JP5484262B2 - Spacecraft attitude control device - Google Patents

Description

  The present invention relates to an attitude control device for a spacecraft that is mounted on a spacecraft such as an artificial satellite or a spacecraft and changes the attitude of the spacecraft by controlling a control moment gyro (CMG).

  A conventional spacecraft attitude control device includes an attitude profile calculator, an attitude estimator, a feedforward controller, a feedback controller, a CMG gimbal target profile generator, and a CMG gimbal controller. At the time of changing the attitude of the artificial satellite, the attitude profile calculator creates a trajectory plan for the attitude change and calculates the final target attitude (attitude angle and attitude angular velocity) of the artificial satellite. The attitude estimator estimates the current attitude (attitude angle and attitude angular velocity) of the artificial satellite.

  The feedforward controller calculates the control torque required for the attitude change in a feedforward manner based on the target attitude of the artificial satellite calculated by the attitude profile calculator. The feedback controller corrects the deviation between the target attitude of the artificial satellite and the current attitude based on the target attitude of the artificial satellite calculated by the attitude profile calculator and the current attitude of the artificial satellite estimated by the attitude estimator. The control torque is calculated in a feedback manner.

  Based on the control torque from the feedforward controller, the corrected control torque from the feedback controller, and the state of each CMG gimbal, the CMG gimbal target profile generator A target value of the rotational angular velocity (gimbal rotational angular velocity) is calculated. The CMG gimbal controller outputs the gimbal torque based on the target value of the gimbal rotation angle and gimbal rotation angular velocity calculated by the CMG gimbal target profile generator, and the current value of the gimbal rotation angle and gimbal rotation angular velocity, and controls the CMG. (For example, refer to Patent Document 1).

JP 2008-189235 A

However, the prior art has the following problems.
The target value of the gimbal rotation angle calculated by the CMG gimbal target profile generator of the conventional spacecraft attitude control device is the upper limit of the gimbal rotation angular velocity or gimbal rotation angular acceleration (rotation angular acceleration around the CMG gimbal axis). Since there is no guarantee that it will always satisfy, the attitude control system may lack reliability.

  Further, depending on the calculated target value of the gimbal rotation angle, there may be a state where a gimbal rotation angular velocity that realizes the control torque and the correction control torque cannot be obtained, that is, a so-called singular state. In this case, not only the desired attitude changing operation cannot be achieved, but also the attitude disturbance which cannot be ignored in the attitude change of the artificial satellite may occur.

  These problems are caused by the fact that, when changing the attitude of the artificial satellite, only the orbit plan for the attitude change is created in advance and the orbit plan for the gimbal rotation angle is not created in advance. In other words, the target values of the gimbal rotation angle and the gimbal rotation angular velocity are calculated each time based on the target attitude of the artificial satellite, the current attitude of the artificial satellite, and the state of each gimbal of the CMG, so the calculated values are too large. There is no means that can reliably avoid inappropriate target values.

As described above, in the conventional spacecraft attitude control device, there is no means for reliably and efficiently determining the trajectory plan of the gimbal rotation angle. Therefore, reliability is ensured in such an attitude control device that controls CMG. There is a problem that you can not.
Further, since a means for approximately calculating the trajectory plan of the gimbal rotation angle is not clear, there is a problem that a long calculation time is required even if a solution is obtained.
Further, since the orbit plan for the gimbal rotation angle and the orbit plan for the attitude change of the artificial satellite are not dynamically matched, there is a problem that an extra operation amount is required for the attitude change.

  The present invention has been made to solve the above-described problems, and the attitude of a spacecraft that can efficiently and reliably execute the attitude change of the spacecraft under the control of the CMG with high reliability. The object is to obtain a control device.

