JPH01182719A - Sun acquisition apparatus of space machine - Google Patents

Abstract

PURPOSE:To obtain a compact apparatus characterized by easy adjustment, by using a sun sensor using one solar cell, rotating a space machine around its axis, which is perpendicular to the null direction of the sun sensor by a specified amount based on the output of an angle sensor, and switching two rotation control means, which are perpendicular to each other at a position, where the angle of the direction of the sun becomes the lowest angle. CONSTITUTION:Rotary axes (x) and (y) of two rotation control means are made perpendicular to each other with respect to the null direction (B) of a sun sensor 11. A coordinate system is formed with the rotary axes (x) and (y). Then, the axis of the null direction (b) of the sensor 11 passes the original point. The angle of the direction of the sun (a) is detected with the sensor 11 based on the output of one solar cell. The angular speed of the space machine is detected with an angular speed sensor 12. The output signals are inputted into an operating device 13. The rotary axis (x) and the rotary axis (y) are switched in this sequence with the device 13 through an actuator 14, and the space machine is rotated. The space machine is stopped at a position, where the angle of the direction of the sun is the lowest angle. This operation is repeated, and the position of the sun is focused to the original point. Thus, the angle of the direction of the sun is kept at zero, and the sun is kept acquired.

Description

[Detailed Description of the Invention] [Purpose of the Invention (Industrial Application Field) This invention is mounted on a spacecraft such as an artificial satellite, detects the direction of the sun, and changes the spacecraft from an arbitrary attitude to a sun-oriented attitude. This invention relates to improvements in solar capture devices used for drive control.

(Prior art) In general, solar capture for an artificial satellite involves detecting the direction of the sun with a sun sensor, and detecting the angular velocity of the sun with respect to the satellite rotation axis with an angular velocity sensor.Based on these detection outputs, a gas jet device or a reaction wheel is This is done by driving an attitude control system consisting of actuators such as The solar sensor consists of four solar cells, each of which has a predetermined angle with respect to the sensor null direction, and detects the solar direction angle by processing the output signal of each cell.

That is, the output characteristic of a solar cell is generally proportional to the cosine of the sunlight incident angle, and as shown in FIG. 4, a pair (two) of solar cells A have an angle α.

If we take the difference in daily output, we get the solid line shown in Figure 5.

In FIG. 5, S^ is the output of cell A, SR is the output of cell B, and Sl is the combined output (difference). As is clear from FIG. 5, by taking the output difference between a pair of solar cells A and B, a signal proportional to the solar direction angle θ around one axis can be obtained. Therefore, two sets of solar cells A to D are arranged orthogonally to each other, and the solar direction angle signals around the two orthogonal axes are obtained by determining the cell output difference between each set, and these signals and the angular velocity signal are It is used to capture the sun. Note that the solar direction angle θ is the angle formed between the solar sensor null direction and the sunlight incident direction.

In general, conventional solar capture devices for spacecraft require at least four solar cells oriented in different directions, making it difficult to adjust the placement of each cell. Furthermore, since the output signal of each cell must be processed, the configuration is complicated and large, which poses a major problem when mounted on a satellite.

(Problems to be Solved by the Invention) As described above, in the conventional solar capture device for a spacecraft, the solar sensor is large and complicated, and the arrangement is difficult to adjust.

The present invention was made to solve the above-mentioned problems, and an object of the present invention is to provide a sun capture device for a spacecraft in which the sun sensor is small and simple, and the adjustment is easy.

[Structure of the Invention] (Means for Solving the Problems) In order to achieve the above object, a solar capture device for a spacecraft according to the present invention includes a solar capture device that detects a solar direction angle from the output of one solar cell. a sensor, an angular velocity sensor that detects the angular velocity of the spacecraft, a first rotation control means that rotates the spacecraft at a constant rate around an axis perpendicular to the null direction of the solar sensor based on the output of the angular velocity sensor, and the angular velocity a second rotation control means for rotating the spacecraft at a constant rate around an axis perpendicular to the null direction of the sun sensor and perpendicular to the rotation axis of the first rotation control means based on the output of the sensor; and switching means for switching the first and second rotation control means at a position where the detected solar direction angle is the lowest.

(Operation) In the spacecraft sun capture device configured as described above, the rotation axes of the first and second rotation control means are respectively perpendicular to the null direction of the sun sensor and mutually perpendicular, and a coordinate system is created by each rotation axis. and the null direction axis of the sun sensor passes through the origin. When the spacecraft is rotated by the first rotation means and stopped at a position where the sun direction angle is the lowest, the sun position is on the rotation axis of the first rotation control means in the coordinate system. Here, if the spacecraft is switched to the second rotation control means and the spacecraft is rotated by the second rotation means and stopped at the position where the solar direction angle is the lowest, the sun position is determined by the second rotation control means in the above coordinate system. on the axis of rotation, that is, at the origin.

From then on, by repeating this operation, the sun position will converge to the origin, thereby reducing the sun direction angle to 0''
can be maintained and solar capture can be performed.

(Embodiment) Hereinafter, an embodiment of the present invention will be described with reference to FIGS. 1 to 3.

