JP7329454B2 - displacement compensation controller - Google Patents

Description

The present invention relates to attitude control of a spacecraft.

An artificial object that flies in space is called a "spacecraft". For example, a satellite is a spacecraft.
Three or more control moment gyros (CMGs) may be used to control the spacecraft's three-axis attitude.
A CMG has one or more gimbal axes and a wheel. The wheels rotate about axes orthogonal to each gimbal axis. The axis of rotation of the wheel is called the "spin axis".
A CMG is an actuator that changes the direction of angular momentum by rotating a wheel that rotates about its spin axis about its gimbal axis.
A rotation angle about the gimbal axis is called a "gimbal angle" or a "gimbal rotation angle."

To adjust the attitude of the spacecraft, the gimbal angle is changed and torque is output. The direction of the output torque is orthogonal to the direction around the gimbal axis.

An interface for installing the CMG in the spacecraft is called an "installation interface".
If the mounting interface for the CMG is not rigid, the output torque will cause displacement of the mounting interface. Any displacement of the mounting interface changes the direction of the angular momentum of the CMG and the direction of the output torque. These changes affect the attitude of the spacecraft.

High-frequency micro-vibrations are generated by the rotation of the CMG wheel. This high-frequency minute vibration is called "disturbance".
Disturbances affect the attitude of the spacecraft. Disturbances also affect the accuracy of each device mounted on the spacecraft. For example, disturbances affect the observation accuracy of observation instruments mounted on spacecraft.
In order to reduce the effects of disturbances, materials with vibration isolation effects are sometimes used in the installation interface. For example, materials such as spring dampers or anti-vibration rubber are used for the installation interface.
A material that has a vibration isolation effect is called an "isolator".
Multiple isolators that are used simultaneously are referred to as a "set of isolators".

A set of isolators is used for multiple CMGs or for a single CMG.
When one set of isolators is used for multiple CMGs, the displacement of the set of isolators is affected by the combined torque output from the multiple CMGs. A composite torque is a composite result of torques output from each of a plurality of CMGs.
When one set of isolators is used for one CMG, the displacement of the set of isolators is affected by the torque output from one CMG. Therefore, the displacement of a set of isolators is different for each CMG.

Patent Literature 1 describes an attitude determination technique that considers the influence of disturbances.
In order to compensate for the influence of the disturbance after attitude determination is completed, an attitude control system that handles frequency components in the same band as the frequency components of the disturbance is required. In general, attitude fluctuations cannot be suppressed when the attitude control system cannot handle the band to be compensated for, not just when compensating for the effects of high-frequency disturbances.

The displacement of the installation interface includes changes in the frequency of the installation interface and changes in the amplitude of the installation interface. The frequency of the installation interface and the amplitude of the installation interface vary depending on the stiffness of the installation interface.
In order to ensure the stability of the attitude control system, the frequency of the installation interface is often higher than the bandwidth of the attitude control system. Therefore, in many cases, the frequency of the installation interface cannot be handled by the attitude control system.

The use of a set of isolators as the installation interface suppresses the transmission of high frequency disturbances, while the displacement of the set of isolators affects the attitude of the spacecraft.

Japanese Patent No. 4509006

Conventionally, the use of a set of isolators as an installation interface suppresses the transmission of high-frequency disturbances, while the displacement of the set of isolators affects the attitude of the spacecraft.

SUMMARY OF THE INVENTION It is an object of the present invention to suppress the influence of displacement of an installation interface on the attitude of a spacecraft.

The displacement correction controller of the present invention is used in an attitude control system that controls the attitude of a spacecraft.
The spacecraft comprises the attitude control system, a plurality of control moment gyros, and a plurality of installation interfaces for installing the plurality of control moment gyros in the spacecraft.
The displacement correction controller comprises:
a displacement estimator that estimates a displacement of each of the plurality of installation interfaces;
and a torque corrector for correcting the total torque of the plurality of control moment gyros calculated for feedback control by the attitude control system based on the estimated displacement of each of the plurality of installation interfaces.

SUMMARY OF THE INVENTION An object of the present invention is to suppress the influence of the displacement of the installation interface on the attitude of the spacecraft.

