JP2006008132A - Posture controller for artificial satellite - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a posture controller for an artificial satellite which posture controller changes the posture of the satellite at high speed, utilizing a large output torque from a thruster, and at the same time stabilizes the satellite quickly after the posture change, utilizing continuous output torques from a wheel. <P>SOLUTION: The posture controller comprises a posture control feedforward calculation section which calculates a feedforward control torque on the basis of an output from a posture angle target generating section and converts the feedforward control torque into a thruster drive signal to put it out, the thruster which is used as an actuator for controlling the posture of the satellite by an external force and controls the posture of the satellite on the basis of the thruster drive signal, a posture control feedback calculation section which receives feedback of a posture angle of the satellite, calculates a feedback posture control torque on the basis of a posture angle feedback output and a posture angle target, and converts the feedback posture control torque into a wheel drive signal to put it out, and the wheel which is used as an actuator for controlling the posture of the satellite by an internal force and controls the posture of the satellite on the basis of the wheel drive signal. <P>COPYRIGHT: (C)2006,JPO&NCIPI

Description

  The present invention relates to an attitude control device for an artificial satellite that changes the attitude of the artificial satellite at high speed.

  FIG. 7, which is disclosed in, for example, Japanese Patent Application Laid-Open No. 06-72396, shows the operating principle of an attitude control device for an artificial satellite that suppresses excitation of structural vibration while controlling the attitude of the artificial satellite using a thruster. FIG.

  When applying a pulse input by a thruster to an artificial satellite having a structural vibration mode, a vibration mode of a low frequency is mainly excited by a pulse injection of the thruster. In order to suppress the excitation of this low frequency vibration, the same pulse input may be given after ½ period of the vibration. As a result, so-called dead beat control is performed, and low-frequency vibration after pulse input is suppressed.

  Furthermore, considering the case where the satellite has a higher frequency vibration mode, it is possible to divide the first pulse and the pulse after 1/2 period of the low frequency vibration in accordance with the period of the high frequency vibration mode. It is done. For example, as shown in FIG. 7, if each pulse is divided into three and the pulse interval is set to 1/3 period of the high-frequency vibration mode, deadbeat control is performed even for the high-frequency vibration mode. Can be achieved.

Fig. 8 shows the attitude of an artificial satellite using both a conventional wheel and thruster based on the explanatory text shown in "Aerospace Engineering Handbook" (edited by the Japan Aerospace Society, Maruzen, p. 1020 (1992)). It is a block diagram for demonstrating the operation principle of a control apparatus.
In FIG. 8, 1 is an artificial satellite body, 2 is an attitude control calculation unit that performs calculations for attitude control of the artificial satellite body 1, and 3 is an actuator that performs attitude control of the artificial satellite body 1 with the reaction torque of the rotor, that is, internal force. Wheel, 12 is an actuator thruster that controls the attitude of the satellite body 1 by gas injection, that is, external force, 13 is a thruster drive arithmetic unit that generates a thruster drive signal from the wheel angular momentum value, and 14 is the attitude angle of the satellite. It is a posture estimation calculation unit that performs a calculation for estimating a posture angular velocity.

Next, the operation according to the above configuration will be described.
In the attitude control of the artificial satellite body 1, normally, the attitude angle and attitude angular velocity output from the attitude estimation calculation unit 14 are fed back, and the attitude control calculation unit 2 sets the attitude angle to a target value (0 in the figure). The control calculation such as PD (proportional differentiation) control necessary to be performed is performed to generate a wheel drive signal to drive the wheel 3. When a disturbance acts on the artificial satellite, the wheel 3 is driven to cancel the action of the disturbance and maintain the posture of the artificial satellite body 1, but the angular momentum of the wheel 3 depends on the nature of the disturbance. There is a possibility that it will increase beyond the allowable value. Therefore, the angular momentum of the wheel 3 is input to the thruster drive calculation unit 13, and the thruster 12 is driven so that the angular momentum of the wheel 3 does not exceed a certain value, thereby releasing the angular momentum of the wheel 3. Control operations are performed simultaneously.

Further, FIG. 9 is an artificial image from the output of the attitude sensor and the output of the attitude angular velocity sensor based on the explanatory text shown in “Introduction to Space Engineering” (written by Masamichi Mobara, Bafukan, pp. 149-150 (1994)). It is a block diagram for demonstrating the operation | movement principle of the attitude estimation calculating part 14 of the conventional artificial satellite which estimates the attitude angle and attitude | position angular velocity of a satellite.
In FIG. 9, 15 is an attitude sensor, 16 is an angular velocity sensor in a low speed mode, 17 is an angular velocity sensor in a high speed mode, 18 is a bias estimation unit of the angular velocity sensor 16 in a low speed mode, and 19 is a bias estimation unit of the angular velocity sensor 17 in a high speed mode. , 20 is a posture estimation unit at the normal posture, and 21 is a posture estimation unit at the time of posture change.

