JP2010188915A - Attitude control device of artificial satellite - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide an attitude control device of an artificial satellite for changing the attitude of the artificial satellite with high accuracy without being affected by a specific point of a CMG, and promptly setting it to the target attitude after changing the attitude. <P>SOLUTION: In the device changing the attitude of the artificial satellite 1 with a plurality of the CMGs 2, a CMG steering calculating part 7 calculating the target gimbal angle and the target gimbal angular velocity of the CMG 2 superposes the correction values of the gimbal angle and the gimbal angular velocity obtained from feed-back attitude control torque. The correction values are made as the first correction values of the gimbal angle and the gimbal angular velocity attaining the feed-back attitude control torque in a range apart from a specific point in which the CMG 2 cannot output the attitude control torque in a specific direction, and as the second correction values of the gimbal angle and the gimbal angular velocity setting only the element in the specific direction of the feed-back attitude control torque to zero in the range adjacent to the specific point. <P>COPYRIGHT: (C)2010,JPO&INPIT

Description

  The present invention relates to an attitude control device for an artificial satellite that changes the attitude of an artificial satellite and sets it to a target attitude.

High-motion artificial satellites are equipped with multiple control moment gyros (abbreviated as CMG) as actuators, and in order to control the three-axis attitude of artificial satellites using this CMG, multiple CMGs are used. It is necessary to perform the appropriate coordinated operation so that the attitude control torque required by the sum of the torques generated by the CMGs is obtained. Normally, each CMG has the same angular momentum, and the three-axis attitude control by the zero momentum system is performed with the point where the total angular momentum becomes zero as the operation center.
The relationship between the CMG gimbal angular velocity and the attitude control torque acting on the artificial satellite is expressed by the following equation (1) using the Jacobian matrix.

  Here, θ is a vector composed of gimbal angles of each CMG (N-dimensional vector, N is the number of CMGs), θ is a gimbal angular velocity vector (N-dimensional vector), and τ is attitude control torque (three-dimensional) acting on the artificial satellite. Vector). Each column of the Jacobian matrix J (θ) (3 × N-dimensional matrix) indicates a torque obtained when the corresponding CMG gimbal is rotated at an angular velocity of 1 rad / s. Hereinafter, J (θ) is described as J for simplicity. When the arrangement of CMGs on the artificial satellite is determined, the Jacobian matrix J can be obtained as a function of the gimbal angle of each CMG.

In a conventional satellite attitude control device, when the attitude control torque τ necessary for attitude change is given, this is multiplied by the pseudo inverse matrix of the Jacobian matrix J to obtain the gimbal angular velocity of each CMG, and the obtained gimbal angular velocity is calculated. By controlling the gimbal of each CMG as a target value with a speed servo system, posture control torque necessary for posture change is generated and posture control is performed.
The target value of the CMG gimbal angular velocity for realizing the attitude control torque τ is obtained by the following equation (2) using a pseudo inverse matrix of the Jacobian matrix J.

Here, J ^T is a transposed matrix of the matrix J, and (JJ ^T ) ⁻¹ is an inverse matrix of the matrix JJ ^T. In this case, depending on a specific combination of CMG gimbal angles, there is a problem that the 3 × 3D matrix JJ ^T in equation (2) is not regular, and there is a state where the inverse matrix cannot be calculated. Further, as the CMG gimbal angle combination approaches this specific CMG gimbal angle combination, the calculated gimbal angular velocity becomes infinite, and there is a problem that the posture cannot be actually controlled.
This combination of specific CMG gimbal angles is generally called a CMG singular point, and a torque direction that cannot be output is called a singular direction. At the singular point, since the posture control torque in the singular direction cannot be generated, arbitrary three-axis posture control cannot be performed.

  As a measure for avoiding such a singular point of CMG, in the satellite attitude control device described in Patent Document 1, the gimbal angular velocity is obtained using the following equations (3) to (5).

