JP5130965B2 - Medium-altitude satellite acquisition method and apparatus - Google Patents

Description

  The present invention relates to a technique for capturing an intermediate altitude satellite equipped with a GPS (Global Positioning System) receiver with a ground antenna.

  In order to track and capture an artificial satellite with an antenna, the motion of the satellite is estimated by processing the data measured from the ground station to the satellite using radio waves (ie, orbit determination), and the equation of motion is Based on the prediction of future motion (ie, orbit prediction) based on this, a procedure of creating information (ie, antenna forecast value) for driving the antenna is required.

  In recent years, GPS receivers are mounted on medium-altitude satellites with altitudes of 500-1000 km, and the positions of the satellites have been determined onboard. This on-board decision value is mainly used for autonomous orbit and attitude control of artificial satellites. In the case of an earth observation satellite, the position data of the artificial satellite may be downloaded to the ground as telemetry data and used to process the earth observation data.

For example, Japanese Patent No. 3348835 discloses a technique that uses position information of an on-board GPS receiver for tracking acquisition of a flying object. This publication is a method of tracking an object moving in the atmosphere (for example, an airship). Basically, it takes a long time to capture and track with an antenna, and the average rate of change of past position data is calculated only when it is out of tracking. Apply to make predictions. The movement of an artificial satellite is described by the equation of motion of an object orbiting around the earth, and is different from the situation of an airship in that the acquisition time of the ground antenna is short and the prediction time is overwhelmingly longer. For example, a medium-altitude artificial satellite orbits the earth 13-15 times a day, but it is 4-5 times visible from one ground station, and the visible time is about 10 minutes per circuit. Therefore, after capturing at one station, it is necessary to perform trajectory prediction of at least 10 hours and in some cases 20 hours or more in order to acquire at the same station. The above publication cannot cope with such a situation.
Japanese Patent No. 3348835

  Conventionally, generating an antenna forecast value as it is using an on-board decision value has not been performed for the following reason. In other words, it is necessary to predict the motion of the satellite to generate the antenna forecast value. Since the on-board decision value is the same principle as car navigation, the position at that moment can be obtained with an accuracy of several tens of meters, but the future motion cannot be accurately predicted.

  Therefore, at present, the position information of the onboard decision value is downloaded as telemetry data for several hours to several tens of hours, and the trajectory is estimated using the data on the ground. In other words, the onboard decision value is used as measurement data for orbit determination. Since the trajectory determination process is an iteration based on the method of least squares, there is a problem that it is difficult to completely eliminate the operator's involvement because the process itself is heavy and may diverge.

  Accordingly, an object of the present invention is to provide a technique for effectively utilizing data from a GPS receiver mounted on a medium-height satellite so that the medium-height satellite can be captured by a ground antenna. It is.

  According to a first aspect of the present invention, there is provided a method for capturing an intermediate satellite from a position and velocity data of the intermediate satellite obtained from a GPS receiver mounted on the intermediate satellite. Calculating the value of the orbital length radius contained in the Kepler element from the multiple values of the orbital period obtained from the Kepler element and the step of substituting it, and replacing it with the multiple values of the orbital period Calculating an average orbit radius change rate using the average motion value obtained, calculating an atmospheric drag coefficient using the average orbit radius change rate, the atmospheric density, and the orbit length radius, and the Kepler element Value and atmospheric resistance coefficient are used to predict the orbit of the medium-height satellite, and based on the result of the orbit prediction, the antenna prediction value used for capturing the medium-height satellite by the antenna And calculating.

  In this way, it is possible to appropriately control the antenna on the ground for acquiring the medium-altitude satellite using the data from the GPS receiver without increasing the calculation load on the ground. Become.

  The Kepler element calculation step described above includes the steps of converting the position and velocity data of the medium altitude satellite into position and velocity data in the equatorial plane coordinate system, and the position and velocity data in the equatorial plane coordinate system. A step of calculating a value of the element.

  The above method may be executed by a combination of a computer and a program executed by the computer, and the program is stored in, for example, a flexible disk, a CD-ROM, a magneto-optical disk, a semiconductor memory, a hard disk, or the like. Stored in a medium or storage device. Moreover, it may be distributed as a digital signal via a network or the like. The intermediate processing result is temporarily stored in a storage device such as a main memory.

  According to the present invention, data from a GPS receiver mounted on a medium-altitude artificial satellite can be effectively used to capture the medium-high-level artificial satellite with a ground antenna.