A spacecraft attitude control device according to the present invention is mounted on a spacecraft and controls the CMG to change the attitude of the spacecraft. When the attitude of the spacecraft is changed, Based on the target posture, calculating the trajectory plan of the rotation angle around the gimbal axis of the CMG as a time function connecting the rotation angles at a plurality of times within the time required for the rotation angle to reach the final value , from the sum of the angular momentum of the entire CMG obtained using the trajectory planning of the rotating angle, calculates the angular momentum of the spacecraft with the conservation of angular momentum, the posture angular velocity of the spacecraft obtained from the angular momentum time A trajectory planning unit that integrates and calculates a trajectory plan of the attitude angle of the spacecraft is provided, and the current value of the attitude angle of the spacecraft and the current value of the rotation angle are fed back to control the CMG.

According to the attitude control device for a spacecraft according to the present invention, the trajectory planning unit determines the trajectory plan for the rotation angle around the CMG gimbal axis based on the target attitude of the spacecraft when changing the attitude of the spacecraft. Was calculated as a time function connecting the rotation angles at multiple times within the time required from the initial value to the final value, and the trajectory plan of the attitude angle of the spacecraft was obtained using the rotation angle trajectory plan The attitude angular velocity of the spacecraft is calculated by time integration.
Therefore, the attitude change of the spacecraft by the control of CMG can be executed efficiently and reliably with high reliability.

It is a perspective view which shows schematic structure of CMG controlled by the attitude control apparatus of the spacecraft concerning Embodiment 1 of this invention. It is explanatory drawing which shows the example of arrangement | positioning in case the multiple CMG of FIG. 1 is arrange | positioned to an artificial satellite. It is a block block diagram which shows the attitude control apparatus of the spacecraft which concerns on Embodiment 1 of this invention with a peripheral device. It is a flowchart which shows operation | movement of the trajectory plan part in the attitude | position control apparatus of the spacecraft concerning Embodiment 1 of this invention. It is explanatory drawing which shows the relationship between the CMG gimbal rotation angle, the angular momentum of the whole CMG, the attitude angular velocity of the artificial satellite, and the attitude angle of the artificial satellite in the attitude control device for a spacecraft according to Embodiment 1 of the present invention. FIG. 3 is an explanatory diagram showing the overall angular momentum in the typical pyramid-arranged CMG shown in FIG. 2. It is explanatory drawing which shows the relationship between the gimbal rotation angle of CMG and the attitude angle of an artificial satellite in the attitude control device for a spacecraft according to Embodiment 2 of the present invention.

Hereinafter, preferred embodiments of the attitude control device for a spacecraft according to the present invention will be described with reference to the drawings. In the drawings, the same or corresponding parts will be described with the same reference numerals.
In the following embodiments, a case where the spacecraft attitude control device is mounted on an artificial satellite will be described as an example. However, the present invention is not limited to this, and the spacecraft attitude control device is a spacecraft or the like. It may be mounted on other spacecraft.

Embodiment 1 FIG.
FIG. 1 is a perspective view showing a schematic configuration of CMG 1 controlled by the attitude control device for a spacecraft according to Embodiment 1 of the present invention. In FIG. 1, CMG1 is an actuator used for attitude control of an artificial satellite, and a wheel that rotates at high speed around a spin axis, and a gimbal that supports a wheel and rotates around a gimbal axis orthogonal to the spin axis. It consists of and.

  In CMG1, the wheel has a certain amount of angular momentum around the spin axis by rotating at high speed. Here, by rotating the gimbal around the gimbal axis and rotating the direction of the angular momentum of the wheel, a large reaction torque acts on the satellite body. This reaction torque is used for attitude control of the satellite.

  FIG. 2 is an explanatory diagram showing an arrangement example when a plurality of CMGs 1 are arranged on an artificial satellite. In FIG. 2, about four CMGs 1 are normally arranged on the artificial satellite, and the attitude control of the artificial satellite is realized by emphasizing and operating these plural CMGs 1. The arrangement of CMG1 shown in FIG. 2 is called a pyramid arrangement because each CMG1 is arranged at the bottom of the quadrangular pyramid, and is one of typical arrangement examples of CMG1.

