetic wave sent from the group of artificial satellites  90  is received by the antenna  108   a , amplified by the low noise amplifier  108   g , converted to an intermediate frequency by the frequency converter  108   h , and decoded to a digital signal by the demodulator  108   i.  Further, by being error-corrected and deciphered and decoded by the decoder  108   j,  traffic information as received data  108   k  can be obtained and input to the car navigator  108   l . Further, traffic information  107  is received by the traffic information antenna  108   n,  received and demodulated by the traffic information receiver  108   m  to be input to the car navigator  108   l.    
         [0317]     In this system, a driver can receive the traffic information  107  such as VICS or the like, but also request desired traffic information from the satellite communication send and receive unit  108  through the base station and the group of satellites  90 , and can obtain the information both through the traffic information notice system  106  and through the group of satellites  90 . Therefore, he can obtain detailed and timely traffic information.  
         [0318]     The traffic information includes the following information.  
         [0319]     That is, ITS information (traffic amount information, traffic speed information, traffic congestion place information, traffic congesting time information, driving environment information, stricken area information, traffic restriction information, optimum route information, information on required time in congesting time, parking lot status information, parking lot reservation information, destination information (weather, travel, sightseeing, meals, recreation information), various kinds of reservation information (public transportation, hotels, amusement facilities)), map information (map information, updated information and so on), car navigation information (car navigation information, updated information and so on), software program information (programs for car navigation, programs for game, OS and so on).  
         [0000]     (4-3) System Example 3  
         [0320]     The system example 3 is an inter-satellite communication system.  
         [0321]      FIG. 33  shows an embodiment of an inter-satellite communication system in accordance with the present invention.  
         [0322]     As shown in  FIG. 33 , the inter-satellite communication system is composed of a group of artificial satellites  90  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; a base station  98  for performing satellite communication through the group of artificial satellites  90 ; a satellite communication send and receive unit  110  and a group of artificial satellites  109 , both of which are capable of performing communication through the group of artificial satellites  90 .  
         [0323]     In this system, in a case where the group of artificial satellites  109  are traveling in such a range that the group of artificial satellites  109  can not directly communicate with the satellite communication send and receive unit  110  or the base station  98  but can communicate with the group of artificial satellites  90 , information of the group of artificial satellites  109  can be obtained through the group of artificial satellites  90 . Therefore, this system can obtain information over a wider area from the group of artificial satellites  109 .  
         [0324]     As an example, in a case where the artificial satellite  109  mounts an earth survey unit, the base station and the satellite communication send and receive unit  110  can receive the survey data, and can request desired survey data. Therefore, this system can obtain desired survey data over a wider area.  
         [0000]     (4-4) System Example 4  
         [0325]     The system example 4 is an earth survey system.  
         [0326]      FIG. 34  shows an embodiment of an earth survey system in accordance with the present invention.  
         [0327]     As shown in  FIG. 34 , the earth survey system is composed of a group of artificial satellites  111  having subsystems suitable for the elliptical orbit of the present invention such as an attitude control system, an electric source system, a communication system, a heat control system and the like; and a base station  98  for receiving a survey result sent from the group of artificial satellites  111 .  
         [0328]     In this system, since the group of artificial satellites  111  travel on the orbit in a high elevation angle to the ground, the earth survey information not affected by shielding objects can be collected by the base station  98 .  
         [0329]     Finally, service with artificial satellites in the United States will be explained briefly with reference to  FIG. 63 .  
         [0330]     What is shown in  FIG. 63  in terms of contour lines is an example of simulation result of the service time ration (%) with which any artificial satellite located in an elevation angle of 70 degree or higher can be seen with an orbital inclination angle of 45 degrees, an eccentricity squared of 0.15 and an argument of perigee of 270 degrees, in case of attempting to provide a full-time, 24 hours a day, service with two sets of four artificial satellites, that is, totally eight satellites, in the United States.  
         [0331]     As for the range of combined parameters for orbital inclination angle and eccentricity squared for two sets of four artificial satellites in the United States, the orbital inclination angle is between 40 degrees and 50 degrees and the eccentricity squared is between 0.15 and 0.25.  
         [0332]     Effects obtained by the present invention are as follows.  