  FIG. 1 shows a system outline diagram according to an embodiment of the present invention. The medium-altitude artificial satellite 1 includes a GPS receiver, and is assumed to be visible on the antenna 3 on the ground at the position 1a, for example. The medium-altitude artificial satellite 1 has a function of transmitting the position and velocity data acquired by the GPS receiver and the data of the orbital period Pi (i = 1 to m) measured every orbit to the antenna 3 as telemetry data. Also have.

  On the ground, the antenna 3 is connected to a known telemetry processing device 5 that processes telemetry data received from the medium-altitude artificial satellite 1 that has become visible. The telemetry processing device 5 uses a GPS receiver from the telemetry data. Position / velocity data and orbital period Pi data are extracted. The telemetry processing device 5 is connected to the antenna forecast value generation device 7 which is the main part in the present embodiment, and the position and velocity data by the GPS receiver and the data of the orbital period Pi are supplied to the antenna forecast value generation device 7. Output. The antenna forecast value generation device 7 performs the process described in detail below to generate an antenna forecast value. The antenna forecast value generation device 7 is connected to a known antenna drive device 9 and outputs an antenna forecast value to the antenna drive device 9. The antenna driving device 9 is connected to the antenna 3 and controls the antenna 3 based on the antenna prediction value so that the medium-height satellite can be captured at the next visible position 1b of the medium-height satellite 1.

  FIG. 2 shows a functional block diagram of the antenna predicted value generation apparatus 7 that performs the main processing in the present embodiment. The antenna forecast value generation device 7 includes a data acquisition unit 71 that receives position and velocity data and data of the orbital period Pi by the GPS receiver from the telemetry processing device 5, and acquisition data that stores data acquired by the data acquisition unit 71. A storage unit 72; a coordinate conversion unit 73 that performs a coordinate conversion process on the data stored in the acquired data storage unit 72; an equatorial plane coordinate system data storage unit 74 that stores a processing result of the coordinate conversion unit 73; An average element generation unit 75 that calculates a value of a Kepler element (also referred to as an average element) using data stored in the equatorial plane coordinate system data storage unit 74, and an average element that stores a processing result of the average element generation unit 75 The data storage unit 76 and the data stored in the acquired data storage unit 72 are used to calculate the trajectory length radius, and the average element data storage unit 76 is overwritten. The trajectory length radius calculation unit 77 to be stored, the average motion data storage unit 78 for storing data in the process of the trajectory length radius calculation unit 77, the data stored in the acquired data storage unit 72 and the average motion data storage unit The trajectory length radius change rate calculation unit 79 that calculates the trajectory length radius change rate using the data stored in 78, and the trajectory length radius change rate value calculated by the trajectory length radius change rate calculation unit 79 are stored. An orbital length radius change rate storage unit 80, an atmospheric density model data storage unit 81 that stores atmospheric density model data, and an atmosphere that calculates the value of the atmospheric density using data stored in the atmospheric density model data storage unit 81 A density calculation unit 82, an atmospheric density storage unit 83 that stores the value of the atmospheric density calculated by the atmospheric density calculation unit 82, an average element data storage unit 76, an atmospheric density storage unit 83, and a gauge An atmospheric resistance coefficient calculation unit 84 that calculates an atmospheric resistance coefficient using data stored in the long radius change rate storage unit 80, and an atmosphere that stores the value of the atmospheric resistance coefficient calculated by the atmospheric resistance coefficient calculation unit 84 A trajectory prediction unit 86 that performs trajectory prediction by a general perturbation method using data stored in the resistance coefficient storage unit 85, the average element data storage unit 76, and the atmospheric resistance coefficient storage unit 85, and a trajectory prediction unit 86 A trajectory prediction data storage unit 87 that stores the generated trajectory prediction data, and an antenna prediction value calculation unit 88 that calculates an antenna prediction value using data stored in the trajectory prediction data storage unit 87.