  FIG. 3 is a block configuration diagram showing the spacecraft attitude control device 10 (hereinafter abbreviated as “attitude control device 10”) according to Embodiment 1 of the present invention, together with CMG1 and the satellite body 2 as peripheral devices. It is. In FIG. 3, each CMG 1 is controlled by a gimbal control torque output from the attitude control device 10.

  As described above, when the wheel and gimbal of the CMG 1 rotate, a reaction torque acts on the artificial satellite body 2 and the attitude of the artificial satellite body 2 is changed. The CMG 1 and the satellite main body 2 output the current value of the gimbal rotation angle and the current value of the attitude angle of the artificial satellite to the attitude control device 10, respectively.

  The attitude control device 10 includes a target attitude calculation unit 11, a trajectory planning unit 12, a subtraction unit 13, an attitude control unit 14, a steering control unit 15, an addition unit 16, a subtraction unit 17, and a gimbal control unit 18. Here, the attitude control device 10 is constituted by a microprocessor (not shown) having a CPU and a memory storing a program, and each block constituting the attitude control device 10 is stored in the memory as software. Yes.

Hereinafter, functions of each part of the posture control apparatus 10 having the above-described configuration will be described.
The target attitude calculation unit 11 calculates the final target attitude of the artificial satellite and outputs it to the orbit planning unit 12 when changing the attitude of the artificial satellite.

  When the target attitude in the attitude change of the artificial satellite is given from the target attitude calculation unit 11, the trajectory planning unit 12 calculates a plan value (orbit plan) of the gimbal rotation angle of each CMG 1 and outputs it to the addition unit 16. A plan value (orbit plan) of the attitude angle of the artificial satellite that dynamically matches the gimbal rotation angle plan value is calculated and output to the subtractor 13. Here, the trajectory planning unit 12 is a main part of the present application, and detailed operations will be described later.

The subtraction unit 13 feeds back and subtracts the current attitude angle value of the artificial satellite input to the attitude control device 10 from the attitude angle plan value of the artificial satellite calculated by the trajectory planning unit 12, and calculates the calculated deviation. Output to the control unit 14.
The attitude control unit 14 generates a feedback attitude control torque so as to make the deviation calculated by the subtraction unit 13 zero, and outputs it to the steering control unit 15.

  The steering control unit 15 calculates a correction value of the gimbal rotation angle of each CMG 1 corresponding to the feedback posture control torque generated by the posture control unit 14. Note that when CMG1 is close to the singular state, the gimbal rotation angle correction value may take an excessive value. However, when CMG1 is close to the singular state, the steering control unit 15 of the present invention may perform gimbal rotation. The angle correction value is configured to be suppressed to a low value.

  The reason why such a configuration is possible is that the trajectory planning unit 12 calculates the gimbal rotation angle plan value of each CMG 1. That is, even when CMG1 is close to the singular state, the attitude change of the artificial satellite is performed without delay because the planned gimbal rotation angle is calculated.

The adding unit 16 adds the gimbal rotation angle planned value calculated by the trajectory planning unit 12 and the gimbal rotation angle correction value calculated by the steering control unit 15 to each other, and subtracts it as the target value of the gimbal rotation angle of each CMG1. To the unit 17.
The subtraction unit 17 feeds back and subtracts the gimbal rotation angle current value input to the attitude control device 10 from the gimbal rotation angle target value calculated by the addition unit 16 and outputs the calculated deviation to the gimbal control unit 18. To do.

The gimbal control unit 18 generates a gimbal control torque so that the deviation calculated by the subtraction unit 17 is zero, and outputs the generated gimbal control torque to each CMG 1.
As described above, the orbital planning unit 12 calculates the attitude angle planned value of the satellite and the gimbal rotation angle planned value of each CMG1, and performs feedback control so as to realize these planned values, thereby achieving high reliability. The attitude control device 10 for the spacecraft can be configured.