Next, the operation according to the above configuration will be described.
In a normal operation mode, the artificial satellite 1 is controlled so as to maintain a steady posture. The attitude angular velocity of the artificial satellite assumed in such a normal attitude is small, and the angular velocity sensor 16 in the low-speed mode that is used usually has a narrow measurement range to improve detection accuracy. When the output of the angular velocity sensor 16 is integrated, a change in angular velocity is obtained. The initial value of the posture angle is added to this to obtain the posture angle.

  The output of the posture sensor 15 is used as the initial value of the posture angle. However, when the output of the angular velocity sensor 16 is integrated over a long period of time, the influence of a bias-like error component included in the output of the angular velocity sensor 16 appears significantly. Therefore, the estimated value of the posture angle is compared with the output of the posture sensor 15, and the bias estimation unit 18 estimates a bias-like error component. The estimated value α of the bias angle is subtracted from the output of the angular velocity sensor 16 to obtain the estimated value α of the posture angular velocity, and the estimated value α of the posture angular velocity is integrated to obtain the estimated value θ of the posture angle. The above is the operation of the posture estimation unit 20 in the steady posture.

  On the other hand, there is an operation mode in which the attitude of the artificial satellite 1 is changed, such as changing the axial direction of the thruster in the inertial space for orbit control or changing the directional axis of the mission sensor in the inertial space for observation. The attitude angular velocity of the artificial satellite assumed at the time of such an attitude change is large, and the angular velocity sensor 17 of the high-speed mode to be used usually decreases the detection accuracy and widens the measurement range. Similar to the posture estimation unit 20 in the steady posture, the angular velocity change is obtained by integrating the output of the angular velocity sensor 17. The initial value of the posture angle is added to this to obtain the posture angle.

  As the initial value of the posture angle, the final value of the posture angle estimated value of the posture estimation unit 20 in the normal posture before switching the operation mode is used. When the posture is changed, it is conceivable that the output of the posture sensor 15 cannot be used because the detection target is out of the field of view of the posture sensor 15 or the usable bandwidth of the posture sensor 15 is exceeded. Therefore, in order to correct a bias-like error component included in the output of the angular velocity sensor 17, the bias estimation unit 19 performs calibration obtained by processing the data acquired in the above-described operation on the ground test or orbit offline. Use the action data. Thereby, the posture angle and the posture angular velocity can be estimated even during the posture change. The above is the operation of the posture estimation unit 21 at the time of posture change.

  The posture estimation calculation unit 14 selects either the posture estimation unit 20 during steady posture or the posture estimation unit 21 during posture change according to the operation mode, and performs posture control on the posture angle estimation value θ and the posture angular velocity estimation value α. Output to the feedback calculation unit 7.

Japanese Patent Laid-Open No. 06-72396 "Aerospace Engineering Handbook" (Japan Aerospace Society, edited by Maruzen, p. 1020 (1992)) “Introduction to Space Engineering” (Masahara Mobara, Bafukan, pp. 149-150 (1994))

  Since conventional attitude control devices for artificial satellites are considered on the premise that attitude control torque is applied in a pulsed manner as described above, the generation of attitude angular acceleration is also pulsed, as in the case of a large attitude angle change. However, there is a problem that it is not possible to cope with the case where the posture angular acceleration needs to be generated for a long time. In addition, since the conventional attitude control method mainly considers suppressing the excitation of structural vibration, no measures are taken for the possibility that an error occurs in the rigid attitude angle due to the presence of structural vibration. There was a problem.

  Further, when attitude control is performed using both the thruster 12 and the wheel 3, the thruster 12 is used only for releasing the angular momentum of the wheel 3, so that a large attitude control torque that can be generated by the thruster 12 is used. There is a problem that even if the thruster 12 is used, the posture can be changed only at the same speed as when only the wheel 3 is used.

  Furthermore, since the conventional satellite attitude control device is based on the assumption that the output of the attitude sensor 15 is used as the initial value of the attitude angle of the attitude estimation units 20 and 21 in the steady attitude, it is operated from the attitude change to the steady attitude. Immediately after the form is switched, the estimated value of the posture angle changes discontinuously. Depending on the magnitude of this change, the command value of the attitude control system changes greatly. In some cases, it takes some time to stabilize the attitude after starting the steady attitude, such as exciting the structural vibration mode. I won't. On the other hand, when the posture and the posture angular velocity are estimated with the same sensor configuration as the posture change after the posture change, the output of the angular velocity sensor 17 in the high speed mode is used, so that sufficient accuracy cannot be obtained. As described above, there is a problem in that the posture cannot be set quickly and with high accuracy after the posture change.

  The present invention was made to solve the problems associated with the above-described conventional example, and suppresses excitation of structural vibration of an artificial satellite even when it is necessary to take a longer time for occurrence of attitude angular acceleration, Another object of the present invention is to obtain an attitude control device for an artificial satellite that can cope with a case where a non-negligible error occurs in the rigid attitude angle of the artificial satellite due to the presence of structural vibration.