1 is a configuration diagram of a spacecraft 100 according to Embodiment 1. FIG. FIG. 2 is a configuration diagram of a CMG 110i according to the first embodiment; FIG. 2 is a configuration diagram of an attitude control system 130 according to Embodiment 1. FIG. 4 is a schematic diagram of attitude control according to Embodiment 1. FIG.

The same or corresponding elements are denoted by the same reference numerals in the embodiments and drawings. Descriptions of elements having the same reference numerals as those described will be omitted or simplified as appropriate.

Embodiment 1.
A configuration for controlling the attitude of the spacecraft 100 will be described with reference to FIGS. 1 to 4. FIG.

Spacecraft 100 is a man-made object that flies in space. A specific example of spacecraft 100 is a satellite.

*** Configuration description ***
The configuration of the spacecraft 100 will be described based on FIG.
The spacecraft 100 includes three or more CMGs 110 , a gimbal control system 120 and an attitude control system 130 .
CMG 110 is a control moment gyro used to control the attitude of spacecraft 100 .
A gimbal control system 120 is a system that controls the gimbal of the CMG 110 .
The attitude control system 130 is a system that controls the attitude of the spacecraft 100 by controlling the CMG 110 via the gimbal control system 120 .
Each of the gimbal control system 120 and the attitude control system 130 is composed of one or more devices. Also, the gimbal control system 120 may be provided in the same equipment as the CMG 110 .

The spacecraft 100 further comprises an installation interface 101P for each CMG110.
The installation interface 101P is a non-rigid installation interface. That is, the installation interface 101P is an installation interface having elasticity or viscoelasticity.
The installation interface is an interface for installing the CMG inside the spacecraft 100 .

Specifically, the installation interface 101P is a set of isolators.
An isolator is a material that has a vibration isolation effect. For example, the isolator is a spring damper or rubber vibration isolator.

In the following description, the number that identifies the CMG 110 is denoted as "i".
"i" is each integer of 1 or more and N or less.
“N” is the number of CMGs 110 .

The configuration of the CMG 110i will be described based on FIG.
CMG 110i comprises gimbal 111 , wheels 112 and body 113 . CMG 110 may comprise multiple gimbals 111 .
A gimbal 111 is connected to a body 113 and rotates about a gimbal shaft 114 as a rotation axis.
The wheel 112 is provided inside the gimbal 111 and rotates about an axis (spin axis) orthogonal to the gimbal axis 114 as a rotation axis.

Fuselage 113 is attached to the interface surface of spacecraft 100 using a set of isolators 101 .
One end of each isolator 101 is attached to the interface surface of the fuselage 113 and the other end of each isolator 101 is attached to the interface surface of the spacecraft 100 .
The interface surface is the surface to which a set of isolators 101 are attached.

In the CMG 110i, the coordinate system based on the interface surface of the spacecraft 100 is called the " _Di system".
The x-axis of the D _i system represents the direction of the gimbal axis 114 when the displacement of the set of isolators 101 is zero.
The z-axis of the D _i system represents the direction of the angular momentum h _i when the displacement of the set of isolators 101 is zero and the gimbal angle θ _i is zero.
The y-axis of the _Di system represents a direction orthogonal to the x-axis of the _Di system and the z-axis of the _Di system.
The x-axis of the D _i system is denoted as "x _di ", the y-axis of the D _i system is denoted as "y _di ", and the z-axis of the D _i system is denoted as "z _di ".

In the CMG 110i, the coordinate system based on the interface surface of the torso 113 is called the " _Ci system".
The x-axis of the C _i system represents the direction of the gimbal axis 114 .
The z-axis of the C _i system represents the direction of the angular momentum h _i of the CMG 110 i when the rotation angle (θ _i ) of the gimbal 111 is zero.
The y-axis of the _Ci system represents a direction orthogonal to the x-axis of the _Ci system and the z-axis of the _Ci system.
The x-axis of the C _i system is denoted as “x _ci ”, the y-axis of the C _i system is denoted as “y _ci ”, and the z-axis of the C _i system is denoted as “z _ci ”.

"h _i " represents the angular momentum of CMG 110i. Angular momentum h _i is generated in the direction of the spin axis by the rotation of wheel 112 of CMG 110i.
“θ _i ” represents the gimbal angle of the CMG 110i. A gimbal angle represents a rotation angle of the gimbal 111 .