Here, kP ^{−1 in the} equation (3) is a minute term added to avoid a singular point. This prevents the inverse matrix from being unable to be calculated from the values of the diagonal components of the matrix P, and provides a time-varying non-diagonal component to prevent the attitude control of the artificial satellite from stopping at a singular point.

JP 2002-506773 A

However, in the satellite attitude control device as described above, a minute term is added to avoid the singular point of CMG. There was a problem that it always had an error, and high-precision attitude control was not possible.

  The present invention has been made to solve the above-described problems, and can change the posture with high accuracy without being affected by the singular point of the CMG, and can quickly settle to the target posture after the posture change. The purpose is to obtain an attitude control device for an artificial satellite.

  An attitude control device for an artificial satellite according to the present invention is mounted on an artificial satellite having a plurality of CMGs for controlling the attitude, and an attitude estimation unit that estimates the attitude angle and attitude angular velocity of the artificial satellite; Plans a target trajectory consisting of a gimbal estimation unit for estimating the gimbal angular velocity, a target attitude angle and target attitude angular velocity for changing the attitude of the satellite, and a CMG gimbal angle plan value, gimbal angular velocity plan value, and gimbal angular acceleration plan value Attitude control feedback to obtain feedback attitude control torque to reduce the error between the target trajectory planning section and the target attitude angle and target attitude angular velocity from the target trajectory planning section and the estimated attitude angle and attitude angular velocity from the attitude estimation section The calculation unit, the feedback attitude control torque and the eye from the attitude control feedback calculation unit From the gimbal angle plan value, the gimbal angular velocity plan value and the gimbal angular acceleration plan value from the trajectory planning unit, the gimbal angle estimation value from the gimbal estimation unit, and the posture angular velocity estimation value from the posture estimation unit, the target gimbal angle of the CMG GMG An artificial satellite attitude control device including a gimbal control calculation unit that obtains a gimbal control torque so that an estimated value of a gimbal angle and an estimated value of a gimbal angular velocity from an estimation unit follow each other. Planned gimbal angle and gin The gimbal angle and gimbal angular velocity correction value obtained from the feedback attitude control torque are superimposed on the plan angle speed value, and the above correction value is fed back in a region away from the singular point where the CMG cannot output the attitude control torque in a specific direction. The first correction value of the gimbal angle and the gimbal angular velocity for realizing the control torque, and the second correction value of the gimbal angle and the gimbal angular velocity in which only the component in the specific direction of the feedback attitude control torque is zero in the region near the singular point It is characterized by doing.

  In the present invention, when the attitude control of the artificial satellite is performed using a plurality of CMGs, the gimbal angle and gimbal angular velocity obtained from the feedback attitude control torque to the gimbal angle planned value and the gimbal angular velocity planned value from the target trajectory planning unit are determined. The correction value is superimposed, and the correction value is set as a first correction value for the gimbal angle and the gimbal angular velocity that realizes the feedback posture control torque in a region away from the singular point where the CMG cannot output the posture control torque in a specific direction. Since the second correction value in which only the component in the specific direction of the feedback attitude control torque is set to zero in the region near the singular point, the attitude control can be performed with high accuracy without being affected by the singular point of the CMG. .

It is a block diagram which shows the structure of the attitude | position control apparatus of the artificial satellite by Embodiment 1 of this invention. FIG. 3 is a diagram showing an example of orbit planning for CMG 2 and artificial satellite 1 at the time of attitude change in order to explain the operation of FIG. 1. 3 is a block diagram illustrating a configuration of a CMG steering calculation unit 7 according to Embodiment 1. FIG. It is a block diagram which shows the structure of the attitude | position control apparatus of the artificial satellite by Embodiment 2 of this invention. 6 is a block diagram showing a configuration of a CMG steering calculation unit 7 according to Embodiment 2. FIG. It is a block diagram which shows the structure of the attitude | position control apparatus of the artificial satellite by Embodiment 3 of this invention. FIG. 10 is a block diagram illustrating a configuration of a CMG steering calculation unit 7 according to a third embodiment. The relationship between the x and the function g according to the fourth embodiment, ε = α ₃ in the case of 0.2 and functions _{g / max (α 3 g,} ε) is a diagram showing the relationship.