  In addition, when changing the posture using both the thruster and the wheel, use the large output torque of the thruster to change the posture at high speed, and at the same time, use the continuous output torque of the wheel to quickly change the posture. The object is to obtain an attitude control device for an artificial satellite capable of setting the artificial satellite.

  Furthermore, even after switching to an operation mode that uses the output of the angular velocity sensor in the low-speed mode with high accuracy after performing the posture change, the posture angle can be quickly settled without the estimated value of the posture angle becoming discontinuous, The purpose is to obtain an attitude control device for an artificial satellite capable of controlling the direction of a mission device such as an antenna or an observation sensor having a directivity control mechanism with high accuracy independently of the main body of the artificial satellite.

  The attitude control device for an artificial satellite according to the present invention is an attitude control device for changing the attitude of an artificial satellite at high speed, and the inertial force accompanying the attitude change of the artificial satellite as an attitude control means for controlling the attitude of the artificial satellite body. The attitude control feedforward arithmetic unit that obtains feedforward control torque based on the output from the attitude angle target value generation unit to compensate for it, converts it into a thruster drive signal and outputs it, and performs attitude control of the artificial satellite with external force A thruster used as an actuator to control the attitude of the artificial satellite based on the thruster drive signal from the attitude control feedforward calculation unit, and the attitude angle from the attitude angle target value generation unit by feeding back the attitude angle of the artificial satellite Based on the posture angle feedback output and the posture angle target value to reduce the error from the angle target value. It is used as an attitude control feedback calculation unit that obtains a back attitude control torque and converts it into a wheel drive signal and outputs it, and an actuator that performs the attitude control of the artificial satellite with an internal force, based on the wheel drive signal from the attitude control feedback calculation unit And a wheel for controlling the attitude of the artificial satellite.

  According to the present invention, the attitude control feedforward calculation unit that compensates the inertial force accompanying the attitude change of the artificial satellite, and the attitude control feedback calculation that reduces the error of the attitude angle target value by feeding back the attitude angle of the artificial satellite A thruster is used for the actuator of the attitude control feedforward calculation unit and a wheel is used for the actuator of the attitude control feedback calculation unit, so the driving torque required for high-speed attitude change is mainly high torque output. The thruster does not interfere with the wheel used for attitude control feedback because the thruster does not perform feedback control, and it is possible to change the attitude faster than using only the wheel. The effect of being able to quickly set the satellite There is.

Embodiment 1 FIG.
An attitude control device for an artificial satellite according to Embodiment 1 of the present invention will be described with reference to FIGS.
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, and FIG. 2 shows an attitude of the attitude control device for an artificial satellite according to Embodiment 1 of the present invention. It is a figure explaining operation | movement of an angle target value generation part. In addition, in each figure, the same code | symbol shows the same or equivalent part.

  In FIG. 1, 1 is an artificial satellite body, 3 is a wheel which is an actuator for controlling the attitude of the artificial satellite body 1 by a reaction torque of the rotor, that is, an internal force, and 7 is an artificial satellite based on a target value for changing the attitude angle and an actual attitude angle. An attitude control feedback calculation unit 8 that performs calculations necessary for attitude control of the vehicle, and 8 indicates an attitude control feedforward calculation unit that obtains in advance a wheel drive torque necessary for attitude change from a target value of attitude angle change. Attitude control that performs attitude control of the satellite body 1 based on target values of attitude angle, attitude angular velocity, and attitude angular acceleration by the configuration of the attitude control feedback calculator 7, attitude control feedforward calculator 8, and attitude estimation calculator 14. Means.

  Further, 9 is a dead beat pattern generation unit for obtaining an angular acceleration target value of the posture angle change at the final stage so as to suppress excitation of structural vibration at the end of the posture angle change, and 10 is a posture angle, posture angular velocity in the posture angle change, It is a posture angle target value generation unit for obtaining a target value of posture angular acceleration.

Next, the operation of the first embodiment will be described.
When the artificial satellite body 1 changes the attitude angle, a trapezoidal velocity pattern with respect to the attitude angular velocity is usually used. This is to increase the posture angular velocity at a certain angular acceleration, then change the posture angle at a constant angular velocity with zero acceleration, and finally change the posture angle when the angular velocity becomes zero by decelerating at a constant angular acceleration. Is to finish. Since the temporal change pattern of the angular velocity at this time becomes a trapezoid, it is called a trapezoidal velocity pattern. However, when the target value for changing the attitude angle is generated in this pattern, if there is a structural vibration that cannot be ignored in the satellite body 1, the structural vibration is also excited when the attitude angle change is completed. This will be described below.

