JPH0470198B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION [Field of Industrial Application] The present invention relates to an attitude control device for an earth-oriented artificial satellite launched into a geostationary orbit.

[Conventional technology]

FIG. 4 is a block diagram for explaining a conventional attitude control device for an earth-oriented artificial satellite (hereinafter simply referred to as a satellite) using an earth sensor. In FIG. 4, 1 is the target attitude toward the center of the earth (pitch angle θ or roll angle φ = 0), 2 is the attitude of the satellite (pitch angle θ or roll angle φ), and the earth sensor 3 After the attitude 2 of the satellite is detected, the output 4 is negatively fed back to the controller 5, converted into a drive signal for the actuator 6 such as a wheel or thruster, and finally through the satellite dynamics 7, the attitude 2 of the satellite is 0. Control is performed so that

Figure 5 shows the reference axis and attitude of the satellite attitude.
FIG. 2 is a diagram for explaining the three Euler angles (pitch, roll, and yaw angles). In Figure 5, 13
is the orbital plane, 14 is the earth, 15 is the satellite, 16 is the direction perpendicular to the orbital plane, 21 is the direction of the center of the earth as seen from the satellite,
22 is a direction in the orbital plane perpendicular to 16 and 21, 2
Coordinate axes 3 to 25 are fixed to the satellite body, with 23 being the Y axis (pitch axis), 24 being the X axis (roll axis), and 25 being the Z axis (yaw axis). Also 22,
16, 21 and the X, Y, and Z axes are three Euler angles, and the rotation angles around the Y, X, and Z axes are respectively pitch angles
6, roll angle φ27, and yaw angle φ28. However, since it is assumed below that θ, φ, and ψ are minute angles, the order of coordinate rotation does not matter here.

Now, in FIG. 5, since the earth sensor in principle detects attitude angles θ and φ around two axes based on the earth center direction 21 as seen from the satellite, θ and φ are detected based on the output signal of the earth sensor. When the attitude control is performed, the Z-axis 25 of the satellite is always directed toward the center of the earth 21. In other words, the Y-axis of the satellite is controlled so as to always form an angle of 90° with the direction 21 of the earth's center.
Furthermore, since the yaw angle ψ28 of the satellite is also normally controlled separately, actively or passively, the Y-axis 23 of the satellite is ultimately perpendicular to the orbital plane in an ideal state with no attitude sensor error or control error. direction 16
will be held.

[Problem that the invention seeks to solve]

The main purpose of attitude control for earth-oriented artificial satellites in geostationary orbit is to keep their onboard mission equipment, such as communication and broadcasting antennas and observation cameras, always pointing toward target areas on the earth.

However, in conventional attitude control systems, when the orbit of a geostationary satellite deviates from the equatorial plane (that is, it has an orbital inclination angle), the orbit period (= daily period) changes in the north-south direction as the satellite orbits. There was a problem in that pointing changes occurred.

FIGS. 6 and 7 are diagrams explaining the above-mentioned problems in detail. In Figure 6, 12 is the equatorial plane, 13 is the orbital plane, 29 is the orbital inclination plane, 30 is the direction of the ascending node (the point where the orbit crosses the equatorial plane from south to north), and 19 is the mission equipment onboard the satellite. 31
20 is the latitude of the target pointing point.
In addition, A, B, C, and D in the figure are 19 when the latitude argument η (the phase angle of the satellite 15 measured in the orbital plane from the ascending node direction 30) is 0°, 90°, 180°, and 270°, respectively.
and the earth's surface, and in the case of a geostationary satellite, the orbital angular velocity of the satellite and the rotational angular velocity of the earth are equal, so these points A, B, C, and D represent points on the same meridian on the earth. There is. As can be seen from FIG. 6, as the satellite orbits, the pointing point of the onboard mission equipment shifts in the north-south direction with respect to the target pointing point at an orbital period (=1 day period).

FIG. 7 shows the time waveform of the pointing error in the north-south direction generated in this way. If minute terms of second order or higher are ignored, the pointing error e _NS in the north-south direction is expressed by the following equation.

e _NS = a _sio η (1) where a is the amplitude of the daily fluctuation of e _NS , η is the latitude argument mentioned above, and is approximately calculated at any time t by the following equation.

η=ω ₀ (t-t ₀ )=ω ₀ t+b ...(2) where ω ₀ is the orbital angular velocity, t ₀ is the time of passing the ascending node, and b (=-ω ₀ t ₀ ) is t=0 is the value of η (hereinafter referred to as the initial phase angle) at The amplitude a of pointing fluctuation is proportional to the orbital inclination angle i30. That is, a=ki...(3) Here, k is a constant that depends on the latitude δ of the target pointing point, and k≒0.14 for δ=30°N.
(deg/deg).