FIG. 1 shows the configuration of the present invention when controlling a spacecraft to a sun-oriented attitude, where 11 is a solar sensor that detects the sun direction angle θ, I2 is an angular velocity sensor that detects the angular velocity of the spacecraft, and 13 is a solar sensor that detects the angular velocity of the spacecraft. An arithmetic processing unit 14 that generates drive signals based on output signals of the sun sensor 11 and angular velocity sensor 12 is an actuator that receives drive signals from the arithmetic processing unit 13 and controls the attitude of the spacecraft.

The solar sensor 11 is composed of one solar cell, and a second
As shown in the figure, a signal proportional to the cosine of the solar direction angle θ is output. That is, the sensor output reaches a maximum when the solar direction angle θ coincides with the solar sensor null direction (direction perpendicular to the light-receiving surface of the solar cell) (θ-00), and decreases as it increases. The arithmetic processing unit 13 is composed of a small computer, etc., and inputs each output signal of the sun sensor 11 and the angular velocity sensor 12, and controls the actuator drive signal in a direction in which the solar direction angle θ obtained by the sun sensor 11 becomes smaller. be.

The operation of the above configuration will be described below with reference to FIG.

First, in the initial state, the null direction of the sun sensor 11 is made to match the sun pointing attitude of the spacecraft. Further, the arithmetic processing unit 13 creates an x-y coordinate system as shown in FIG. The origin of this coordinate system is a point on a straight line indicating the solar sensor null direction, and the X-axis and y-axis are perpendicular to the null direction.

In the above coordinate system, the solar direction angle θ is plotted as a distance from the origin on the solar direction axis, and this point is defined as the sun position. Assume that the sun position is at point A at the start of sun capture. Here, when the spacecraft is rotated a certain amount around the X axis through the actuator 14 by the arithmetic processing unit 13 based on the outputs of the sun sensor 11 and the angular velocity sensor 12, the sun position moves on an ellipse as shown in the figure. Therefore, the solar sensor 11 is stopped at a position where the output of the solar sensor 11 is maximum, that is, at a position where θ is minimum. At this time, the sun position is at point B in the figure, that is, on the X axis.

Next, the spacecraft is similarly rotated around the y-axis and stopped at a position where the output of the solar sensor 11 is maximized and θ is minimized. At this time, the sun position moves on the X-axis and stops at the 0 point, that is, the origin, on the y-axis in the figure. The position of this origin is θ-06, indicating that the sun is in the sensor null direction.

Furthermore, by sequentially repeating the detection around the X-axis and the detection around the y-axis, the sun position will always converge to the origin (0 point) even if the spacecraft is moving on an orbit, for example. As a result, the direction of the sun continues to match the direction of the sensor null, and the spacecraft captures the sun and is controlled to a sun-oriented attitude.

Therefore, the above-mentioned solar capture device is small, simple, and lightweight since the solar sensor 11 is composed of a solar battery cell, and it is only necessary to align the solar sensor null direction with the solar pointing attitude of the spacecraft, so its angle can be adjusted. It is also extremely easy.

Incidentally, since a spacecraft is normally equipped with a solar cell for power generation, a dedicated solar sensor can be omitted by using the output of this solar cell as a sensor output. Further, if an angular velocity sensor is used that is mounted on the spacecraft, there is no need to provide a special angular velocity sensor. The present invention is applicable not only to controlling a spacecraft to a sun-oriented attitude, but also to various space power generation systems, such as when directing a solar panel of an artificial satellite toward the sun.

[Effects of the Invention] As described above, according to the present invention, it is possible to provide a sun capture device for a spacecraft in which the sun sensor is small and simple, and the adjustment is easy.

[BRIEF DESCRIPTION OF THE DRAWINGS] Fig. 1 is a block diagram showing an embodiment of a sun capture device for a spacecraft according to the present invention, and Fig. 2 shows the output characteristics of the sun sensor of the same embodiment with respect to the sun direction angle. Characteristic diagram, 3rd
This figure is a diagram for explaining the operation of the same embodiment, and FIGS. 4 and 5 are diagrams for explaining the operation of the conventional solar capture device for a spacecraft, respectively. 11... Sun sensor, 12... Angular velocity sensor, 13
... Arithmetic processing unit, 14... Actuator, θ・
...Sun direction angle. Applicant's Representative Patent Attorney Takehiko Suzue Kaheki Figure 1

Claims (1)

[Claims] A solar sensor that detects the solar direction angle from the output of one solar cell, an angular velocity sensor that detects the angular velocity of the spacecraft, and a spacecraft that moves perpendicular to the null direction of the solar sensor based on the output of the angular velocity sensor. a first rotation control means for constant rotation around an axis; and a first rotation control means for rotating the spacecraft based on the output of the angular velocity sensor around an axis perpendicular to the null direction of the sun sensor and perpendicular to the rotation axis of the first rotation control means. a second rotation control means for rotating the sun at a constant speed; and a switching means for switching between the first and second rotation control means at a position where the sun direction angle detected by the sun sensor is the lowest. Capture device.