Considering the case where there is only one main structural vibration mode of the artificial satellite body 1, the equation of motion can be simplified as the following equation (1).
d ² x / dt ² + 2ζωdx / dt + ω ² x = f (1)
Where x is the vibration amplitude of the main vibration mode, d ² x / dt ² is the second-order time derivative of x, dx / dt is the first-order time derivative of x, ω is the natural frequency, and ω ² is the natural frequency. The square of ω, ζ is a damping coefficient, and f is an inertial force acting on a main vibration mode when the satellite body 1 changes its attitude at a constant angular acceleration.

Consider a case where the posture angular acceleration changes from one constant value to another constant value. Accordingly, the inertial force f is also changed from f ₁ to f ₂ .
Assuming that the structural vibration attenuation coefficient ζ is large to some extent due to the action of the posture control system, it may be considered that the vibration amplitude x of the structural vibration is substantially constant immediately before the posture angular acceleration changes. The value of x ₀ when the value and x ₀ is determined from the above equation as follows.
x ₀ = f ₁ / ω ² (2)

Now, when the inertial force f changes from f ₁ to f ₂ from the state where the vibration amplitude x of the structural vibration is the above value and takes a constant value, the time of change is set as time t = 0 and the value x of the subsequent x (T) is obtained from equation (1) as follows.
x (t) = f ₂ / ω ² − [exp (−ζωt) {ω (1-ζ ² ) cos (Ωt) + ζΩsin (Ωt)} / ω ³ / (1-ζ ² )] × (f ₂ − f ₁ )
(3)
However, Ω = sqrt (1-ζ ² ) ω, exp is a power of the base e of the natural logarithm, and sqrt is a square root.

For example, considering the case where the posture angular acceleration changes from a certain value to zero, the inertial force accompanying the posture angular acceleration becomes zero after the change, and therefore f ₂ is zero in equation (3). As is clear from Equation (3), x (t) is not zero. That is, in the final state of posture change, in the case of a trapezoidal velocity pattern in which the posture angular acceleration is reduced from a certain value to zero, when the posture angle change is finished, the structural vibration does not become zero and the vibration continues. Will be. If this vibration continues, the attitude angle of the artificial satellite body 1 becomes oscillating, which makes it difficult to achieve high attitude accuracy.

Therefore, the dead beat pattern generation unit 9 generates the target value of the posture angular acceleration at the final stage as follows.
First, when the time derivative of equation (3) is taken to be t = π / Ω, it can be seen that the value of the time derivative dx / dt becomes zero. That is, after a half cycle of vibration, the time derivative of the vibration amplitude of the structural vibration becomes zero. Therefore, at this same time, if the inertial force f ₂ during the half cycle is set so that the value of x is also zero, and the inertial force thereafter is zero, the inertial force becomes zero. Since the vibration amplitude x of the structural vibration and its time differential dx / dt are both zero at the time, a deadbeat state in which no structural vibration occurs is achieved thereafter. The value of the inertial force f _{2 that} satisfies this condition can be obtained from the equation (3) as follows.
f ₂ = exp (−ζωπ / Ω) / {1 + exp (−ζωπ / Ω)} × f ₁
(4)

Since the inertial force acting on the structural vibration is proportional to the value of posture angular acceleration, the value of posture angular acceleration is a period of π / Ω corresponding to a half cycle of vibration in the final stage of posture change. it is sufficient to decrease so as to satisfy the ratio of f ₂ and f _1.
The function of the dead beat pattern generation unit 9 is to obtain a rate of decrease in posture angular acceleration and a duration during which the decreased posture angular acceleration lasts.

  If the above value is obtained by the dead beat pattern generation unit 9, the posture angle target value generation unit 10 determines the posture from the angle and time to change the posture and the final angular acceleration pattern of the dead beat pattern generation unit 9. A change in time (target value) of angular acceleration for change is determined. This angular acceleration is obtained by slightly changing the angular acceleration of a normal trapezoidal velocity pattern. If the target value of angular acceleration is obtained, the target value of angular velocity and the target value of angle can be obtained by time integration.

The state at this time is shown in FIG.
In FIG. 2, the upper diagram is the target value of the posture angular acceleration, and the portion indicated by A is the change in angular acceleration obtained by the deadbeat pattern generation unit 9.
As described above, immediately before the end of the posture change, the angular acceleration is decreased according to the equation (4) for a period corresponding to a half cycle of the structural vibration.
When the target value of the posture angular acceleration is integrated for the first floor, the target value of the posture angular velocity is obtained as shown in the middle diagram. Further, when the first-order time integration is performed, a target value of the posture angle is obtained as shown in the lower diagram. Thus, the posture angle target value generating unit 10 generates target values for the posture angle, posture angular velocity, and posture angular acceleration.