Embodiment 1 FIG.
FIG. 1 is a block diagram showing a configuration of an attitude control device for an artificial satellite according to Embodiment 1 of the present invention. The artificial satellite 1 is equipped with a plurality of CMGs as attitude control torque actuators. In the following description, a part for controlling one of the CMGs will be described. Of course, the same embodiment is applied to other CMGs.
In the block diagram of FIG. 1, the attitude of the CMG 2, the gimbal estimation unit 3 for estimating the gimbal angle and the gimbal angular velocity of the CMG 2, the attitude estimation unit 4 for estimating the attitude angle and the attitude angular velocity of the artificial satellite 1, and the attitude of the artificial satellite 1 are changed. Target posture angle and target posture angular velocity for generating a target, and a target trajectory plan unit for generating a gimbal angle plan value, a gimbal angular velocity plan value, and a gimbal angular acceleration plan value of CMG2 that dynamically match these target posture angle and target posture angular velocity 5 and attitude control feedback calculation for obtaining feedback attitude control torque to reduce the error between the target attitude angle and target attitude angular velocity from the target trajectory planning unit 5 and the estimated values of attitude angle and attitude angular velocity from the attitude estimation unit 4 Part 6 is provided.

The feedback attitude control torque from the attitude control feedback calculation unit 6, the gimbal angle plan value, the gimbal angular velocity plan value and the gimbal angle acceleration plan value from the target trajectory plan unit 5, and the gimbal angle of the CMG 2 from the gimbal estimation unit 3 From the estimated value and the estimated value of the attitude angular velocity of the artificial satellite 1 from the attitude estimation unit 4, the feedforward gimbal control that compensates for the CMG2 target gimbal angle and target gimbal angular velocity, and the inertia force and gyro term of the CMG2 associated with the attitude change. In order to reduce errors between the CMG steering calculation unit 7 that calculates torque, and the target gimbal angle and target gimbal angular velocity of the CMG 2 from the CMG steering calculation unit 7 and the estimated values of the gimbal angle and gimbal angular velocity from the gimbal estimation unit 3 Gin for gimbal control torque And a le control arithmetic unit 8.

The gimbal estimation unit 3 can estimate the gimbal angle as an angle sensor output by using, for example, a CMG gimbal shaft with a general angle sensor (such as an encoder or resolver). Further, the gimbal angular velocity can be estimated by differentiating the gimbal angle with respect to time.
The attitude estimation unit 4 includes, for example, an attitude sensor (such as a start tracker or an earth sensor) and an attitude angular velocity sensor (gyroscope) mounted on the artificial satellite 1 and constructs a Kalman filter using these sensor outputs, thereby constructing the artificial satellite 1. It is possible to estimate the posture angle and posture angular velocity.

When the artificial satellite 1 changes its attitude, the target trajectory planning unit 5 performs a trajectory plan for changing the attitude. Here, in addition to the target attitude angle and target attitude angular velocity of the artificial satellite 1, the CMG2 gimbal angle plan value, gimbal angular velocity plan value, and gimbal angular acceleration plan value that dynamically match them under the law of conservation of angular motion are also included. Generate.
When the initial (current) attitude of the artificial satellite 1 and the terminal attitude that is the target of attitude change are given, the attitude change from the initial attitude to the terminal attitude is, for example, as shown in Non-Patent Document 1, Euler axis ( It can be realized by rotating around one axis called Euler Axis.
When this rotation axis and u _a, angular momentum direction u _c to be generated CMG2 for attitude change can be determined by the following equation (6).
P.C. Hughes, “Spacecraft Attitude Dynamics”, John Wiley & Sons, 1986, page 11, Figure 2.3.