  Next, processing contents of the antenna forecast value generation device 7 shown in FIG. 2 will be described with reference to FIGS. First, the data acquisition unit 71 receives GPS measurement values (position and velocity data) and orbit period data from the telemetry processing device 5, and stores them in the acquisition data storage unit 72 (step S1). Next, the coordinate conversion unit 73 converts the GPS measurement value of the WGS 84 system stored in the acquired data storage unit 72 into a value of the equatorial plane coordinate system, and the converted data is the equatorial plane coordinate system data storage unit 74. (Step S3). The WGS84 system is a fixed earth coordinate system. As shown in Fig. 4, the X axis (XR) is a straight line passing through the intersection of the equatorial plane and the Greenwich meridian and the Earth's center of gravity. The axis parallel to the direction of the conventional international origin is defined as the Z axis (ZR), and the Z axis, the X axis, and the Y axis (YR) forming a right-handed orthogonal system are defined. Since the fixed earth coordinate system such as the WGS84 system is a coordinate system that rotates with the earth, it is necessary to convert it to an inertial coordinate system in order to handle the equation of motion. Usually, an artificial satellite uses an inertial coordinate system called an equatorial plane coordinate system, and is converted to this coordinate system. In the equatorial plane coordinate system, as shown in FIG. 4, the equatorial plane direction on the equator plane is the X axis (XI), the direction perpendicular to the equator plane is the Z axis (ZI), and the direction perpendicular to the X axis and the Z axis is It is defined by the Y axis (YI). Such coordinate transformation processing is well known, and is disclosed, for example, in “Aerospace Guidance and Control” published by the Society of Instrument and Control Engineers 1995.

  Further, the average element generation unit 75 converts the values of the equatorial plane coordinate system stored in the equatorial plane coordinate system data storage unit 74 into Kepler elements (ie, average elements), and stores the converted data as average element data. The data is stored in the unit 76 (step S5). The GPS measurement value includes a periodic fluctuation of perturbation acting on the artificial satellite. Perturbation refers to the part of the forces acting on the actual movement of an artificial satellite that deviates from the two-body approximation (an ideal model that treats the satellite and the earth as mass points). (A) Earth shape is not a perfect sphere (particularly flat factor is large), (b) Sun / moon attractive force, (c) Atmospheric resistance (cannot be ignored as a force acting continuously on medium altitude satellites), (D) There is solar radiation pressure. In the case of medium-altitude satellites, the maximum contribution of periodic fluctuations is the effect of earth flatness (referred to as J2 term), so an average element is generated by subtracting this amount. As shown in FIGS. 5 and 6, these average elements are various parameters when the orbit of the artificial satellite is approximated as a simple elliptical orbit. As shown in FIG. 5, among the end points of a line segment that passes through the center of the earth 1000 and connects two points on the orbit 1002 and has the maximum length, a point closer to the earth 1000 is a peripheral point. The farther point is called the apogee point. The distance between the center of the Earth 1000 and the Perige point is called the Perige radius, and the distance between the center of the Earth 1000 and the Apogee point is called the Apogee radius. The track length radius a is half the length of the line segment. That is, the length of the line segment is twice the track length radius a. Furthermore, when the eccentricity is e, the apogee radius is calculated as a × (1 + e), and the periphery radius is calculated as a × (1−e). A vector having a length e from the center of the earth 1000 toward the peripheral point is referred to as an e vector.

  As shown three-dimensionally in FIG. 6, the point where the artificial satellite penetrates the equatorial plane of the earth 1000 from the bottom to the top is called the ascending intersection, and the angle formed by the plane of the orbit 1002 and the equatorial plane is the orbital inclination angle i. Call it. In addition, the angle of a straight line connecting the center of the earth 1000 and the ascending intersection with the equinox of the earth 1000 as a reference is called ascending intersection red meridian Ω. In addition, an angle formed by a straight line connecting the center of the earth 1000 and the rising intersection and the e vector is referred to as a near point argument ω. Further, an angle formed by the e vector and a straight line connecting the artificial satellite and the center of the earth 1000 is referred to as a near point separation angle f.

  The average element shown in this way is calculated. The specific calculation method is well known, and is disclosed in, for example, the above-mentioned literature, “Artificial Satellite Orbit”, Yoshihide Kozai, Space Development Agency, Central Tracking Control Center, and 1979. Further explanation is omitted.

  Next, the trajectory length radius calculation unit 77 calculates the average motion ni from the trajectory period Pi (i = 1 to m) stored in the acquired data storage unit 72, and calculates the trajectory length radius a from the average value of ni. Calculated and overwritten and saved in the average element data storage unit 76 so as to replace the orbital length radius in the average element calculated in step S5 (step S7). For the orbital period Pi, the orbital period is counted at the timing of the ascending intersection in the artificial satellite, for example, data of about 100 laps is accumulated, downloaded as telemetry data, and used for this processing.

Here, the average motion ni of each lap is expressed as follows. The average motion ni data is stored in the average motion data storage unit 78.
ni = 2π / Pi (i = 1 to m)
Further, the average value of the average motion is expressed as follows.