  The attitude angle and attitude angular velocity target values of the satellite body 1 output from the attitude angle target value generator 10 are input to the attitude control feedback calculator 7 by taking the difference between the actual attitude angle and attitude angular velocity. . The attitude control feedback calculation unit 7 generates the driving torque of the wheel 3 so that this difference becomes zero. At the same time, the posture angle, posture angular velocity, and posture angular acceleration target values output from the posture angle target value generation unit 10 are input to the posture control feedforward calculation unit 8, and the driving torque of the wheel 3 necessary for posture change is obtained in advance. This is obtained and added to the output of the attitude control feedback calculation unit 7 to obtain the final driving torque of the wheel 3.

  In the first embodiment, the case where the wheel 3 is used for the attitude control of the artificial satellite body 1 is shown. However, the same effect can be obtained even when the attitude control is performed by the thruster. Further, in the dead beat pattern generation unit 9, the case where the period during which the angular acceleration decreases is half of the vibration period has been shown. However, the period is not necessarily limited to half of the vibration period. If the vibration amplitude and its time derivative are set to be zero, the same effect can be obtained.

  Therefore, according to the first embodiment, the posture angle target value is set so that the posture angular acceleration value is decreased stepwise immediately before the posture change is finished, based on the frequency and damping coefficient of the main structural vibration. Therefore, at the time when the attitude angle change is finished after giving a constant attitude angular acceleration to the satellite continuously to change the attitude angle, for example, the value of the attitude angular acceleration is set for a period of a half cycle of vibration. The main structural vibration of the satellite excited by the attitude angular acceleration accompanying the attitude angle change is reduced when the attitude angle change is completed. Even when it is necessary to suppress the occurrence of attitude angular acceleration for a long time, the excitation of structural vibration of the satellite is suppressed, and the presence of structural vibration ignores the rigid attitude angle of the satellite. Can There may also deal with the case, such as occurs with error.

Embodiment 2. FIG.
An artificial satellite attitude control apparatus according to Embodiment 2 of the present invention will be described with reference to FIG.
FIG. 3 is a block diagram showing the configuration of the attitude control device for an artificial satellite according to Embodiment 2 of the present invention.
In FIG. 3, the same parts as those of the first embodiment shown in FIG. As a new code, reference numeral 11 denotes an inverse characteristic filter unit having characteristics almost opposite to those of dynamics including structural vibration of the satellite body 1.

Next, an operation according to Embodiment 2 will be described.
When changing the attitude of an artificial satellite at a high speed, at the time when the attitude change is completed, both a bias error and a vibration error do not remain in the rigid attitude angle of the satellite body 1. It is necessary. Among them, the vibrational error can be removed by the function of the dead beat pattern generation unit 9 shown in FIG. However, it is not possible to remove the bias angle error of the posture angle alone, and further ingenuity is required. The reason is described below.

Now, assuming that there is only one main structural vibration mode of the satellite body 1, its vibration amplitude, natural frequency, and damping coefficient are x, ω, and ζ, respectively, and the rigid attitude angle of the satellite body 1 is If θ and the moment of inertia of the satellite body 1 are J, the equation of motion of the entire satellite can be considered as follows.
Jd ² θ / dt ² + Md ² x / dt ² = n
Md ² θ / dt ² + d ² x / dt ² + 2ζωdx / dt + ω ² x = 0 (5)
In this equation, n is the driving torque of the satellite body 1, M is a dynamic interference term between the rigid attitude angle θ and the vibration amplitude x, and d ² θ / dt ² is the second-order time derivative of θ.

Here, if the provisional target value of the attitude angle θ is θ ₂ , the driving torque n of the artificial satellite body 1 can be expressed as follows, for example.
n = Jd ² θ ₂ / dt ²
+ Kd (dθ ² / dt−dθ / dt) + kp (θ ₂ −θ) (6)
In the above equation, d ² θ ₂ / dt ² is the second-order time derivative of θ ₂ , dθ ₂ / dt is the first-order time derivative of θ ₂ , dθ / dt is the first-order time derivative of θ, and kd is the derivative of attitude control. The control gain, kp, is a proportional control gain for attitude control.

In the above equation, the first term on the right side corresponds to the action of the attitude control feedforward calculation unit 8, and the second term on the right side and the third term on the right side correspond to the action of the attitude control feedback calculation unit 7. Here, the driving torque n is selected so that the actual attitude angle θ of the artificial satellite body 1 is equal to the provisional target value θ ₂ , but an error occurs due to the presence of the vibration term x. When the driving torque n is substituted into the original equation of motion, the following equation is obtained.
Jd ² θ / dt ² + kddθ / dt + kpθ + Md ² x / dt ²
= Jd ² θ ₂ / dt ² + kddθ ₂ / dt + kpθ ₂
Md ² θ / dt ² + d ² x / dt ² + 2ζωdx / dt + ω ² x = 0 (7)