"τ _i " represents the CMG torque. CMG torque is the torque produced by CMG 110i. Torque is represented by the magnitude in each direction of the three axes (x, y, z). In other words, torque is represented by a vector.
The CMG torque τ _i is produced by the rotation of the gimbal 111 and the rotation of the wheel 112 in the direction of the outer product of the gimbal angular velocity and the angular momentum _hi . The gimbal angular velocity is the angular velocity of the gimbal 111 .

“τ _gi ” represents the gimbal torque of CMG 110i. Gimbal torque is the torque applied by gimbal control system 120 to rotate gimbal 111 .
A gimbal torque τ _gi is produced in the direction of the gimbal axis 114 of the CMG 110i.

“k _* ” represents the stiffness value of the set of isolators 101 . The stiffness value represents the strength of stiffness. The stiffness value k _* is common to all CMGs 110 . Also, the stiffness value k _* is assumed to be a fixed value. A fixed value is a predefined value. However, the stiffness value k _* may be defined for each CMG 110i.
The stiffness value in the direction of rotation about the x-axis is denoted as " _kx ".
The stiffness value in the direction of rotation about the y-axis is denoted as "k _y ".
A stiffness value in the direction of rotation about the z-axis is denoted as "k _z ".

Rotation of the wheel 112 generates high-frequency minute vibrations (disturbances). Disturbance transmission is suppressed by a set of isolators 101 .
A set of isolators 101 are displaced under the influence of CMG torque _τi . Displacement of the set of isolators 101 causes the fuselage 113 to tilt.
The displacement in the direction of rotation about the x-axis is denoted as "ψ _xi ".
The displacement in the direction of rotation about the y-axis is denoted as "ψ _yi ".
The displacement in the direction of rotation about the z-axis is denoted as "ψ _zi ".

The configuration of the attitude control system 130 will be described based on FIG.
The attitude control system 130 includes a target generator 131 , an attitude estimator 132 , an attitude control regulator 133 , a steering law calculator 134 , an adder 135 and a displacement correction controller 140 .

Symbols shown in FIG. 3 will be described.
"θ" represents the gimbal angle vector. The gimbal _angle vector θ is ^a _vector {θ ₁ , .
The suffix "FF" means a component of feedforward control. The feedforward control calculates the amount of control in advance for posture change and the like. A conventional method may be used as a calculation method for feedforward control.
The suffix "FB" means the component of feedback control. Feedback control is to calculate a control amount for compensating an attitude error or the like.
The suffix "FF+FB" means a combination of the feedforward control component and the feedback control component.

“ω” represents the attitude angular velocity vector. The attitude angular velocity vector ω is a vector {ω _x , ω _y , ω _z } ^T in which the attitude angular velocities ω _* are arranged. The attitude angular velocity ω _* is the attitude angular velocity of the spacecraft 100 .
The suffix “S” means the S system. The S system is a coordinate system fixed to the spacecraft 100 .

"ψ _* " represents the displacement vector.
The _displacement vector ψ _x is a ^vector {ψ _x1 , _.
The displacement vector ψ _y is a ^vector _{{ψ y1 , .}
The displacement vector ψ _z is a ^vector _{ ψ _z1 , _.

"τ" represents the total torque. The total torque is the torque generated in the spacecraft 100 by all of the multiple CMGs 110 .
"τ _FB " represents the total torque τ for feedback control. τ _FB is the torque required to compensate for the attitude error, and may be calculated by a conventional method.
“τ _FBcor ” represents the correction torque. The corrected torque τ _FBcor is the total torque τ _FB after being corrected according to the present embodiment.

Each element of the attitude control system 130 will be described.
The target generator 131 is an element that calculates various target values.
The target generator 131 calculates a gimbal angle vector θ, a first derivative of the gimbal angle vector θ (gimbal angular velocity vector), and a second derivative of the gimbal angle vector θ (gimbal angular acceleration vector).
A first derivative value is a value obtained by first derivative. A second derivative value is a value obtained by second derivative.

The target generator 131 calculates the attitude angular velocity vector ^Sω .