Next, the operation of the trajectory planning unit 12 in the spacecraft attitude control device 10 according to the first embodiment of the present invention will be described in detail with reference to the flowchart of FIG.
First, the trajectory planning unit 12 sets a time T required for the gimbal rotation angle θ _i to reach the final value from the initial value when the gimbal rotation angle of the i-th CMG 1 is θ _i (step S1).

The initial value and final value of the gimbal rotation angle θ _i can be obtained from the conditions of the attitude angle, angular velocity, and angular acceleration of the artificial satellite. Note that the time T required for the gimbal rotation angle θ _i from the initial value to the final value can be estimated from the attitude change of the artificial satellite. In order to obtain the value, a relatively small time T is initially set.

Next, the trajectory planning unit 12 sets the gimbal rotation angle theta _i at a certain time t _j, excluding the time of the initial and final values of the gimbal angle of rotation θ _{i (t j) (step} S2).
Here, the trajectory planning unit 12 may set any number of times t _j during the time T required for the gimbal rotation angle θ _i to reach the final value from the initial value. The gimbal rotation angle θ _i (t _j ) at time t _{j in} the middle is a parameter for calculating the trajectory plan of the gimbal rotation angle θ _i .

For example, setting the two times _{t 1} and time _{t 2,} except the time of the initial and final values of the gimbal rotation angle theta _i, the number of time _t gimbal rotation angle _j θ _{i (t j)} is, Cmgl the number 2N times 2 times N. In other words, the 2N parameters can be used to control the final attitude of the artificial satellite to a desired value, as will be described later.

At this time, the time t _j does not necessarily have to be a certain moment, and is assumed to have a constant value between time t ₁ and time t ₂ as a trajectory plan of the gimbal rotation angle θ _i. For example, θ _i (t ₁ ) = θ _i (t ₂ ), which is a constant value therebetween, can be handled as one parameter.

Subsequently, when the gimbal rotation angle θ _i (t _j ) at an intermediate time t _j is set, the trajectory planning unit 12 sets a time function θ _i (t) () representing the gimbal rotation angle θ _i at an arbitrary time t. i is an integer from 1 to N) (hereinafter referred to as “gimbal rotation angle trajectory plan θ _i (t)”) (step S3).

More specifically, first, if the gimbal rotation angle θ _i (t _j ) at an intermediate time t _j is set as a parameter, the gimbal rotation angle θ _i (t) at an arbitrary time t is Can be obtained by connecting the gimbal rotation angles θ _{i of the two} by a time polynomial.

For example, assuming that the initial time is t ₀ , the final time is t _f, and the gimbal rotation angle θ _i (t ₁ ) at the intermediate time t ₁ is obtained, any time between time t ₀ and time t ₁ is obtained. gimbal pivot angle in t θ _i (t) is, θ _{i (t 0)} at t = _{t 0,} and the given by θ _{i (t 1)} to become such n next time polynomial in t = _{t 1.}

Here, the degree n of the time polynomial gimbal rotation angle at time t ₀ and time t _1, the gimbal rotational angular velocity and or gimbal angular acceleration, the posture angular velocity of the satellite from the time t ₀ until time t _1, the posture It depends on the number of constraints on the angular acceleration and attitude angular jerk, and on the gimbal rotation angular velocity and gimbal rotation angular acceleration.

The same applies to the gimbal rotation angle θ _i (t) at any time t between time t ₁ and time t _f , and becomes θ _i (t ₁ ) at t = t ₁ , and θ at t = t _f . It is given by an n-th order time polynomial such that _i (t _f ). The degree n of the time polynomial may be different from the degree of the time polynomial between time t ₀ and time t ₁ . The time polynomial is determined so that not only the gimbal rotation angle is continuous but also the gimbal rotation angular velocity and the gimbal rotation angular acceleration are continuous at time t ₁ .