As apparent from the above equation (7), the posture angle θ does not become equal to the temporary target value θ ₂ due to the influence of the vibration amplitude x (the influence of the term Md ² x / dt ² ).
To remove this bias error, consider the inverse system of equation (7). Substituting the output θ ₁ of the attitude angle target value generation unit 10 in place of θ in Equation (7) and moving θ ₂ to the left side, the following system is obtained.
Jd ² θ2 / dt ² + kddθ ₂ / dt + kpθ ₂ −Md ² x / dt ²
= Jd ² θ ₁ / dt ² + kddθ ₁ / dt + kpθ ₁
d ² x / dt ² + 2ζωdx / dt + ω ² x
= −Md ² θ ₁ / dt ² (8)

This is because θ and θ ₁ are interchanged in equation (7), and therefore the relationship between θ ₁ and θ ₂ is just the opposite of the relationship between θ ₂ and θ in equation (7). Therefore, if the output θ ₁ of the posture angle target value generation unit 10 is once converted into the temporary target value θ ₂ by the equation (8) and then the posture control is performed as in the equation (6), the equation (7) ) And Expression (8), the relationship between θ ₁ and θ _{2 and} the relationship between θ ₂ and θ are just canceled, and the actual posture angle θ becomes equal to the output θ ₁ of the posture angle target value generation unit 10. That is, the influence of the structural vibration included in the artificial satellite body 1 can be removed by inserting a filter having characteristics opposite to those of the artificial satellite body 1 in advance. The filter that inputs θ ₁ represented by the equation (8) and outputs θ ₂ is called an inverse characteristic filter unit 11 here.

  However, in Equation (8), when the vibration attenuation coefficient ζ is small, the filter becomes vibrational. In order to avoid this, the inverse characteristic filter unit 11 artificially gives a large attenuation. . For this reason, the reverse characteristics of the plant are not completely realized, but in that case, there is an effect of improving the response.

In this example, the case where the wheel is used for the attitude control of the artificial satellite body 1 is shown, but the same effect can be obtained even when the attitude control is performed by the thruster.
Further, when the deadbeat pattern generation unit 9 shown in FIG. 1 is inserted in front of the attitude angle target value generation unit 10, both a bias error and a vibration error in the rigid attitude angle of the artificial satellite body 1 are obtained. There is an effect to prevent it from remaining.

  Therefore, according to the second embodiment, the target value for changing the attitude angle given to the rigid attitude angle of the artificial satellite is set to the inverse characteristic having almost the opposite characteristic to the dynamics including the structural vibration of the artificial satellite. By passing the filter 11 to the target value for the final attitude angle change of the artificial satellite, the influence of the structural vibration of the artificial satellite is almost canceled, and the bias to the rigid attitude angle of the artificial satellite is caused by the influence of the structural vibration. Generation of a typical error can be prevented.

  Further, by inserting the dead beat pattern generation unit 9 shown in FIG. 1 before the attitude angle target value generation unit 10, bias errors and vibration errors in the rigid attitude angle of the artificial satellite body 1 can be obtained. It has the effect of not leaving both.

Embodiment 3 FIG.
An artificial satellite attitude control apparatus according to Embodiment 3 of the present invention will be described with reference to FIG.
FIG. 4 is a block diagram showing the configuration of the attitude control device for an artificial satellite according to Embodiment 3 of the present invention.
4, the same parts as those in the first and second embodiments shown in FIGS. 1 and 2 are denoted by the same reference numerals, and the description thereof is omitted. As a new code, reference numeral 12 denotes a thruster that is an actuator that performs the attitude control of the satellite body 1 by gas injection, that is, by external force.

Next, the operation of the third embodiment will be described.
When changing the attitude of the artificial satellite body 1 at high speed, assuming that the attitude angle of the artificial satellite body 1 is θ and the attitude angle output of the attitude angle target value generation unit 10 is θ ₂ , To give the attitude control torque n of the satellite body 1.
n = Jd ² θ ₂ / dt ² + gyro term + kd (dθ ₂ / dt−dθ / dt)
+ Kp (θ ₂ −θ) (9)

In this equation (9), the first term on the right side, Jd ² θ ₂ / dt ² , is the attitude angular acceleration output d ² θ ₂ / dt ² of the attitude angle target value generation unit 10 to the moment of inertia of the artificial satellite body 1. J can be obtained without feedback of the attitude angle θ of the artificial satellite body 1. Further, the gyro term of the second term on the right side is between the angular momentum of the entire satellite 1 (the sum of the angular momentum of the satellite body 1 and the angular momentum of the wheel 3) when the satellite body 1 changes its attitude at high speed. In the case where the thruster 12 is used as an actuator, the angular momentum of the entire satellite changes, and this term cannot be ignored. This gyro term can also be obtained by using the attitude angle output of the attitude angle target value generation unit 10 without feedback of the attitude angle θ of the artificial satellite body 1.