As is clear from the above explanation, when the orbital plane of the satellite is tilted from the equatorial plane, a sinusoidal pointing error in the north-south direction occurs with an amplitude proportional to the orbital inclination angle. Therefore, the higher the pointing precision required of the satellite, the
In order to reduce the pointing error caused by the orbital inclination angle, it is necessary to keep the permissible range of the orbital inclination angle small, which causes problems such as the need to frequently perform orbit maintenance control in the north-south direction. It was hot.

This invention was made to solve this problem, and aims to prevent pointing errors in the north-south direction from occurring even if the orbit of the satellite is tilted from the equator plane.

[Means for solving problems]

The attitude control device for an artificial satellite according to the present invention is provided with a signal generator that generates a sinusoidal signal that changes with the orbit period, and uses the sum signal of the output of the signal generator and the roll output of the earth sensor as a reference for the controller. This is the input.

[Effect]

In this invention, by using the roll output of the earth sensor as the reference input for direct control of the attitude control system, the direction fluctuation in the north-south direction that occurs when the roll attitude of the satellite is maintained at zero is reversed with the same amplitude. By generating a phase sine wave signal with a signal generator and using the sum signal of this signal and the roll output of the earth sensor as the reference input of the controller,
This realizes a roll attitude that changes in a sinusoidal manner so that no north-south pointing variation occurs even when the orbital inclination angle is not zero.

〔Example〕

The present invention will be specifically explained below using an embodiment shown in FIGS. 1 to 3.

FIG. 1 is a block diagram of a satellite attitude control device (roll system) according to the present invention. Conventional equipment (Figure 4) is used for the satellite pitch system.
Regarding the yaw system, we have developed a method (so-called bias momentum method) that provides bias angular momentum around the Y axis (pitch axis) and passively controls it without using a yaw sensor using dynamic roll/yaw coupling. Suppose.

In FIG. 1, 1 to 7 are exactly the same as the conventional device described above, 8 is a command input from the ground (amplitude a and initial phase angle b), and 9 is a sinusoidal signal 1.
1 is a signal generator according to the present invention that generates a zero.
Sine wave signal 10 and earth sensor roll output φ
The addition output 11 with 4 is negatively fed back to the controller 5. The controller drives the actuator 6, and control is performed via the attitude dynamics 7 of the satellite so that the added output 11 eventually becomes 0. As a result, the attitude of the satellite (roll angle φ)2 is
Excluding earth sensor errors, control errors (including the influence of disturbance torque), etc., the time-varying waveform has the same amplitude and opposite phase as the sinusoidal signal 10 output from the signal generator 9. That is, φ=−a _sio (ω ₀ t+b)=−a _sio η ...(4) FIG. 2 is an explanation showing the inertial space of the satellite and its attitude with respect to the earth realized by the embodiment shown in FIG. 1. It is a diagram. In Figure 2, 12 is the equatorial plane,
13 is the orbital plane, 14 is the earth, 15 is η=90°,
The satellite at 270° (corresponding to 6 hours and 18 hours after passing the ascending node), 16 is the direction perpendicular to the orbital plane, and 17 is the Y axis (pitch axis) of the satellite. According to the embodiment shown in Fig. 1, the angle (=90° + φ) between the Y-axis and the direction of the earth's center as seen from the satellite changes sinusoidally, and the yaw changes due to the bias angular momentum around the Y-axis. Since the attitude is passively controlled, a constant Y-axis attitude in inertial space is achieved regardless of the orbital position of the satellite. Therefore, the Y axis 17 and the direction perpendicular to the raceway surface 16
The angle 18 formed by the angle 18 remains constant and its magnitude is equal to the amplitude a of the sinusoidal signal 10 (FIG. 1). Therefore, if this amplitude a is selected to be equal to the amplitude a of the pointing error in the north-south direction that occurs when a conventional attitude control device is used for the roll system (see Figure 7), the satellite The pointing direction 19 of the onboard mission equipment can be directed to a target point on the same latitude 20 on the earth.