Here, I is the moment of inertia of the artificial satellite 1 and ‖Iu _a ‖ is the norm of the vector Iu _a . Since the vector Iu _a is not _a zero vector, ‖Iu _a ‖ never becomes zero.
In order to quickly change the attitude with high accuracy, it is necessary to cause the artificial satellite 1 to generate a large attitude angular velocity by the angular momentum near the maximum value of the angular momentum that can be generated by the CMG 2. If the angular momentum direction u _c required position change is obtained, and the maximum angular momentum that direction CMG2 can be generated, the gimbal angle of CMG2 at that time may be determined by the following procedure.
A set of maximum angular momentum that can be generated by the CMG 2 is generally called a maximum angular momentum envelope, and forms a closed curved surface in the angular momentum space. Since the maximum angular momentum envelope surface can be obtained by the method described in Chapter 3 of Non-Patent Document 2, for example, detailed description thereof is omitted.

Next, when the angular momentum is given, the gimbal angle of CMG2 corresponding to the angular momentum can be obtained by, for example, the method described in Non-Patent Document 3 (repetitive calculation using the Newton-Raphson method). Is omitted.
Wie, B., “Singularity Analysis and Visualization for Single-Gimbal Control Moment Gyro Systems,” Journal of Guidance, Control, and Dynamics, Vol. 27, No. 2, 2004, pp.271-282. Kurokawa, “Control Law of 3-Unit CMG for Small Satellites”, Bulletin of Mechanical Engineering Laboratory, Vol. 53, No. 6, 1999, pp. 203-209.

When the gimbal angle of CMG2 near the maximum angular momentum is obtained by the above-described procedure, the trajectory of the CMG2 gimbal angle in the posture change section typically has a trapezoidal pattern as shown in FIG. In other words, the CMG2 is driven from the initial state zero (the CMG generation angular momentum is zero) to the point A in FIG. 2A to generate the maximum angular momentum in the CMG2, thereby causing the artificial satellite 1 to rotate around the Euler axis. The maximum attitude angular velocity is generated.
After the maximum attitude angular velocity is generated, the gimbal is held and the generated angular momentum of CMG 2 is maintained, and the artificial satellite 1 continues to generate the maximum attitude angular velocity.
Near the end of the attitude change section, the CMG2 gimbal is driven again to return to zero, the angular momentum generated by CMG2 is released, and the attitude change to the target attitude is completed with the attitude angular velocity of the artificial satellite 1 set to zero. FIG. 2 shows an example of the history of the CMG 2 gimbal angle and generated angular momentum and the attitude angular velocity and attitude angle of the artificial satellite 1 during the attitude change section.

The trajectory of the CMG2 gimbal angle in FIG. 2A is called a gimbal angle planned value. The gimbal angular velocity planned value can be obtained by time differentiating the gimbal angular planned value, and the gimbal angular acceleration planned value can be obtained by time differentiating the gimbal angular velocity planned value.
If the gimbal angle plan value of CMG2 is obtained, the attitude orbit of the artificial satellite 1 can be obtained from the angular momentum conservation law. The angular momentum of the entire satellite is given by the sum of the angular momentum of the artificial satellite 1 and the generated angular momentum of the CMG 2. As shown in the following equation (7), the angular momentum of the entire satellite is preserved in the inertial space by the angular momentum conservation law.

Here, ω _b is the attitude angular velocity of the artificial satellite 1, and h _cmg is the generated angular momentum of the CMG 2. Since h _cmg can be obtained from the gimbal angle planned value of CMG 2 obtained previously, the attitude angular velocity of the artificial satellite 1 that is dynamically matched can be obtained using equation (7). This is the target posture angular velocity. The target posture angle can be obtained by time integration of the target posture angular velocity.
Posture control is performed using an error (posture angle error and posture angular velocity error) between the target posture angle and target posture angular velocity from the target trajectory planning unit 5 and the estimated values of the posture angle and posture angular velocity obtained from the posture estimation unit 4. The feedback calculation unit 6 obtains a feedback attitude control torque so as to reduce this error.
If the posture angle error is e _θ and the posture angular velocity error is e _ω , the feedback posture control torque τ _FB can be obtained by the following equation (8) using PD control, for example.

Here, K _P and K _D are feedback gains for the attitude angle error and the attitude angular velocity error, respectively. Alternatively, feedback attitude control torque τ _FB may be obtained by PID control by adding integral compensation instead of PD control.
Next, the CMG steering calculation unit 7 uses the feedback attitude control torque τ _FB to obtain target values for the gimbal angle and gimbal angular velocity of the CMG 2 necessary for attitude change. When the Jacobian matrix J given as a function of the gimbal angle of CMG2 is subjected to singular value decomposition, it is expressed by the following equations (9) to (13).

  Here, U is a 3 × 3 dimensional orthogonal matrix, V is an N × N dimensional orthogonal matrix, and S is a 3 × N dimensional matrix. By definition, the pseudo inverse matrix of J is expressed by the following equation (14).

Since σ ₃ → 0 in the vicinity of the singular point in the equation (14), the third term of the equation diverges, and the correction value θ _correct for the gimbal angular velocity obtained using the pseudo inverse matrix according to the following equation (15): Will also diverge near the singularity.

Therefore, the matrix J ⁺ is defined using the following equations (16) to (18), where ε is a minute positive number.

Here, max (A, B) is a function that outputs the maximum value of A and B. Using this matrix J ⁺ , the correction value θ _correct for the gimbal angular velocity is calculated using the following equation (19).

That is, the correction value of the gimbal angular velocity of the CMG 2 that realizes the feedback attitude control torque is obtained by multiplying the feedback attitude control torque τ _FB obtained by the attitude control feedback calculation unit 6 by the matrix J ⁺ expressed by the equation (16). In the region away from the singular point, J ^{+ in} equation (16) is equal to the right side of equation (14). Therefore, the correction value of the gimbal angular velocity is calculated only from the component that realizes the feedback attitude control torque, and does not include an extra error for avoiding the singular point by the CMG2. Further, in the region near the singularity (16) while the alpha ₃ → 0 becomes, J ⁺ third term is zero in formula, J ⁺ first term of the second term remains as a finite value. Therefore, the correction value of the gimbal angular velocity calculated using the equation (19) is a value obtained by setting only the component in the singular direction of the feedback attitude control torque as zero. Thereby, the divergence of the correction value of the gimbal angular velocity is suppressed in the region near the singular point.
Further, from the equations (16) to (18), the correction value of the gimbal angular velocity calculated using the equation (19) continuously changes including the singular point vicinity region.

FIG. 3 shows a detailed configuration of the CMG steering calculation unit 7. The target gimbal angle of the CMG 2 is obtained by superimposing the time-integrated gimbal angular velocity correction value obtained above on the gimbal angle planned value generated by the target trajectory planning unit 5. Similarly, the gimbal angular velocity is superimposed on the gimbal angular velocity plan value generated by the target trajectory planning unit 5 to obtain the target gimbal angular velocity of CMG2.
In addition, the feedforward compensation calculation unit 9 obtains the sum of the gyro term based on the inertial force of the CMG 2 based on the gimbal angular acceleration planned value generated by the target trajectory planning unit 5 and the attitude angular velocity of the artificial satellite 1 and the angular momentum of the CMG 2. Feedforward gimbal control torque.

  Next, the control system is configured so that the gimbal of the CMG 2 follows the trajectory at the target gimbal angle and the target gimbal angular velocity obtained by the CMG steering calculation unit 7 in the gimbal control calculation unit 8. FIG. 1 shows an example of the configuration. An error (gimbal angle error) between the gimbal angle of the CMG 2 estimated by the gimbal estimation unit 3 and the target gimbal angle obtained by the CMG steering calculation unit 7 is obtained, and an angle control system is configured using the gimbal angle error. As the angle control system, for example, a proportional control that multiplies a gimbal angle error by a proportional gain, or a combination of proportional control and integral control (PI control) that multiplies a gimbal angle error by a proportional gain and an integral gain is used. But you can.

Further, an error (gimbal angular velocity error) between the gimbal angular velocity of CMG 2 estimated by the gimbal estimating unit 3 and the target gimbal angular velocity obtained by the CMG steering calculating unit 7 is obtained, and an angular velocity control system is configured using the gimbal angular velocity error. . As the angular velocity control system, for example, a system based on proportional control in which a gimbal angular velocity error is multiplied by a proportional gain can be used.
The sum of the output torque from the angle control system and the angular velocity control system described above and the feedforward gimbal control torque obtained by the CMG steering calculation unit 7 is obtained and used as the gimbal control torque to perform the gimbal control of the CMG2.

Since the correction values for the gimbal angle and the gimbal angular velocity calculated by the CMG steering calculation unit 7 in this way do not include an extra error for avoiding the singular point by the CMG 2 in a region away from the singular point, 3 axis feedback attitude control becomes possible.
Further, in the region near the singular point, at least partial feedback attitude control excluding the singular direction is possible, and the effect of feedback can be utilized to the maximum.

According to this embodiment, the correction value of the gimbal angle and the gimbal angular velocity superimposed on the gimbal angular velocity planned value and the gimbal angular velocity planned value from the target trajectory planning unit 5 in the CMG steering calculation unit 7 realizes the feedback attitude control torque. In the region near the singular point, only the component in the singular direction of the feedback attitude control torque can be set to zero.
Thereby, the attitude of the artificial satellite can be changed with high accuracy without being affected by the singular point by using the high output torque of the CMG 2, and the target attitude can be quickly settled after the attitude change.
Further, since the correction values for the gimbal angle and the gimbal angular velocity are continuous values even when passing through the region near the singular point, continuous feedback attitude control is realized.
As a result, the attitude of the artificial satellite can be changed with high accuracy, and after the attitude change, the target attitude can be quickly settled.

Embodiment 2. FIG.
Next, with reference to FIGS. 4 and 5, the configuration of the attitude control device for the artificial satellite according to the second embodiment of the present invention will be described. In the first embodiment, the attitude control device in the case of changing the attitude of the artificial satellite has been described, but in the present embodiment, a case in which the attitude of the artificial satellite is maintained in a certain steady attitude will be described.

FIG. 4 is a block diagram showing the configuration of the attitude control device for an artificial satellite according to the present embodiment. When the attitude of the artificial satellite 1 is maintained at a certain steady attitude, the target trajectory planning unit 5 does not need to generate the planned values of the CMG 2 gimbal angle, gimbal angular velocity, and gimbal angular acceleration in the target trajectory. This is equivalent to setting the gimbal angle planned value, the gimbal angular velocity planned value, and the gimbal angular acceleration planned value from the target trajectory planning unit 5 in Embodiment 1 to zero, and the processing of the target trajectory planning unit 5 is simplified. Can do. Accordingly, the CMG steering calculation unit 7 can be configured as shown in FIG.
The CMG steering calculation unit 7 obtains the correction values for the gimbal angle and the gimbal angular velocity in the same procedure as in the first embodiment, and superimposes them on the gimbal angle planned value and the gimbal angular velocity planned value set to zero as described above. The gimbal angle and the target gimbal angular velocity are used. The feedforward compensation calculation unit 9 obtains a gyro term based on the attitude angular velocity of the artificial satellite 1 by the attitude estimation unit 4 and the angular momentum of the CMG 2 to obtain a feedforward gimbal control torque.
The gimbal control calculation unit 8 uses this feedforward gimbal control torque to configure a control system so that the gimbal of the CMG 2 follows the target gimbal angle and the target gimbal angular velocity.
As described above, when maintaining the attitude of the artificial satellite in a certain steady attitude, the attitude control device of the artificial satellite 1 can be simplified as compared with FIG. 1 as shown in FIG.

  According to the present embodiment, in the CMG steering calculation unit 7, the gimbal angle and gimbal angular velocity correction values superimposed on the planned gimbal angle value and the gimbal angular velocity plan value from the target trajectory plan unit 5 are the same as those in the first embodiment. Since the same procedure is used, the same effect as in the first embodiment can be obtained.