  The calculation of the first term on the right side and the second term on the right side in these equations (9) is the function of the attitude control feedforward calculation unit 8, and the obtained feedforward attitude control torque is artificial during the attitude change. It occupies the main part of the attitude control torque of the satellite body 1 and enables the attitude change of the artificial satellite body 1 at high speed. The obtained feedforward attitude control torque is further converted into a thruster drive signal by the attitude control feedforward calculation unit 8 to drive the thruster 12.

On the other hand, the third term on the right side and the fourth term on the right side in Equation (9) are terms obtained by feeding back the attitude angle θ and the attitude angular velocity dθ / dt of the artificial satellite body 1. It works to reduce the error between the output θ ₂ and the attitude angle θ of the artificial satellite body 1. The attitude control feedback calculation unit 7 obtains this feedback attitude control torque, converts it into a wheel drive signal, and drives the wheel 3.

  As described above, when the attitude control torque by the attitude control feedforward calculation unit 8 is realized by the thruster 12 and the attitude control torque by the attitude control feedback calculation unit 7 is realized by the wheel 3, the feedback is transmitted to the wheel 3. Therefore, the thrust control system of the thruster 12 and the wheel 3 does not interfere with each other, and a large part of the large posture control torque required during the posture change is the posture control feedforward calculation unit 8. Therefore, it is possible to realize a high-speed posture change by effectively utilizing the high torque output of the thruster 12. Further, after the posture change, the output of the posture control feedforward calculation unit 8 becomes zero, so that the thruster 12 is not injected, and the posture can be quickly settled by the continuous output torque of the wheel 3. It becomes possible.

  In this example, the case where the gyro term of the feedforward attitude control torque is compensated by the thruster 12 is shown. However, since the gyro term can also be controlled by the normal wheel 3, the gyro term is adjusted by the wheel 3 in that case. Even if it is included in the control torque, the same effect can be obtained. Further, the dead beat pattern generation unit 9 shown in FIG. 1 is inserted before the posture angle target value generation unit 10, and the inverse characteristic filter unit 11 shown in FIG. 3 is inserted after the posture angle target value generation unit 10. These are all effective and effective in removing the influence of structural vibration on the attitude angle of the satellite body 1.

  Therefore, according to the third embodiment, the thruster 12 is used as the actuator of the attitude control feedforward calculation unit 8 and the wheel 3 is used as the actuator of the attitude control feedback calculation unit 7, so that the high speed of the artificial satellite is increased. The large control torque required to change the attitude of the satellite is mainly used to compensate the inertial force of the artificial satellite. Therefore, this is obtained by the attitude control feedforward calculation unit 8 and the necessary torque is generated by the thruster 12. Thus, it is possible to change the posture at a higher speed than using only the wheel 3 without interfering with the wheel 3 used for posture control feedback.

Embodiment 4 FIG.
An artificial satellite attitude control apparatus according to Embodiment 4 of the present invention will be described with reference to FIG.
FIG. 5 is a block diagram showing the configuration of the attitude estimation calculation unit of the attitude control device for an artificial satellite according to Embodiment 4 of the present invention.
In FIG. 5, the same parts as those in the conventional example shown in FIG. As new codes, 14A represents a posture estimation calculation unit according to the fourth embodiment, 22 represents a posture estimation time estimation unit, and the posture estimation calculation unit 14A according to the fourth embodiment includes a posture estimation time posture estimation. A built-in unit 22 is configured to output the estimated values of the posture angle and posture angular velocity at the time of posture setting to the posture control feedback calculation unit 7 at the time of posture setting in which the posture is quickly set after changing the posture.

Next, the operation of the fourth embodiment will be described.
At the time of posture setting for quickly setting the posture after the posture change, the posture estimation calculation unit 14 outputs the estimated values of the posture angle and posture angular velocity from the posture estimation unit 22 during posture setting to the posture control feedback calculation unit 7. . Since the attitude change is finished, the assumed attitude angular velocity of the artificial satellite is small, and the angular velocity sensor 16 in the low speed mode can be used. A bias-like error included in the output of the angular velocity sensor 16 is corrected using the output of the bias estimation unit 18 to obtain the estimated value α of the posture angular velocity. By adding an initial value of the posture angle to this, an estimated value θ of the posture angle is obtained. As the initial value of the posture angle, the final value of the posture angle of the posture estimation unit 21 at the time of posture change before switching the operation mode is used. In the bias estimation unit 18, the estimated value of the posture angle of the posture estimation unit 20 in the steady posture and the output of the posture sensor 15 are compared, and a bias-like error included in the output of the angular velocity sensor 16 is estimated.

  Note that the angular velocity sensor 16 in the low speed mode and the angular velocity sensor 17 in the high speed mode are the same angular velocity sensor and the angular velocity sensor in the narrow measurement range and the low resolution even when the resolution and the measurement range are switched by switching the sensor gain. It goes without saying that the same effect can be obtained even when an angular velocity sensor with a wide measurement range is combined.