The target generator 131 calculates target attitude information of the spacecraft 100 .
The target attitude information includes an attitude angle vector and an attitude angular velocity vector ^Sω .

The attitude estimator 132 is an element that calculates various estimated values.
The attitude estimator 132 calculates estimated attitude information of the spacecraft 100 .
The estimated attitude information includes an attitude angle vector and an attitude angular velocity vector.

The attitude control regulator 133 receives the target attitude information and the estimated attitude information and calculates the total torque τ _FB .

The steering law calculator 134 receives as inputs the gimbal angle vector θ, the first derivative of the gimbal angle vector θ, and the second derivative of the gimbal angle vector θ.
The steering law calculator 134 receives as inputs the previous gimbal angle vector θ _FB and the first derivative of the previous gimbal angle vector θ _FB .
A steering law calculator 134 receives the correction torque τ _FBcor as an input.
The steering law calculator 134 receives as input the Jacobian matrix A _FF+FB . The Jacobian matrix A _FF+FB will be described later.
Then, the steering law calculator 134 calculates the gimbal angle vector θ _FB , the first derivative of the gimbal angle vector θ _FB , and the second derivative of the gimbal angle vector θ _FB .

The adder 135 receives as inputs the gimbal angle vector θ _FF , the first derivative of the gimbal angle vector θ _FF , and the second derivative of the gimbal angle vector θ _FF .
The adder 135 receives as inputs the gimbal angle vector θ _FB , the first derivative of the gimbal angle vector θ _FB , and the second derivative of the gimbal angle vector θ _FB .
The adder 135 then calculates gimbal control information. The gimbal control information includes a gimbal angle vector θ, a first derivative of the gimbal angle vector θ, and a second derivative of the gimbal angle vector θ.
The gimbal angle vector θ is obtained by adding the gimbal angle vector _θFF and the gimbal angle vector _θFB (=θFF _+FB ).
The first-order differential value of the gimbal angle vector θ is obtained by adding the first-order differential value of the gimbal angle vector _θFF and the first-order differential value of the gimbal angle vector _θFB .
The second-order differential value of the gimbal angle vector θ is obtained by adding the second-order differential value of the gimbal angle vector _θFF and the second-order differential value of the gimbal angle vector _θFB .

The displacement correction controller 140 has a displacement estimator 141 and a torque corrector 142 .

Displacement estimator 141 is an element that calculates the displacement vector ψ _* of the set of isolators 101 for CMG 110i.
The displacement estimator 141 receives as inputs the gimbal angle vector θ _FF , the first derivative of the gimbal angle vector θ _FF , and the second derivative of the gimbal angle vector θ _FF .
The displacement estimator 141 receives as inputs the gimbal angle vector θ _FB and the first derivative of the gimbal angle vector θ _FB .
The displacement estimator 141 receives as input the first derivative of the gimbal angle vector θ _FF+FB .
The displacement estimator 141 receives the attitude angular velocity vector ^Sω as an input.
Then, the displacement estimator 141 calculates a displacement vector ψ _x , a displacement vector ψ _y , and a displacement vector ψ _z .

Torque corrector 142 is an element that corrects the overall torque τ _FB based on the displacement vector ψ _* of the set of isolators 101 for CMG 110i.
The torque compensator 142 receives as inputs the gimbal angle vector θ _FF , the first derivative of the gimbal angle vector θ _FF , and the second derivative of the gimbal angle vector θ _FF .
The torque compensator 142 receives the gimbal angle vector θ _FB and the first derivative of the gimbal angle vector θ _FB as inputs.
Torque compensator 142 receives as inputs displacement vector φ _x , displacement vector φ _y , and displacement vector φ _z .
Torque compensator 142 receives the overall torque τ _FB as an input.
Then, the torque corrector 142 calculates the correction torque _τFBcor .

In the attitude control system 130, the elements other than the displacement correction controller 140 are conventional elements, and the calculation method for each element may be the conventional method.

***Description of operation***
A calculation method in the displacement correction controller 140 will be described below.

First, a calculation method for calculating the displacement vector ψ _*i will be described.
A displacement vector ψ _*i is calculated by the displacement estimator 141 .