Next, the trajectory planning unit 12, based on the trajectory planning of the gimbal rotation angle obtained in step S3 θ _{i (t),} and calculates the final attitude of the satellite at the final time t _f (step S4).
Specifically, first, if the trajectory plan θ _i (t) of the gimbal rotation angle is obtained, the sum of the angular momentums of the CMG 1 can be calculated at an arbitrary time t.

  At this time, according to the law of conservation of angular momentum, even if CMG1 moves, the sum of the angular momentum of the entire satellite is kept at a constant value. Therefore, the change from the initial value of the angular momentum of the artificial satellite body 2 at the time t is equal to the change from the initial value of the angular momentum of the CMG 1 at the time t.

Thereby, the angular momentum of the artificial satellite body 2 at the time t can be calculated, and the attitude angular velocity of the artificial satellite body 2 at the time t can be calculated from the angular momentum. Further, by integrating this attitude angular velocity from the initial time t ₀ to the final time t _f , the final attitude of the artificial satellite body 2 can be calculated.

Subsequently, the trajectory planning unit 12, the final orientation of the satellite in the final time t _f calculated in step S4 determines whether matches the target satellite's attitude (step S5).
Here, as the target attitude of the artificial satellite, a value given from the target attitude calculation unit 11 when the attitude of the artificial satellite is changed is used.

In step S5, when the final position of the satellite is determined to be equal to the target satellite's attitude (i.e., Yes), the trajectory planning unit 12, from the initial time t ₀ to the final time t _f Then, it is determined whether or not the attitude angular velocity, attitude angular acceleration, attitude angular jerk, and gimbal rotation angular velocity and gimbal rotation angular acceleration of the artificial satellite satisfy the constraint conditions (step S6).

In step S6, when it is determined that the constraint condition is satisfied (that is, Yes), the trajectory planning unit 12 ends the process of FIG. The trajectory plan θ _i (t) of the gimbal rotation angle at this time is output to the adding unit 16 as the planned value of the gimbal rotation angle of each CMG 1.

On the other hand, in step S5, the final orientation of the satellite is not equal to the target satellite's attitude (i.e., No) if it is determined that the trajectory planning unit 12, the time t _j of the gimbal rotation angle theta _{i (t It} is determined whether or not the number of _j ) corrections (described later) is less than a predetermined upper limit value (step S7).

In step S7, when it is determined that the number of corrections of the gimbal rotation angle θ _i (t _j ) at time t _j is less than the predetermined upper limit value (that is, Yes), the trajectory planning unit 12 performs step S2. The gimbal rotation angle θ _i (t _j ) at time t _j set in (or step S8) is corrected (step S8), and the process proceeds to step S3.

Here, in the correction of the gimbal rotation angle θ _i (t _j ) at time t _j in step S8, a value obtained by adding a small amount Δθ _i (t _j ) to θ _i (t _j ) using, for example, the Newton method or the like. Is a new gimbal rotation angle θ _i (t _j ).

On the other hand, when it is determined in step S7 that the number of corrections of the gimbal rotation angle θ _i (t _j ) at time t _j is greater than or equal to a predetermined upper limit value (that is, No), the trajectory planning unit 12 The time T required for the gimbal rotation angle θ _i from the initial value to the final value set in step S1 (or step S9 described later) is corrected (step S9), and the process proceeds to step S2.

More specifically, if the attitude of the artificial satellite does not converge even when the gimbal rotation angle θ _i (t _j ) at the time t _j is corrected a certain number of times (predetermined upper limit value) or more, a step is performed. It is considered that the time T set for the gimbal rotation angle θ _i from the initial value to the final value set in S1 or step S9 is too short to obtain a solution.

Therefore, in the correction of the time T in step S9, a value obtained by adding the minute time ΔT to the time T is set as a new time T required for the gimbal rotation angle θ _i to reach the final value from the initial value.
Note that even if the value of the minute amount Δθ _i (t _j ) added to θ _i (t _j ) in step S8 becomes too large, it is considered that the time T is too short and no solution can be obtained. You may put the magnitude | size of the value of micro amount ( _DELTA ) ( _theta ) _i (tj) into the judgment object in S7.

  If it is determined in step S6 that the constraint condition is not satisfied (that is, No), the trajectory planning unit 12 immediately proceeds to step S9.

More specifically, _if the orbit plan of the attitude angle of the artificial satellite and the orbit plan of the gimbal rotation angle of each CMG ₁ from the initial time t ₀ to the final time t _f does not satisfy the constraint condition, As in the case where the attitude of the satellite does not converge, it is considered that the time T required for the gimbal rotation angle θ _i from the initial value to the final value is too short to obtain a solution. Therefore, when it determines with No in step S6, time T is correct | amended by step S9 and the process from step S2 is performed again.

Since the solution is always obtained by increasing the time T required for the gimbal rotation angle θ _i from the initial value to the final value, the processing in FIG. That is, in step S6, the condition determination is Yes, and the process in FIG. 4 can be completed. The trajectory plan θ _i (t) of the gimbal rotation angle at this time is output to the adding unit 16 as the planned value of the gimbal rotation angle of each CMG 1.

The trajectory planning unit 12 also determines the attitude angular velocity and attitude angular acceleration of the satellite based on the trajectory plan θ _i (t) of the gimbal rotation angle and a dynamic law such as an equation of motion of the satellite or an angular momentum conservation law. Is calculated as a time function, and the time function is further time integrated to calculate the planned attitude angle of the artificial satellite and output it to the subtractor 13.

In this way, the planned value of the gimbal rotation angle is given as a time function, and the time function is expressed as the time T required for the gimbal rotation angle θ _i to reach the final value from the initial value and the gimbal rotation angle θ at a certain time t _j . _i (t _j ) is obtained as a parameter. As a result, the time required for the gimbal rotation angle θ _i to reach from the initial value to the final value (the time required for the artificial satellite to reach the final value from the attitude initial value) is as short as possible, and with a small number of parameters, The orbit plan θ _i (t) for the gimbal rotation angle and the orbit plan for the attitude angle of the artificial satellite corresponding to the orbit plan can be calculated efficiently and reliably.

In addition, the trajectory plan θ _i (t) of the gimbal rotation angle that deviates from the above-mentioned constraint condition does not appear, and even if the CMG1 is close to a singular state, the trajectory plan of the gimbal rotation angle θ _i (t ) Is calculated, an inappropriate target value does not appear, and the attitude change of the satellite is performed without delay.

As described above, according to the first embodiment, when changing the attitude of the spacecraft, the trajectory planning unit calculates the trajectory plan of the rotation angle around the gimbal axis of the CMG based on the target attitude of the spacecraft. Calculated as a time function connecting the rotation angles at multiple times within the time required from the initial value to the final value, and the spacecraft attitude angle trajectory plan obtained using the rotation angle trajectory plan The attitude angular velocity of the aircraft is calculated by time integration.
Therefore, the attitude change of the spacecraft by the control of CMG can be executed efficiently and reliably with high reliability.

  That is, with such a configuration, the change in attitude of the spacecraft and the control of the gimbal rotation angle of each CMG do not interfere with each other at the planned value stage. Control can be integrated without difficulty, and the attitude change of a spacecraft using a plurality of CMGs can be performed efficiently.

Next, a specific example in which the interval between the initial value and the final value of the gimbal rotation angle θ _i is divided into three sections will be described with reference to FIG. FIG. 5 is a diagram illustrating the relationship between the CMG1 gimbal rotation angle, the angular momentum of the entire CMG1, the attitude angle velocity of the artificial satellite, and the attitude angle of the artificial satellite in the attitude control device for a spacecraft according to the first embodiment of the present invention. FIG.