FIG. 3 is an explanatory diagram, corresponding to FIG. 2, showing the time waveforms of the satellite roll attitude φ, the signal generator output, and the north-south direction pointing error e _NS obtained by this embodiment. As is clear from comparing FIG. 3 and FIG. 7, according to this embodiment, even if the orbital inclination angle is not zero (excluding minute terms of second order or higher), e _NS =0... …(5)

In the above embodiment, among the constants used in the signal generator, the amplitude a and the initial phase angle b are input as commands from the ground, and the orbital angular velocity ω ₀ is an internal constant, but a, b, ω ₀ It is also possible to use all or any part of them as command input, or to omit command input altogether and set everything internally. Furthermore, instead of a and b, i or t ₀ may be used according to equation (3) or equation (2).
The values of a and b actually change over time, albeit more slowly than the orbital period, so they need to be updated from time to time. However, if the temporal changes in both or one of a and b are modeled and generated on-board, the frequency of commands from the ground can be made lower than in the above embodiment.

Furthermore, in the above embodiment, a method was assumed in which the yaw attitude is passively controlled by providing bias angular momentum around the Y axis (pitch axis), but the yaw attitude is estimated or detected by a separately provided yaw sensor, and the yaw The present invention is also applicable to systems that are actively controlled by system controllers and actuators (three-degree-of-freedom controlled bias momentum system or zero momentum system). In the latter case, the satellite's Y-axis attitude does not remain constant with respect to inertial space, but changes in a sinusoidal manner. The relationship shown holds true as is.

〔Effect of the invention〕

As described above, according to the present invention, in an earth-oriented artificial satellite in a geostationary orbit, even if the orbital plane is tilted from the equator plane, the pointing error in the north-south direction that occurs in the conventional attitude control device that uses an earth sensor can be solved. It has the effect of being able to suppress Therefore, there is also the effect that the permissible range of the orbital inclination angle can be widened without deteriorating the pointing error in the north-south direction.

[Brief explanation of the drawing]

FIG. 1 is a block diagram showing an embodiment of the attitude control device (roll system) according to the present invention, and FIG.
An explanatory diagram of the attitude of the satellite realized by the embodiment shown in the figure, FIG. 3 is an explanatory diagram showing the effect of the above embodiment,
Figure 4 is a block diagram of a conventional attitude control system (pitch system and roll system) using an earth sensor, Figure 5 is a diagram for explaining the definition of the satellite attitude, and Figure 6 is a conventional attitude control system. FIG. 7 is an explanatory diagram of the attitude of the satellite realized by the method, and FIG. 7 is a diagram illustrating the problems of the conventional attitude control device. In the figure, 1 is the target attitude (pitch angle θ or roll angle φ = 0), 2 is the satellite attitude (pitch angle θ or roll angle φ), 3 is the earth sensor, 4 is the earth sensor output (pitch or roll), 5 is a controller,
6 is an actuator, 7 is an attitude dynamics of the satellite, 8 is a command input from the ground, 9 is a signal generator, 10 is a sinusoidal signal, 11 is a sum signal of 10 and 4, 12 is an equatorial plane, 13 is an orbital plane, 14 is the earth, 15 is the artificial satellite, 16 is the vertical direction of the orbital plane, 1
7 is the Y axis (pitch axis) of the satellite, 18 is 16 and 17
19 is the pointing direction of the satellite-mounted mission equipment, 20 is the latitude to which the target pointing point on the earth's surface belongs, 21 is the direction of the center of the earth as seen from the satellite, and 22 is 1
6 and 21, 23 is the Y axis (pitch axis) of the satellite, 24 is the X axis (roll axis) of the satellite, 25 is the Z axis (yaw axis) of the satellite, 26 is the pitch angle θ, 27
is the roll angle φ, 28 is the yaw angle ψ, 29 is the orbital inclination angle, 30 is the ascending node direction, and 31 is the mission equipment onboard the satellite. Note that the same reference numerals in the figures indicate the same or corresponding parts.

Claims (1)

[Claims] 1 In a device that detects the attitude of an artificial satellite using an earth sensor installed on the artificial satellite and operates an actuator based on the detection result to control the attitude of the artificial satellite with respect to the earth, a sinusoidal signal that changes with the orbit period. A signal generator is provided that generates a When trying to maintain the roll attitude at zero by directly using the roll output of the sensor as the reference input for the attitude control system, the north-south direction changes sinusoidally with the orbit period caused by a non-zero orbital inclination angle. By adjusting the amplitude and opposite phase to the pointing fluctuation of the satellite, the roll attitude of the satellite is changed in a sinusoidal manner so that the pointing fluctuation in the north-south direction does not occur even when the orbital inclination angle is not zero. An artificial satellite attitude control device characterized by: