JPH11325951A - Method and device for determining orbit with attitude sensor of space navigating object - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide the orbit determination value of a space navigating body by providing the orbit determination value through an autonomous orbit determining method and providing this value in on-board state. SOLUTION: An earth sensor 8 outputs a geocentric vector 9 in the machine body coordinate system of the space navigating body. Based on an attitude determination value 6 provided by an attitude determination system 5 from an inertia coordinate system to the machine body coordinate system, the coordinate of this output is transformed and a geocentric vector 10 in the inertia coordinate system is calculated. By defining this vector as an observation vector on a Kalman filter 1, an orbit determination value 4 in the inertia coordinate system is provided. In filtering processing 2, the predictive value in the geocentric direction of the inertia coordinate system is calculated by the estimated orbit determination value 4 and an earth direction model 11 and this is compared with the real geocentric vector in the inertia coordinate system so that the orbit determination value 4 can be corrected. Based on the orbit determination value determined in the past, the predictive value of the orbit determination value is found while using an orbital dynamic model 3, this is compared with the real orbit determination value and the orbit determination value 4 is corrected.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

The present invention relates to a method for determining the trajectory of a spacecraft such as a rocket or an artificial satellite. Orbit determination method by tracking from the ground and GPS (global
Unlike the orbit determination method by the positioning system, the ground tracking station and NAVSTAR (navigation)
satellite time and ranging global positionin
g system), which is an autonomous orbit determination method that does not require an orbit determination infrastructure.

[0002]

2. Description of the Related Art In a conventional method for determining the orbit of a spacecraft,
At a tracking station on the ground, the distance and direction angle to the spacecraft were measured using a large tracking antenna, and the data were processed off-line to calculate the orbit determination value of the spacecraft. Another conventional orbit determination method uses a space infrastructure such as GPS.

[0003]

However, the above-mentioned conventional system has the following problems. The first problem is that the spacecraft needs a dedicated device for orbit determination. The reason is that, in the case of the orbit determination method using a tracking station on the ground, a device for transmitting radio waves to the ground station is required, and in the case of the orbit determination method by GPS or the like, a GPS receiver or the like is required. That's why.

A second problem is that autonomous orbit determination cannot be performed. The reason is that whether the orbit determination method using a tracking station on the ground or the orbit determination method using space infrastructure such as GPS, a system required for orbit determination other than a space vehicle is required. NAVSTAR or NAV for tracking stations on the ground, GPS, etc.
This is STAR's ground operation equipment.

A third problem is that it is not possible to obtain a trajectory determination value on-board and in real time. The reason is that the orbit determination method using a tracking station on the ground includes an off-line process. Furthermore, when using this orbit data on-board, it is necessary to transmit the orbit determination value from the ground to the spacecraft, but real-time performance cannot be realized because instantaneous transmission is not possible. .

The present invention has been made under such a background, and an object of the present invention is to provide an orbit determination value by an autonomous orbit determination method that does not require any other system besides a spacecraft. Another object of the present invention is to provide a trajectory determination value that can be used in real time and onboard by realizing this trajectory determination method on-board.

[0007]

According to a first aspect of the present invention, in the method for determining the trajectory of a spacecraft, information on the earth acquired by a posture sensor mounted on the spacecraft is used. The coordinates are converted from the body coordinate system of the spacecraft to the inertial coordinate system based on the determined value, and the sensor value converted in this way is used as an observation vector, and a filter process is performed to determine a trajectory determined value. An orbit determination method for a spacecraft using a Kalman filter, characterized in that the orbit determination value is corrected by an observation model that predicts the value and the orbit determination value is corrected by an orbital dynamics model of the spacecraft flying around the earth. .

According to a second aspect of the present invention, in the trajectory determining method for a spacecraft according to the first aspect, the trajectory determining method for the spacecraft includes a method for determining a trajectory other than the equipment mounted on the spacecraft. It is characterized by the autonomous determination of the orbit without the need for equipment.

According to a third aspect of the present invention, in the trajectory determining method for a spacecraft according to the first aspect, the trajectory determining method for the spacecraft includes a process of obtaining the attitude determination value;
Equipped with a computer that performs arithmetic processing by an observation model corresponding to the attitude sensor, arithmetic processing by an orbital dynamics model of a spacecraft flying around the earth, and arithmetic processing by a Kalman filter, the It is characterized in that a determined value can be output.

According to a fourth aspect of the present invention, in the trajectory determining method for a spacecraft according to the third aspect, the trajectory determining method for the spacecraft performs necessary processing in real time. The feature is that the trajectory determination value can be output in real time.