In FIG. 5, the trajectory planning unit 12 divides the interval between the initial value and the final location in the trajectory plan θ _i (t) of the gimbal rotation angle into three sections, and two times excluding the time of the initial value and the final value. The gimbal rotation angle θ _i (t ₁ ) and the gimbal rotation angle θ _i (t ₂ ) of the i-th CMG _{1 at} t ₁ and time t ₂ are used as parameters.

At this time, the total number of parameters is 2N if the number of CMG1 is N, but if the gimbal rotation angle θ _i (t ₁ ) is equal to the gimbal rotation angle θ _i (t ₂ ), the number of parameters is N. Become. Further, if the gimbal angle of rotation of Cmgl between time t ₁ and time t ₂ is constant, also constant angular momentum of the entire Cmgl.

  Therefore, as described above, since the sum of the angular momentum of the entire satellite is maintained at a constant value, the angular momentum of the artificial satellite body 2 is obtained by adding the opposite sign to the angular momentum of the entire CMG1. As shown in FIG. 5, the artificial satellites change their attitudes with a substantially trapezoidal velocity pattern.

  Here, taking a typical pyramid-arranged CMG 1 as shown in FIG. 2 as an example, the overall angular momentum has a magnitude as shown in FIG. 6 according to the direction of angular momentum. . FIG. 6 shows how much angular momentum can be generated in the CMG 1 as a whole with respect to a certain direction from the origin, and each CMG 1 has 1 angular momentum.

The attitude change of the artificial satellite can be expressed by rotation around a certain rotation axis called Euler axis. Accordingly, the direction obtained by multiplying the inertial matrix of the artificial satellite and the Euler axis is the direction of the angular momentum to be generated by the artificial satellite body. Therefore, the gimbal rotation angle of each CMG 1 is obtained so that the sum of the angular momentums of the CMG 1 is maximized in the opposite direction, and the gimbal rotation angle is determined as the gimbal rotation angle θ _i (t ₁ ) and the gimbal rotation angle θ. By setting _i (t ₂ ), the angular velocity for posture change around the Euler axis in this case is the maximum angular velocity that can be realized by the given CMG 1. Therefore, the time required for rotation around the Euler axis is a minimum value.

In this way, the planned value of the gimbal rotation angle is determined so that the sum of the angular momentum of the entire CMG1 is reversed and maximized in the direction of angular momentum to be given to the artificial satellite determined based on the target attitude in the attitude change. did. Thereby, in the given CMG1, a trajectory plan θ _i (t) of the gimbal rotation angle that can realize the attitude change in the shortest time and a trajectory plan of the attitude angle of the artificial satellite according to the trajectory plan are calculated. be able to.
Actually, the state in which the sum of the angular momentum in the entire CMG 1 is the maximum is one of the singular states in the sense that the angular momentum cannot be generated any more. Therefore, the planned value of the gimbal rotation angle may be determined so that the angular momentum generated by the entire CMG 1 becomes a value slightly smaller than the maximum generated angular momentum.

Embodiment 2. FIG.
In the first embodiment, the case where only the final attitude (target attitude) in the attitude change of the artificial satellite is given to the trajectory planning unit 12 from the target attitude calculating unit 11 has been described. In the second embodiment, a case will be described in which not only the final attitude of the attitude angle of the artificial satellite but also a trajectory plan from the initial attitude to the final attitude is given to the orbit planning unit 12 in advance.

FIG. 7 is an explanatory diagram showing the relationship between the gimbal rotation angle of CMG1 and the attitude angle of the artificial satellite in the attitude control device for a spacecraft according to Embodiment 2 of the present invention. As shown in FIG. 7, when the trajectory planning unit 12 is given not only the final attitude of the attitude angle of the artificial satellite but also the movement from the initial attitude to the final attitude as the orbit plan, It is difficult to analytically obtain a gimbal rotation angle trajectory plan θ _i (t) that matches the attitude angle trajectory plan.