In the CMG 110i, the direction of the angular momentum h _i is controlled by controlling the gimbal angle θ _i .
Then, the direction of the CMG torque τ _i changes as the direction of the angular momentum _hi changes. The CMG torque τ _i is output in a direction perpendicular to the direction of rotation of the gimbal angle θ _i and the direction of the angular momentum _hi .

When the CMG torque τ _i is output, a displacement ψ _*i is produced in the set of isolators 101 . The magnitude of the displacement ψ _*i depends on the stiffness value k _* of the set of isolators 101 .
Normally, in order to stabilize each of the attitude control system 130 and the gimbal control system 120, the relationship "attitude control band<<gimbal control band<<isolator resonance frequency" often holds.
The attitude control band is a frequency band that can be controlled by the attitude control system 130 .
A gimbal control band is a frequency band that can be controlled by the gimbal control system 120 .
The isolator resonance frequency is the resonance frequency of each isolator 101 .
Since the attitude control band is lower than the isolator resonance frequency, the attitude control system 130 can hardly compensate for the displacement ψ _*i due to the vibration of each isolator 101 .
Therefore, the vibration of each isolator 101 is removed and attention is paid to the displacement ψ _*i .

The displacement ψ _*i can be expressed by equations (1-1) to (1-3).
“h _c ” represents the angular momentum of CMG 110i. The magnitude of the angular momentum _hc is common to all CMGs 110i. Also, the angular momentum h _c is assumed to be a fixed value. However, the angular momentum _hc may be defined for each CMG 110i. The angular momentum h _c may vary with time.

Equation (1-1) includes the gimbal torque τ _gi .
It is difficult for the attitude control system 130 to accurately estimate the gimbal torque τ _gi in the relationship of “attitude control band << gimbal control band”.
Also, the gimbal torque τ _gi is very small compared to the CMG torque τ _i . Therefore, the influence of the gimbal torque τ _gi on the displacement ψ _*i is minor.

Equations (1-2) and (1-3) include the attitude angular velocity ω _x of the C _i system.
The attitude angular velocity ω _x of the C _i system depends on the displacement ψ _x of the set of isolators 101 . However, the displacement ψ _x is very small.
Also, usually, the maximum value of the attitude angular velocity ω _* is often smaller than the maximum value of the gimbal angular velocity.
Furthermore, the relationship of formula (2-1) holds.

Therefore, the posture angular velocity ω _x of the C _i system can be approximated as in Equation (2-2).
“U _ab ” means a coordinate transformation matrix from the b system to the a system. That is, "U _sdi " means the coordinate transformation matrix from the D _i system to the S system.
“g _ci ” represents the gimbal axial direction vector. The gimbal axis direction vector g _ci is the direction vector of the gimbal axis 114 of CMG 110 i when the displacement ψ _*i is zero. The gimbal axial vector g _ci is a fixed value.
"c _c0i " represents the angular momentum direction vector. The angular momentum direction vector c _c0i is the direction vector of the angular momentum _{h c} when the displacement ψ _*i is zero and the gimbal angle θ _i is zero. The angular momentum direction vector c _c0i is a fixed value.

Equation (2-2) includes the attitude angular velocity vector ω of the S system.
The posture angular velocity vector ω of the S system may be either a target value calculated by the target generator 131 or an estimated value calculated by the posture estimator 132 . However, it is preferable to use the target value because the estimated value is affected by unwanted noise.

The gimbal angle _θi and the gimbal angular velocity can be considered separately into a feedforward component and a feedback component.
The feedforward component is preplanned for attitude changes and the like.
A feedback component is needed for attitude error compensation.

For the above reasons, formulas (1-1) to (1-3) can be replaced with formulas (3-1) to (3-3).
The suffix “z” means the previously calculated value.
“Δt” represents the processing time of one cycle of the attitude control system 130 .

Displacement estimator 141 calculates equations (2-1) and (2-2) to calculate equations (3-2) and (3-3).
The displacement estimator 141 calculates the displacement vector ψ _*i by calculating the equations (3-1) to (3-3).

Next, a calculation method for calculating the correction torque _τFBcor will be described.
The correction torque τ _FBcor is calculated by the torque corrector 142 .