         [0000]     (1) Effects with regard to the method of setting an argument of perigee  
         [0333]     According to the present invention, with regard to an elliptical orbit having orbital elements of an arbitrary orbit inclination angle, it is possible to set the orbital elements in taking into consideration change in the argument of perigee caused by the effect of gravitational field of the earth in advance.  
         [0000]     (2) Effects with regard to the method of arranging the orbits of a plurality of artificial satellites  
         [0334]     According to the present invention, in a case where communication service or broadcast service is concentratively and continuously performed using an arbitrary number of artificial satellites to a specified area on a celestial body of the artificial satellites traveling around, or where surveillance of a specified area on a celestial body of the artificial satellites traveling around or weather of the specified area is concentratively and continuously performed using an arbitrary number of artificial satellites, it is possible to easily set a semi-major axis, an eccentricity, an orbit inclination angle, an argument of perigee, a right ascension of north-bound node and a true anomaly of orbital elements of the artificial satellite.  
         [0335]     Further, according to the present invention, using an arbitrary number of artificial satellites, communication service or broadcast service can be concentratively and continuously performed to a specified area on a celestial body which the artificial satellites travel around.  
         [0336]     Furthermore, according to the present invention, using an arbitrary number of artificial satellites, surveillance of a specified area on a celestial body of the artificial satellites traveling around or weather of the specified area can be concentratively and continuously performed.  
         [0000]     (3) Effects with regard to arrangement of a plurality of artificial satellites having orbital elements employing the orbital element obtained using the above-described items (1) and (2)  
         [0337]     According to the present invention, technical requirements to the satellite communication send and receive unit imposed due to short reachable distance of electromagnetic wave can be moderated, and a communication system having s short communication delay time can be constructed.  
         [0338]     Further, according to the present invention, since the orbit is nearly circular, it is possible to shorten a breakdown time of communication and broadcast which may occur at service switching time between the plurality of artificial satellites.  
         [0339]     Furthermore, according to the present invention, since using three or four artificial satellites, the artificial satellites are arranged so that any one of the artificial satellites is coincidentally visual in the sky in an elevation angle above 70 degrees over the territory of Japan from Nemuro to Naha, communication and broadcast service to a mobile object can be easily performed using the artificial satellites.  
         [0000]     (4) Effect common to the above items from (1) to (3)  
         [0340]     According to the present invention, it is possible to provide an orbit control system by which orbit control of the artificial satellite can be performed based on the orbital element set by the method described above.  
         [0341]     To assist the understanding of the drawings, reference numerals will be explained hereinafter.  
         [0342]     In the attached figures, reference numeral  1  is an orbit projected on the ground over a span of one day from reference time,  2  . . . orbit projected on the ground over a span from 1095 day to 1096 day after reference time,  3  . . . orbit projected on the ground over a span from 2190 day to 2191 day after reference time,  4  . . . orbit projected on the ground over a span from 3650 day to 3651 day after reference time,  5  . . . setting of reference time,  6  . . . setting of number n of artificial satellites,  7  . . . setting of a temporary value of argument of perigee ω,  8  . . . setting of a temporary value of orbital semi-major axis a,  9  . . . setting of a temporary value of eccentricity e,  10  . . . setting of a temporary value of orbital inclination angle i,  11  . . . setting of temporary values of right ascension of north-bound node Ω 1  and true anomaly θ 1  of an orbit of an artificial satellite of satellite number  1 ,  12  . . . setting of temporary values of right ascension of north-bound node Ω k  and true anomaly θ k  of an orbit of an artificial satellite of satellite number k,  13  . . . simulation by a computer,  14  . . . evaluation,  15  . . . orbital elements of each artificial satellite at reference time,  17  . . . six orbital elements at reference time,  18  . . . artificial satellite tracking and control facility,  19  . . . control command,  20 ,  60 ,  61 ,  62  . . . artificial satellite,  21  . . . launch vehicle tracking and control facility,  22  . . . target injecting orbit element,  23  . . . launch vehicle,  24  . . . send and receive system,  25  . . . telemetering system,  26  . . . computer system,  27  . . . telemetry, ranging signal,  28  . . . ranging signal,  29  . . . distance and change rate of distance,  30  . . . orbit determining program,  31  . . . six orbital