  Therefore, in this case, the trajectory planning unit 12 obtains the gimbal rotation angle corresponding to the initial attitude and the final attitude of the artificial satellite, and the gimbal rotation angle is connected smoothly (continuously) between the intervals. The trajectory plan is expressed by an appropriate time function (time polynomial).

  The trajectory planning unit 12 calculates a trajectory plan of the attitude angle of the artificial satellite that dynamically matches the gimbal rotation angle given by this time function, and the calculated trajectory plan is given from the target attitude calculation unit 11. A parameter (time polynomial coefficient) included in the time function indicating the trajectory plan of the gimbal rotation angle is determined so as to be sufficiently close to the trajectory plan of the posture change.

Thus, the parameters included in the time function indicating the trajectory plan of the gimbal rotation angle so that the trajectory plan calculated by the trajectory planner 12 is sufficiently close to the trajectory plan of posture change given from the target posture calculator 11. Even when a trajectory plan of the attitude angle of the artificial satellite is given, the trajectory plan θ _i (t) of the continuous gimbal rotation angle corresponding to the trajectory plan can be approximately calculated. it can.

  1 CMG, 2 artificial satellite body, 10 spacecraft attitude control device, 11 target attitude calculation unit, 12 orbit planning unit, 13 subtraction unit, 14 attitude control unit, 15 steering control unit, 16 addition unit, 17 subtraction unit, 18 Gimbal control unit.

Claims (4)

A spacecraft attitude control device that is mounted on a spacecraft and controls the CMG (control moment gyro) to change the attitude of the spacecraft.
When changing the attitude of the spacecraft, based on the target attitude of the spacecraft, a trajectory plan of the rotation angle around the gimbal axis of the CMG is calculated in a plurality of times within the time required for the rotation angle to reach the final value from the initial value. and calculates as a function of time for connecting the rotation angle at a time, from the CMG total angular momentum of the sum obtained with the trajectory planning before Symbol rotation angle, the angular momentum of the spacecraft with the law of conservation of angular momentum And calculating a trajectory plan of the attitude angle of the spacecraft by time integrating the attitude angular velocity of the spacecraft calculated from the angular momentum ,
An attitude control device for a spacecraft, wherein the CMG is controlled by feeding back a current value of an attitude angle of the spacecraft and a current value of the rotation angle.   The trajectory planning unit divides between the initial value and final value of the rotation angle into a plurality of sections, the rotation angle at the end points of each section excluding the initial value and final value, and the rotation angle from the initial value. Using the time required to reach the final value as a parameter, a constraint condition including at least one of the attitude angular velocity, attitude angular acceleration and attitude angular jerk of the spacecraft, and the rotation angular velocity and rotation angular acceleration around the gimbal axis The spacecraft attitude control device according to claim 1, wherein the rotation angle trajectory plan is calculated by expressing the rotation angle of each section as a time function so as to satisfy.   The trajectory planning unit divides the interval between the initial value and final value of the rotation angle into three sections, and calculates the two rotation angles at the end points of each section excluding the initial value and final value of the spacecraft. In the direction of angular momentum to be given to the spacecraft determined based on a target attitude, the trajectory plan of the rotation angle is calculated so that the sum of the angular momentum of the entire CMG is maximized in the opposite direction. The attitude control device for a spacecraft according to claim 1.   The trajectory planning unit obtains a rotation angle corresponding to the initial attitude and the final attitude of the attitude angle of the spacecraft when a trajectory plan of the attitude angle of the spacecraft is given in advance, and the initial attitude and the final attitude The time function indicating the rotation angle trajectory plan is set so that the sections between the two are smoothly connected, and the spacecraft attitude angle trajectory plan is calculated based on the time function, and the calculated space 4. The parameter of the time function of the rotation angle is determined so that a trajectory plan of the attitude angle of the aircraft approaches a trajectory plan of the attitude angle of the spacecraft given in advance. The attitude control device for a spacecraft according to any one of the above.

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