そして、平均運動の平均値を用いて軌道長半径は、以下のとおりに計算される。

The orbital length radius is calculated as follows using the average value of the average motion.

但し、μは地心重力定数である。このように算出した軌道長半径の方が精度が高いので、このようにして算出された軌道長半径ａを以下の処理では採用する。

Where μ is the geocentric gravity constant. Since the orbital length radius calculated in this way has higher accuracy, the orbital length radius a thus calculated is adopted in the following processing.

  Further, the track length radius change rate calculation unit 79 calculates the average track length radius change rate adot from the change rate Δni of the average motion ni stored in the average motion data storage unit 78, and the track length radius change rate storage unit. It stores in 80 (step S9).

The rate of change Δni of the average motion ni for each turn is calculated as follows.
Δni = n _{i + 1} −n _i
Moreover, the orbital length radius for each round is calculated as follows.

Moreover, when converted into a change in the orbital length radius for each round, it is expressed as follows.
Δai = − (2ai / 3ni) Δni (i = 1 to m−1)
And the average orbital length radius change rate is expressed as follows.

  On the other hand, the atmospheric density calculation unit 82 calculates the atmospheric density ρ using the data stored in the atmospheric density model data storage unit 81 and stores it in the atmospheric density storage unit 83 (step S11). Since this calculation is well known, further explanation is omitted. For details, see “MSIS-86 THERMOSPHERIC MODEL,” Hedin. A. E. , Journal of geophysical research, Vol.92, No. A5, 1987.

  Further, the atmospheric resistance coefficient calculation unit 84 includes an average orbital length radius change rate adot, an atmospheric density ρ, and an orbital length radius change rate storage unit 80, an atmospheric density storage unit 83, and an average element data storage unit 76. An atmospheric resistance coefficient is calculated from the orbital length radius a, and the atmospheric resistance coefficient is stored in the atmospheric resistance coefficient storage unit 85 (step S13).

Here, acceleration due to atmospheric resistance is expressed by the following equation.
αdrag = C _D SρV ² / 2W
Where C _D is the satellite resistance coefficient (coefficient due to surface characteristics), S is the effective cross-sectional area of the artificial satellite, W is the satellite mass, V = (μ / a) ^1/2 , The average speed of the satellite. From the relationship between the orbital length radius change rate and atmospheric resistance force acceleration, the following equation holds.
adot = C _D SρVa / W
Therefore, the atmospheric resistance coefficient C _D S / W is expressed by the following equation.
C _D S / W = adot / (ρVa) = adot / (ρa (μ / a) ^1/2 )
The atmospheric resistance coefficient C _D S / W is calculated by this equation and stored in the atmospheric resistance coefficient storage unit 85.

  Then, the trajectory prediction unit 86 calculates a general perturbation method (for example, SGP4 / SDP4) from the value of the average element stored in the average element data storage unit 76 and the atmospheric resistance coefficient stored in the atmospheric resistance coefficient storage unit 85. ) Is performed, and the result of trajectory prediction is stored in the trajectory prediction data storage unit 87 (step S15). Orbit prediction by the general perturbation method is well known, for example, see http://www.celestrak.com/software/tskelso-sw.asp.

  Further, the antenna prediction value calculation unit 88 calculates an antenna prediction value (eg, an azimuth / elevation angle necessary for capturing an artificial satellite) using the orbit prediction data stored in the orbit prediction data storage unit 87, and the antenna driving device. 9 (step S17). The method of calculating the antenna forecast value is well known, see the Japan Aerospace Exploration Agency, orbit determination program, algorithm specifications, published in 1994.

  Then, the antenna driving device 9 drives the antenna by a known method from the antenna forecast value so as to capture the artificial satellite at the timing when the next artificial satellite becomes visible.

  By performing the processing as described above, the orbit determination processing on the ground becomes unnecessary. The processing performed on the ground is basically only algebraic calculation as described above, and there is no concern about divergence and the like, so that it can be fully automated.

  Further, conventionally, it has been necessary to download an on-board determination value from a GPS receiver over several hours, which is a load in telemetry processing, but in this embodiment, only a few GPS receivers are used. Since the on-board decision value and about 100 orbit period information are sufficient, the load on the telemetry process can be reduced. In addition, since the processing load of the entire system combining artificial satellites and the ground is reduced as a whole, it is effective for artificial satellites with limited operational resources such as small artificial satellites.