Embodiment 3 FIG.
In the second embodiment, when the control band of the gimbal control calculation unit 8 can be secured sufficiently high, the satellite attitude control device can be configured as follows.
FIG. 6 is a block diagram showing the configuration of the attitude control device for an artificial satellite according to the present embodiment. Similar to the second embodiment, the target trajectory planning unit 5 can set the planned gimbal angle value, the gimbal angular velocity planned value, and the gimbal angular acceleration planned value of the CMG 2 to zero, thereby simplifying the process. Further, when the control band of the gimbal control calculation unit 8 can be secured sufficiently high, the feedforward gimbal control torque from the CMG steering calculation unit 7 can be omitted. Therefore, the CMG steering calculation unit 7 in the present embodiment can be configured as shown in FIG. 7, and can be further simplified than FIG.

The CMG steering calculation unit 7 obtains correction values for the gimbal angle and the gimbal angular velocity in the same procedure as in the first embodiment, and superimposes it on the gimbal angle planned value and the gimbal angular velocity planned value that are set to zero as described above. And the target gimbal angular velocity as an input to the gimbal control calculation unit 8. In the gimbal control calculation unit 8, the control system is configured so that the gimbal of the CMG 2 follows the target gimbal angle and the target gimbal angular velocity.
As described above, when the attitude of the artificial satellite 1 is maintained at a certain steady attitude, when the control band of the gimbal control calculation unit 8 can be secured sufficiently high, the attitude control device of the artificial satellite 1 is as shown in FIG. 4 can be further simplified.

  According to the present embodiment, in the CMG steering calculation unit 7, the gimbal angle and gimbal angular velocity correction values superimposed on the planned gimbal angle value and the gimbal angular velocity plan value from the target trajectory plan unit 5 are the same as those in the first embodiment. Since the same procedure is used, the same effect as in the first embodiment can be obtained.

Embodiment 4 FIG.
If the excitation of the vibration mode specific to the flexible structure accompanying the artificial satellite can be suppressed when changing the attitude of the artificial satellite, the attitude can be changed with higher accuracy. In order to suppress excitation of this vibration mode, it is necessary to suppress jerk of the target orbit of the attitude angle of the artificial satellite. In the case of using CMG, this corresponds to suppressing the acceleration of the target trajectory of the gimbal angle, that is, the gimbal angular acceleration. For this purpose, it is necessary to generate a trajectory in which the gimbal angular acceleration is continuous in the vicinity of the singular point.
In the present embodiment, a configuration capable of generating a trajectory in which the gimbal angular acceleration is continuous will be described.

In the first embodiment, an example of the configuration of the matrix J ⁺ has been described, but the configuration is not limited thereto. ε may be a minute positive number, and the matrix J ⁺ may be configured as in the following equation (20).

Here, as the function g, for example, configurations of the following formulas (21) to (23) are conceivable.

FIG. 8A shows changes in the functions g ₁ , g ₂ , and g ₃ depending on x. FIG. 8 (b) shows the change due to α ₃ of the coefficient g / max (α ₃ g, ε) of the third term of the matrix J ⁺ when ε = 0.2.
When g ₂ or g ₃ is used as the function g, the coefficient g / max (α ₃ g of the third term of J ⁺ in the singular point vicinity region (α ₃ = ε) and the singular point (α ₃ = 0). , Ε) is continuous and differentiable. For this reason, the correction value of the gimbal angular velocity obtained by multiplying by J ⁺ is also continuous and differentiable, and the resulting gimbal angular acceleration is continuous.
In the above example, the function g ₁ is given for comparison. In the case of using g ₁ as a function g is different from the case of using the g _2, g _3, the singularity neighborhood area and singularity coefficients J ⁺ third term of g / max (alpha ₃ g, epsilon ) Is continuous but not differentiable. For this reason, the correction value of the gimbal angular velocity obtained by multiplying by J ⁺ cannot be differentiated, and the resulting gimbal angular acceleration is discontinuous.