According to a fifth aspect of the present invention, in the orbit determination device for a spacecraft, an attitude sensor for detecting information on the earth, and the information is converted from the body coordinate system of the spacecraft to the inertial coordinate system. A coordinate transformation processing device that transforms the coordinates into a coordinate, a converted sensor value, an observation model that predicts the value of the sensor after the coordinate transformation, and an orbital dynamics model of a spacecraft flying around the earth. And a Kalman filter device for processing.

According to a sixth aspect of the present invention, in the trajectory determining device for a spacecraft according to the fifth aspect, the trajectory determining device for the spacecraft includes a space navigation device as an attitude sensor for detecting information on the earth. It is characterized by using an earth sensor for detecting the direction of the earth's center of gravity with respect to the body's body coordinate system.

According to a seventh aspect of the present invention, in the trajectory determining apparatus for a spacecraft according to the fifth aspect, the trajectory determining apparatus for a spacecraft includes a spacecraft as an attitude sensor for detecting information on the earth. It is characterized by using a geomagnetic sensor that detects an earth magnetic field vector with respect to the body's body coordinate system, and using a geomagnetic model obtained from a geomagnetic potential as an observation model that predicts the value of the sensor after coordinate conversion.

[0014]

Next, embodiments of the present invention will be described in detail with reference to the drawings. FIG. 1 shows a functional block diagram of a trajectory determination method as one embodiment of the present invention. In the figure, an earth sensor 8 outputs a geocentric vector 9 in the body coordinate system of a spacecraft by observing the direction of the earth. This output is separately subjected to coordinate transformation from the body coordinate system to the inertial coordinate system by coordinate transformation 7 based on the posture determination value 6 with respect to the body coordinate system from the inertial coordinate system obtained by the posture determination system 5, The ground-center direction vector 10 in the coordinate system is calculated. By using this as the observation vector in the Kalman filter 1,
The trajectory determination value 4 in the inertial coordinate system is obtained by the filtering process 2. In the filtering process 2, the estimated orbit determination value 4 and the earth direction model 11
By calculating the predicted value of the geocentric direction in the inertial coordinate system and comparing this with the geocentric direction vector in the actual inertial coordinate system, the trajectory determination value 4 in the inertial coordinate system can be corrected. Further, based on the orbit determination value determined in the past, a predicted value of the future orbit determination value is obtained by using the orbit dynamic model 3, and this is compared with the actual orbit determination value to obtain the inertial coordinate system. Can be corrected. The Kalman filter 1 employs a sequential Kalman filter, and can calculate the orbit determination value 4 in real time every time the earth-centered vector 10 in the inertial coordinate system as an input to the filter is updated. Although the configuration of the embodiment has been described in detail above, the detailed configuration of the Kalman filter 1 in the figure is omitted because it is well known to those skilled in the art.

Next, the operation of one embodiment of the trajectory determination method shown in FIG. 1 is shown in FIGS. In the figure, FIG.
(A) to (c) show the position determination error X in the trajectory determination.
/ Y / Z component. Here, X / Y / Z means Z for the earth-center direction, and Y and Y for the directions perpendicular to the orbital plane of the spacecraft with respect to the angular momentum vector of the spacecraft with respect to the earth.
The direction defined by the right-handed system in the direction orthogonal to Z and Z is defined as X, and this coordinate system will be referred to as a trajectory coordinate system. Similarly, FIGS. 3A to 3C show the X / Y / Z components of the speed determination error in the trajectory determination. These errors are 1 as errors contained in the output of the earth sensor.
[Deg]. As is clear from the figure, a position determination error of about 20 [km] and a speed determination error of about 40 [m / s] are realized. Thereby, it was confirmed that the orbit determination value can be obtained by performing the Kalman filter processing by performing coordinate conversion on the output of the earth sensor based on the attitude determination value. That is, the effectiveness of the autonomous orbit determination method in the sense that no other system other than the equipment mounted on the spacecraft is required was confirmed.

In another embodiment of the present invention, the basic configuration is as described above, but a geomagnetic sensor may be used instead of the earth sensor. The configuration is shown in FIG. In the figure, the geomagnetic vector 13 in the body coordinate system measured by the geomagnetic sensor 12 is coordinate-transformed into the geomagnetic vector 14 in the inertial coordinate system by the coordinate conversion 7 as described above. By performing the filter processing 2 in the Kalman filter 1 on the value after the coordinate conversion, a trajectory determination value 4 in the inertial coordinate system is obtained. As an observation equation in the Kalman filter, a geomagnetic model 15 is applied, and a predicted value of a geomagnetic vector calculated from the model and a predicted value of an orbit determination value in an inertial coordinate system and a geomagnetic vector 14 in an actual inertial coordinate system are used. To correct the posture determination value 4.