  Although the embodiment of the present invention has been described above, the present invention is not limited to this. For example, the functional block diagram of the antenna forecast value generation device 7 is an example, and does not necessarily match the actual program configuration.

  Further, the antenna forecast value generation device 7 is a computer device. In this computer device, as shown in FIG. 7, a memory 2501 (storage unit), a CPU 2503 (processing unit), a hard disk drive (HDD) 2505, A display control unit 2507 connected to the display device 2509, a drive device 2513 for a removable disk 2511, an input device 2515, and a communication control unit 2517 for connecting to a network are connected by a bus 2519. Application programs including the OS and the Web browser are stored in the HDD 2505, and are read from the HDD 2505 to the memory 2501 when executed by the CPU 2503. If necessary, the CPU 2503 controls the display control unit 2507, the communication control unit 2517, and the drive device 2513 to perform necessary operations. Further, data in the middle of processing is stored in the memory 2501 and stored in the HDD 2505 if necessary. Such a computer realizes various functions as described above by organically cooperating hardware such as the CPU 2503 and the memory 2501 described above with the OS and necessary application programs.

(Appendix 1)
A Kepler element calculation step for calculating a Kepler element value from the position and velocity data of the intermediate altitude satellite obtained from a GPS receiver mounted on the medium altitude satellite;
The value of the orbital length radius of the intermediate altitude satellite is calculated from a plurality of values of the orbital period obtained from the intermediate altitude artificial satellite, and the orbital length radius of the Kepler element calculated in the Kepler element calculating step is calculated. A trajectory length radius calculating step for replacing with a trajectory length radius;
An average orbit radius change rate is calculated using an average motion value obtained from a plurality of values of the orbit period, the average orbit radius change rate, the atmospheric density, and the orbit length radius calculated in the orbit length radius calculating step. And calculating the atmospheric resistance coefficient using
Using the Kepler element value with the orbital length radius replaced and the atmospheric resistance coefficient, orbit prediction of the medium-height satellite is performed, and based on the orbit prediction result, Calculating an antenna predicted value used for acquisition;
A computer-implemented method for capturing mid-satellite satellites.

(Appendix 2)
In the Kepler element calculation step,
Converting the medium altitude satellite position and velocity data into position and velocity data in the equatorial plane coordinate system;
From the position and velocity data in the equatorial plane coordinate system, the value of the Kepler element is calculated.
The method for capturing a medium-altitude satellite described in Appendix 1.

(Appendix 3)
A Kepler element calculation step for calculating a Kepler element value from the position and velocity data of the intermediate altitude satellite obtained from a GPS receiver mounted on the medium altitude satellite;
The value of the orbital length radius of the intermediate altitude satellite is calculated from a plurality of values of the orbital period obtained from the intermediate altitude artificial satellite, and the orbital length radius of the Kepler element calculated in the Kepler element calculating step is calculated. A trajectory length radius calculating step for replacing with a trajectory length radius;
An average orbit radius change rate is calculated using an average motion value obtained from a plurality of values of the orbit period, the average orbit radius change rate, the atmospheric density, and the orbit length radius calculated in the orbit length radius calculating step. And calculating the atmospheric resistance coefficient using
Using the Kepler element value with the orbital length radius replaced and the atmospheric resistance coefficient, orbit prediction of the medium-height satellite is performed, and based on the orbit prediction result, Calculating an antenna predicted value used for acquisition;
A medium-altitude satellite acquisition program to run a computer.

(Appendix 4)
Kepler element calculation means for calculating the value of the Kepler element from the position and velocity data of the intermediate altitude satellite obtained from the GPS receiver mounted on the medium altitude satellite,
The value of the orbital length radius of the intermediate altitude satellite is calculated from a plurality of values of the orbital period obtained from the intermediate altitude artificial satellite, and the orbital length radius of the Kepler element calculated in the Kepler element calculating step is calculated. Means to replace the orbital length radius;
An average orbit radius change rate is calculated using an average motion value obtained from a plurality of values of the orbit period, the average orbit radius change rate, the atmospheric density, and the orbit length radius calculated in the orbit length radius calculating step. Means for calculating the atmospheric resistance coefficient using
Using the Kepler element value with the orbital length radius replaced and the atmospheric resistance coefficient, orbit prediction of the medium-height satellite is performed, and based on the orbit prediction result, Means for calculating an antenna predicted value used for acquisition;
A medium-altitude satellite capture device.