elements,  32  . . . orbit control program,  33  . . . attitude control variable, orbit control variable,  34  . . . command generating program,  35  . . . control command,  36  . . . command, ranging signal,  37  . . . communication system,  38  . . . data processing system,  39  . . . attitude and orbit control system,  40  . . . gas jet unit,  41  . . . attitude control actuator drive,  42  . . . thruster valve drive,  50  . . . the earth,  51  . . . equatorial plane of the earth,  63  . . . orbit of artificial satellite  60 ,  64  . . . orbit of artificial satellite  61 ,  65  . . . orbit of artificial satellite  62 ,  66  . . . ascending node of orbit  63 ,  67  . . . ascending node of orbit  64 ,  68  . . . ascending node of orbit  65 ,  70   a  . . . artificial satellite a,  70   b  . . . artificial satellite b,  70   c  . . . artificial satellite c,  70   d  . . . artificial satellite d,  71   a  . . . orbit a of artificial satellite a,  71   b  . . . orbit b of artificial satellite b,  71   c  . . . orbit c of artificial satellite c,  71   d  . . . orbit d of artificial satellite d,  72   a  . . . ascending node of orbit a,  72   b  . . . ascending node of orbit b,  72   c  . . . ascending node of orbit c,  72   d  . . . ascending node of orbit d,  90  . . . artificial satellite having subsystems such as an attitude control system, a power supply system, a communication system, a heat control system and the like suitable for an elliptical orbit in accordance with the present invention,  91  . . . base station for sending satellite broadcast through artificial satellite  90 ,  92  . . . satellite broadcast terminal for receiving satellite broadcast through artificial satellite  90 ,  93  . . . ground broadcast station,  94  . . . satellite broadcast terminal for receiving satellite broadcast through the artificial satellite  90  and a ground broadcast,  95  . . . ground communication network such as public line and cellular phone,  96  . . . base station for sending satellite broadcast through the artificial satellite  90  capable of receiving request from satellite broadcast terminal through the base communication network  95 ,  97  . . . satellite broadcast terminal having a function of communication with the ground communication network  95  and receiving satellite broadcast through the artificial satellite  90 ,  98  . . . base station performing satellite communication through the artificial satellite  90 ,  99  . . . satellite communication send and receive unit for performing satellite communication through the artificial satellite  90 ,  100  . . . artificial satellites composing a global positioning system,  101  . . . satellite communication send and receive unit having a function capable of measuring its own position using a telemetry signal from the artificial satellites composing a global positioning system and performing satellite communication through the artificial satellite  90 ,  102  . . . satellite communication send and receive unit having functions of measuring consumed amounts of electricity, city gas and city water and performing satellite communication through the artificial satellite  90 ,  103  . . . satellite communication send and receive unit having functions of collecting and relaying information of an information network and performing satellite communication through the artificial satellite  90 ,  104  . . . satellite communication send and receive unit having functions of monitoring environment and performing satellite communication through the artificial satellite  90 ,  105  . . . satellite communication send and receive unit having functions of detecting abnormality and performing satellite communication through the artificial satellite  90 ,  106  . . . traffic information informing system,  107  . . . traffic information,  108  . . . satellite communication send and receive unit having functions of receiving the traffic information  107  from the traffic information informing system  106  and performing satellite communication through the artificial satellite  90 ,  109  . . . artificial satellite,  110  . . . satellite communication send and receive unit for performing satellite communication through the artificial satellite  90 ,  111  . . . artificial satellite mounting a earth survey sensor and having subsystems such as an attitude control system, a power supply system, a communication system, a heat control system and the like suitable for an elliptical orbit in accordance with the present invention,  151  . . . artificial satellite orbit projected on the ground on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 210 degrees,  152  . . . artificial satellite orbit projected on the ground on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 230 degrees,  153  . . . artificial satellite orbit projected on the ground on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 250 degrees,  154  . . . artificial satellite orbit projected on the ground on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 270 degrees,  155  . . . artificial satellite orbit projected on the ground on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 290 degrees,  156  . . . artificial satellite orbit projected on the ground on which an artificial satellite travels with a 24-hour orbital period, an orbital inclination angle of 42.5 degrees, an eccentricity of 0.21 and an argument of perigee of 310 degrees.