As described above, the gimbal angular acceleration obtained using the function g ₂ or g ₃ is continuous even when passing through the region near the singular point. Therefore, a trajectory with continuous gimbal angular acceleration can be generated even in the region near the singular point, and the acceleration of the target trajectory with the gimbal angle can be suppressed. As a result, the jerk of the attitude angle of the artificial satellite can be suppressed, so that excitation of the vibration mode of the flexible structure associated with the artificial satellite can be suppressed.
As described above, the attitude of the artificial satellite can be changed with higher accuracy even when passing through the singular point vicinity region or the singular point region, and can be quickly set to the target posture after the posture change.

According to the present embodiment, in the CMG steering calculation unit 7, the gimbal angle and gimbal angular velocity correction values superimposed on the planned gimbal angle value and the gimbal angular velocity plan value from the target trajectory plan unit 5 are the same as those in the first embodiment. Since the same procedure is used, the same effect as in the first embodiment can be obtained.
Further, by using the correction values for the gimbal angle and the gimbal angular velocity, it is possible to suppress the jerk of the artificial satellite when passing through the region near the singular point or passing through the singular point, and the vibration of the flexible structure characteristic accompanying the artificial satellite. Mode excitation can be suppressed.
As a result, the attitude of the artificial satellite can be changed with higher accuracy even in the vicinity of the singularity area or passing through the singularity area, and can be quickly set to the target attitude after the attitude change.

  1 artificial satellite, 2 CMG, 3 gimbal estimation unit, 4 attitude estimation unit, 5 target trajectory planning unit, 6 attitude control feedback calculation unit, 7 CMG steering calculation unit, 8 gimbal control calculation unit, 9 feedforward compensation calculation unit.

Claims (3)

It is mounted on an artificial satellite having multiple CMGs for controlling the attitude,
An attitude estimation unit for estimating the attitude angle and attitude angular velocity of the artificial satellite;
A gimbal estimation unit for estimating the gimbal angle and gimbal angular velocity of the CMG;
A target trajectory planning unit for planning a target trajectory composed of a target attitude angle and a target attitude angular velocity for changing the attitude of the artificial satellite, and a gimbal angle plan value, a gimbal angular velocity plan value and a gimbal angular acceleration plan value of the CMG;
Attitude control feedback calculation for obtaining feedback attitude control torque to reduce errors between the target attitude angle and the target attitude angular velocity from the target trajectory planning unit and the estimated values of the attitude angle and attitude angular velocity from the attitude estimation unit And
The feedback attitude control torque from the attitude control feedback calculation unit, the gimbal angle plan value from the target trajectory plan unit, the gimbal angular velocity plan value and the gimbal angular acceleration plan value, and the gimbal angle from the gimbal estimation unit. From the estimated value and the estimated value of the attitude angular velocity from the attitude estimation unit, a feedforward gimbal control torque that compensates for the CMG inertial force and gyro term due to the CMG target gimbal angle and target gimbal angular velocity and attitude change. A CMG steering computing unit for computing;
A gimbal control calculation for obtaining a gimbal control torque so that the estimated value of the gimbal angle and the estimated value of the gimbal angular speed from the gimbal estimating unit follow the target gimbal angle and the target gimbal angular velocity from the CMG steering calculating unit. In an attitude control device for an artificial satellite equipped with a
The CMG steering computing unit superimposes the gimbal angle and gimbal angular velocity correction values obtained from the feedback attitude control torque on the gimbal angle planned value and the gimbal angular velocity planned value from the target trajectory planning unit,
The above correction value
In a region away from a singular point where the CMG cannot output attitude control torque in a specific direction, the first correction value of the gimbal angle and gimbal angular velocity for realizing the feedback attitude control torque is used.
The artificial satellite attitude control device characterized in that in the region near the singular point, the gimbal angle and the second correction value of the gimbal angular velocity are set such that only the component in the specific direction of the feedback attitude control torque is zero. 2. The attitude control device for an artificial satellite according to claim 1, wherein the correction values of the gimbal angle and the gimbal angular velocity change continuously when passing through the region near the singular point. 3. The attitude control device for an artificial satellite according to claim 2, wherein the correction values of the gimbal angle and the gimbal angular velocity continuously change including the higher-order differential value when passing through the region near the singular point.

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