It is a system outline figure concerning an embodiment of the invention. It is a functional block diagram of an antenna forecast value generation device. It is a figure which shows the main flow in embodiment of this invention. It is a figure for demonstrating coordinate transformation. It is a figure for demonstrating a Kepler element. It is a figure for demonstrating a Kepler element. It is a functional block diagram of a computer.

Explanation of symbols

DESCRIPTION OF SYMBOLS 1 Artificial satellite 3 Antenna 5 Telemetry processing apparatus 7 Antenna forecast value generation apparatus 9 Antenna drive apparatus 71 Data acquisition part 72 Acquisition data storage part 73 Coordinate conversion part 74 Equatorial plane coordinate system data storage part 75 Average element generation part 76 Average element data storage Unit 77 orbital length radius calculation unit 78 average motion data storage unit 79 orbital length radius change rate calculation unit 80 orbital length radius change rate storage unit 81 atmospheric density model data storage unit 82 atmospheric density calculation unit 83 atmospheric density storage unit 84 atmospheric resistance coefficient Calculation unit 85 Atmospheric resistance coefficient storage unit 86 Orbit prediction unit 87 Orbit prediction data storage unit 88 Antenna forecast value calculation unit

Claims (4)

A Kepler element calculation step for calculating a value of a Kepler element including an orbital length radius from the position and velocity data of the intermediate altitude satellite obtained from a GPS receiver mounted on the medium altitude artificial satellite;
A plurality of values of the orbital period obtained from highly satellites in said calculates a semi-major axis of the middle altitude satellites, the semi-major axis in the value of the Kepler elements calculated by the Keplerian element calculation step, A trajectory length radius calculating step for replacing the calculated trajectory length radius;
An average orbit radius change rate is calculated using an average motion value obtained from a plurality of values of the orbit period, the average orbit radius change rate, the atmospheric density, and the orbit length radius calculated in the orbit length radius calculating step. And calculating the atmospheric resistance coefficient using
Using the Kepler element value with the orbital length radius replaced and the atmospheric resistance coefficient, orbit prediction of the medium-height satellite is performed, and based on the orbit prediction result, Calculating an antenna predicted value used for acquisition;
A computer-implemented method for capturing mid-satellite satellites. In the Kepler element calculation step,
Converting the medium altitude satellite position and velocity data into position and velocity data in the equatorial plane coordinate system;
From the position and velocity data in the equatorial plane coordinate system, the value of the Kepler element is calculated.
A method for capturing a medium-altitude satellite according to claim 1. A Kepler element calculation step for calculating a value of a Kepler element including an orbital length radius from the position and velocity data of the intermediate altitude satellite obtained from a GPS receiver mounted on the medium altitude artificial satellite;
A plurality of values of the orbital period obtained from highly satellites in said calculates a semi-major axis of the middle altitude satellites, the semi-major axis in the value of the Kepler elements calculated by the Keplerian element calculation step, A trajectory length radius calculating step for replacing the calculated trajectory length radius;
An average orbit radius change rate is calculated using an average motion value obtained from a plurality of values of the orbit period, the average orbit radius change rate, the atmospheric density, and the orbit length radius calculated in the orbit length radius calculating step. And calculating the atmospheric resistance coefficient using
Using the Kepler element value with the orbital length radius replaced and the atmospheric resistance coefficient, orbit prediction of the medium-height satellite is performed, and based on the orbit prediction result, Calculating an antenna predicted value used for acquisition;
A medium-altitude satellite acquisition program to run a computer. Kepler element calculation means for calculating a value of a Kepler element including an orbital length radius from the position and velocity data of the intermediate altitude satellite obtained from a GPS receiver mounted on the medium altitude artificial satellite;
A plurality of values of the orbital period obtained from highly satellites in said calculates a semi-major axis of the middle altitude satellites, the semi-major axis in the value of the Kepler elements calculated by the Keplerian element calculation step, Means for replacing the calculated orbital length radius;
An average orbit radius change rate is calculated using an average motion value obtained from a plurality of values of the orbit period, the average orbit radius change rate, the atmospheric density, and the orbit length radius calculated in the orbit length radius calculating step. Means for calculating the atmospheric resistance coefficient using
Using the Kepler element value with the orbital length radius replaced and the atmospheric resistance coefficient, orbit prediction of the medium-height satellite is performed, and based on the orbit prediction result, Means for calculating an antenna predicted value used for acquisition;
A medium-altitude satellite capture device.

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