Next, the operation of the trajectory determination method as another embodiment shown in FIG. 4 is shown in FIGS. FIG.
The meanings of the symbols in (a) to (c) and FIGS. 6 (a) to (c) are shown in FIGS. 2 (a) to (c) and FIGS.
The same as those in (c). As can be seen from the drawing, a position determination error of about 1500 [km] and a speed determination error of 1500 [m / s] are realized. However, the simplest dipole model is used as the geomagnetic model in this case. By using a more accurate geomagnetic model, it is possible to reduce the error of the orbit determination value.

[0018]

As described above, the present invention has the following effects. The first effect is that the trajectory of the spacecraft can be determined autonomously without requiring any other system other than the spacecraft. The reason is that the devices used in the present invention are only those mounted on the spacecraft.

The second effect is that the trajectory can be determined on-board. The reason is that the devices used in the present invention are limited to those mounted on a spacecraft.

A third effect is that the orbit of the spacecraft can be determined in real time. The reason is that a sequential Kalman filter is used as a filter process capable of real-time processing.

[Brief description of the drawings]

FIG. 1 is a functional block diagram of a trajectory determination method according to the present invention.

FIG. 2 relates to an operation example when the orbit determination method of the present invention shown in FIG. 1 is applied to a spacecraft,
It is explanatory drawing which shows a Y / Z component.

FIG. 3 relates to an operation example when the orbit determination method of the present invention shown in FIG. 1 is applied to a spacecraft,
It is explanatory drawing which shows a Y / Z component.

FIG. 4 is a functional block diagram of a trajectory determination method according to another embodiment of the present invention.

FIG. 5 is an explanatory diagram showing X / Y / Z components of position estimation accuracy in an operation example when the orbit determination method according to another embodiment of the present invention shown in FIG. 4 is applied to a spacecraft.

FIG. 6 is an explanatory diagram showing X / Y / Z components of speed estimation accuracy in an operation example when the orbit determination method according to another embodiment of the present invention shown in FIG. 4 is applied to a spacecraft.

[Explanation of symbols]

 Reference Signs List 1 Kalman filter 2 Filter processing 3 Orbital dynamic model 4 Orbit determination value 5 Attitude determination system 6 Attitude determination value 7 Coordinate transformation 8 Earth sensor 9 Geocentric vector in inertial coordinate system 10 Geocentric vector in inertial coordinate system 11 Earth direction model

Claims (7)

[Claims] 1. A method for determining the trajectory of a spacecraft, wherein information on the earth obtained by a posture sensor mounted on the spacecraft is converted from the body coordinate system of the spacecraft into inertial coordinates by the attitude determination value of the spacecraft. The coordinates are transformed into a system, the converted sensor values are used as observation vectors, and a filtering process is performed to determine the trajectory determination value.The trajectory determination value is corrected by an observation model that predicts the sensor value after the coordinate transformation, An orbit determination method for a spacecraft using a Kalman filter, wherein an orbit determination value is corrected by an orbital dynamics model of a spacecraft flying around the earth. 2. The method for determining the trajectory of a spacecraft does not require any equipment other than the equipment mounted on the spacecraft,
The orbit determination method for a spacecraft according to claim 1, wherein the spacecraft determines an orbit autonomously. 3. A method for determining the trajectory of a spacecraft, comprising: a process for obtaining a determined value of the attitude of the spacecraft; a calculation process using an observation model corresponding to a attitude sensor; and a spacecraft flying around the earth. 2. The spacecraft according to claim 1, wherein an orbit determination value can be output on-board by installing a computer that executes an arithmetic process using an orbital dynamic model of the body and an arithmetic process of a Kalman filter. Orbit determination method. 4. The method according to claim 3, wherein the orbit determination method for the spacecraft can output the orbit determination value in real time by performing necessary processing in real time.
Orbit determination method for the spacecraft described. 5. A trajectory determining device for a spacecraft mounted on a spacecraft, wherein the attitude sensor is for detecting information on the earth, and the information is converted from an airframe coordinate system of the spacecraft to an inertial coordinate system. A coordinate transformation processing device that transforms the coordinates into a coordinate, a filter processing of the converted sensor values, an observation model that predicts the values of the sensors after the coordinate transformation, and an orbital dynamics model of a spacecraft that flies around the earth. An orbit determination device for a spacecraft, comprising a Kalman filter that performs arithmetic processing by using a computer. 6. The orbit determination device for a spacecraft uses an earth sensor that detects a direction of the earth's center of gravity with respect to a body coordinate system of the spacecraft as a posture sensor for detecting information on the earth. The orbit determination device for a spacecraft according to claim 5, wherein 7. The orbit determination device for a spacecraft uses a geomagnetic sensor for detecting an earth magnetic field vector with respect to the body coordinate system of the spacecraft as a posture sensor for detecting information on the earth, and performs a coordinate conversion. The orbit determination device for a spacecraft according to claim 5, wherein a geomagnetic model obtained from a geomagnetic potential is used as an observation model for predicting the value of the sensor.