The influence of the displacement vector ψ _*i calculated by the displacement estimator 141, that is, the estimated displacement vector ψ _*i on attitude control is divided into a first influence error and a second influence error.
The first influence error is "torque error due to error in Jacobian matrix".
The second effect error is "the angular momentum error caused by the change in direction of the angular momentum of the CMG".
The first effect error and the second effect error are compensated by the following control law.

The first effect error is compensated by correction of the Jacobian matrix.
Equation (4) shows the relational expression for the Jacobian matrix.
"A" represents the Jacobian matrix.

Equations (5-1) and (5-2) are defined based on Equation (4).
“A _FF0 ” represents the Jacobian matrix for feedforward control when the displacement ψ _*i is zero.
“A _FF+FB ” represents the Jacobian matrix for synthesis of feedforward control and feedback control when there is a displacement ψ _*i . A _FF+FB is used as the Jacobian matrix in the steering law calculator 134 to compensate for the first effect error.
“g _i ” represents the gimbal axial vector. The gimbal axis direction vector g _i is the direction vector of the gimbal axis 114 of CMG 110 i with displacement ψ _*i .
“c _0i ” represents the angular momentum direction vector. The angular momentum direction vector c _0i is the direction vector of the angular momentum h _c when there is a displacement ψ _*i and the gimbal angle θ _i is zero.
“h _0i ” represents the angular momentum h _i in the absence of displacement ψ _*i .

The estimated displacement vector ψ _*i is very small. Therefore, by calculating the equation (6), it is possible to calculate the gimbal axial direction vector g _i of the S system and the angular momentum direction vector c _0i of the S system.

Equation (7) is defined for the difference between the total angular momentum of CMGs 110 with displacement ψ _*i and the total angular momentum of CMGs 110 without displacement ψ _*i .
By calculating the equation (7), the correcting angular momentum of the S system can be calculated.
"Δh" represents the angular momentum for correction.
“Δh _i ” represents the corrective angular momentum for the displacement ψ _*i .

The correction torque τ _FBcor can be expressed by Equation (8).

Torque compensator 142 calculates equations (5-1), (5-2), (6), and (7) to calculate equation (8).
Torque corrector 142 calculates correction torque τ _FBcor by calculating equation (8).

The above Jacobian matrix A _FF+FB is also used in steering law calculator 134 .
The steering law calculator 134 can calculate the first derivative of the gimbal angle vector θ _FB by calculating equation (9).
"A ⁺ _FF+FB " represents the pseudo-inverse of the Jacobian matrix AFF _+FB .

Equation (9) is a basic relational expression, and Equation (9) can also be replaced with an equation that considers singularity avoidance rules and the like.

Based on FIG. 4, the outline of the displacement correction control will be described.
"CMG#i" means CMG 110i.
"Isolator #i" refers to the set of isolators 101 for CMG 110i.
“CMG#i gimbal angular velocity” in the displacement estimator 141 in _FIG .
"CMG#i output torque" means torque corresponding to each numerator of formulas (3-2) and (3-3). Usually, the dominant element of the CMG#i output torque is the product of the first derivative of the gimbal angular velocity θFF _+FBzi and the CMG angular momentum _hC .
“Isolator #i displacement estimation” means executing each calculation of equations (3-1) to (3-3).
“Jacobiian matrix correction” means calculating the Jacobian matrix by equation (5-2).
"Angular momentum displacement estimation" means calculating the corrective angular momentum according to equation (7).
"Torque correction" means calculating the correction torque according to equation (8).
"CMG#i gimbal angular velocity" in the steering law calculator 134 in the drawing means the first order differential value of the gimbal angle vector θ _FB calculated by equation (9).

The second effect error is the angular momentum error caused by the change in angular momentum direction due to the displacement φ _*i of the set of isolators 101 for each CMG 110i. An attitude angular velocity error and an attitude angle error occur due to the angular momentum error.
The only controllable element for each CMG 110i is the gimbal angle _θi .
The direction in which the angular momentum h _i changes with changes in the gimbal angle θ _i is orthogonal to the direction in which the angular momentum h i changes with the displacement ψ _{*i of} the set of isolators 101 .
Therefore, it is impossible to independently control only the gimbal angle _θi to compensate for the angular momentum error due to the displacement ψ _*i of the set of isolators 101 for each CMG 110i.
Therefore, the displacement estimator 141 calculates the displacement ψ _*i produced in the set of isolators 101 for the CMG 110i. A torque compensator 142 then integrates the displacement ψ _*i of the CMG 110i to correct the overall torque τ _FB .
Thereby, it is realized to compensate the angular momentum error due to the displacement φ _*i of the set of isolators 101 for CMG 110i in all of the plurality of CMGs 110 .

*** Effect of Embodiment 1 ***
The displacement correction controller 140 is used for attitude control of the spacecraft 100 having multiple CMGs 110 . Each CMG 110i is mounted on a non-rigid mounting interface 101P.
The displacement correction controller 140 has a displacement estimator 141 and a torque corrector 142 .
A displacement estimator 141 estimates the installation interface displacement.
A torque corrector 142 corrects the attitude control torque.
As a result, the influence of the displacement of the installation interface 101P on the attitude of the spacecraft 100 can be compensated.

The torque compensator 142 performs an integrated calculation for the entire system of the multiple CMGs 110 with respect to the displacement of the installation interface 101P generated by each CMG 110i. As a result, even under conditions where it is impossible to compensate for the influence of the displacement of the installation interface 101P for each CMG 110i, the influence of the displacement of the installation interface 101P generated in each CMG 110i is compensated.

*** Supplement to Embodiment 1 ***
Each element of attitude control system 130 may be a dedicated circuit, or a processor that executes a program stored in memory.
In other words, attitude control system 130 can be implemented in hardware, software, firmware, or a combination thereof.

Embodiment 1 is an illustration of a preferred embodiment, and is not intended to limit the technical scope of the present invention. Embodiment 1 may be partially implemented, or may be implemented in combination with other modes. The procedures described using flowcharts and the like may be changed as appropriate.

100 spacecraft, 101 isolator, 101P installation interface, 110 CMG, 111 gimbal, 112 wheel, 113 fuselage, 114 gimbal axis, 120 gimbal control system, 130 attitude control system, 131 target generator, 132 attitude estimator, 133 attitude control regulator, 134 steering law calculator, 135 adder, 140 displacement correction controller, 141 displacement estimator, 142 torque compensator.

Claims (4)

A displacement correction controller used in an attitude control system that controls the attitude of a spacecraft,
the spacecraft comprises the attitude control system, a plurality of control moment gyros, and a plurality of installation interfaces for installing the plurality of control moment gyros in the spacecraft;
The displacement correction controller comprises:
a displacement estimator that estimates a displacement of each of the plurality of installation interfaces;
a torque corrector for correcting the total torque of the plurality of control moment gyros calculated for feedback control by the attitude control system based on the estimated displacement of each of the plurality of installation interfaces;
with
The displacement estimator calculates the displacement of the installation interface for the control moment gyro based on the stiffness value of the installation interface for the control moment gyro, the gimbal angle of the control moment gyro, and the angular momentum of the control moment gyro. to estimate
Displacement compensation controller. The torque corrector calculates a correction angular momentum for the displacement of each installation interface based on the displacement of each installation interface and the torque of each control moment gyro when there is no displacement of each installation interface, and calculates the correction. 2. The displacement correction controller according to claim 1, wherein the total corrective angular momentum is calculated by summing the total corrective angular momentum, and the total torque is corrected based on the total corrective angular momentum. A displacement correction controller used in an attitude control system that controls the attitude of a spacecraft,
the spacecraft comprises the attitude control system, a plurality of control moment gyros, and a plurality of installation interfaces for installing the plurality of control moment gyros in the spacecraft;
The displacement correction controller comprises:
a displacement estimator that estimates a displacement of each of the plurality of installation interfaces;
a torque corrector for correcting the total torque of the plurality of control moment gyros calculated for feedback control by the attitude control system based on the estimated displacement of each of the plurality of installation interfaces;
with
The torque corrector calculates a correction angular momentum for the displacement of each installation interface based on the displacement of each installation interface and the torque of each control moment gyro when there is no displacement of each installation interface, and calculates the correction. calculating a total corrective angular momentum by summing the total angular momentum, and correcting the total torque based on the total corrective angular momentum;
Displacement compensation controller. 4. A displacement compensation controller according to any preceding claim, wherein each of said plurality of mounting interfaces is a set of isolators.

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