  Further, in order to correct a bias-like error included in the output of the angular velocity sensor 16, the bias estimation unit 18 performs calibration obtained by processing the data acquired in the ground test or the previous operation on the orbit offline. The same effect can be obtained when using data.

  Although only the posture estimation calculation unit 14A has been described in the fourth embodiment, it can be applied to the third embodiment shown in FIG. Inserting the dead beat pattern generating unit 9 shown in FIG. 1 and inserting the inverse characteristic filter unit 11 shown in FIG. 3 after the attitude angle target value generating unit 10 are both effective. It is effective in removing the influence of structural vibration on the attitude angle of

  Therefore, according to the fourth embodiment, the posture estimation unit 20 in the normal posture that estimates the posture angle and posture angular velocity in the normal posture, and the posture estimation unit 21 in posture change that estimates the posture angle and posture angular velocity in the posture change. In addition to the above, a posture estimating unit 22 for estimating the posture angle and posture angular velocity at the time of posture setting is provided, and the posture of the posture changing unit 21 at the time of posture change is initialized by the posture estimating unit 22 at the time of posture setting. By using the final value of the angle and integrating the output of the angular velocity sensor 16 in the low-speed mode to estimate the posture, the posture angle estimated value is continuously obtained during operation switching from posture change to posture settling. It is possible to estimate the change in posture angle with high accuracy.

Embodiment 5. FIG.
An artificial satellite attitude control apparatus according to Embodiment 5 of the present invention will be described with reference to FIG.
FIG. 6 is a block diagram showing the configuration of the attitude estimation calculation unit of the attitude control device for an artificial satellite according to Embodiment 5 of the present invention.
In FIG. 6, the same parts as those of the fifth embodiment shown in FIG. As a new code, reference numeral 23 denotes a directional control device that performs directional control of an antenna, an observation sensor, or the like that has a directional control mechanism.

Next, the operation of the fifth embodiment will be described.
When the directivity direction of the antenna or observation sensor can be controlled electrically or by driving a light active member such as a small mirror and practically no reaction torque is exerted on the satellite body 1 as a result of the control Even if the estimated value of the attitude angle of the artificial satellite body 1 is discontinuous, there is no fear of exciting the structural vibration of the artificial satellite by the directivity control. Therefore, when switching from the posture change to the posture settling, the calculation of the posture estimation unit 20 in the steady posture is executed in parallel with the calculation of the posture estimation unit 22 in the posture settling, and the highly accurate posture determination value is transmitted to the pointing control device. To 23. The directional control device 23 can perform directional control based on this highly accurate posture determination value.

  In the fifth embodiment, only the posture estimation calculation unit 14A and the directivity control device 23 have been described. However, the posture angle target value generation unit can be applied to the third embodiment shown in FIG. It is effective to insert the dead beat pattern generation unit 9 shown in FIG. 1 before the stage 10 and the inverse characteristic filter unit 11 shown in FIG. 3 after the attitude angle target value generation part 10. This is effective in removing the influence of structural vibration on the attitude angle of the satellite body 1.

  Therefore, according to the fifth embodiment, at the time of posture setting by posture control feedback based on the output of the posture estimation unit 22 at the time of posture setting, the directivity control of the antenna, the observation sensor, etc. based on the output of the posture estimation unit 20 at the normal posture. As soon as the attitude change is completed, the attitude change of the satellite body 1 is controlled with high accuracy and the direction of the mission equipment such as the antenna and the observation sensor is controlled with high accuracy. There is an effect that becomes possible.

  As described above, according to the attitude control device for an artificial satellite according to the present invention, after providing a deadbeat pattern generation unit and continuously giving a constant attitude angular acceleration to the artificial satellite for changing the attitude angle, At the end of the posture angle change, for example, the posture angle acceleration value is decreased only for a half period of vibration, and then the posture angle acceleration value is set to zero to stop the posture angle change. The main structural vibration of the satellite excited by the accompanying attitude angular acceleration can be suppressed at the end of the attitude angle change, and even if it is necessary to take a longer time for the attitude angular acceleration to occur, There is an effect that it is possible to cope with a case where an error that cannot be ignored is generated in the rigid attitude angle of the artificial satellite due to the existence of the structural vibration, while suppressing the excitation of the structural vibration.

  In addition, the target value of the attitude angle change given to the rigid attitude angle of the satellite is passed through an inverse characteristic filter unit having almost the opposite characteristics to the dynamics including the structural vibration of the satellite, and Because the target value for the final attitude angle change is set, the influence of the structural vibration of the satellite is almost cancelled, and the occurrence of a bias error in the rigid attitude angle of the artificial satellite due to the influence of the structural vibration is prevented. There is an effect that can be.

  In addition, it includes an attitude control feedforward calculation unit that compensates for the inertial force accompanying the attitude change of the satellite, and an attitude control feedback calculation unit that feeds back the attitude angle of the satellite and reduces the error from the target value of the attitude angle. Since a thruster is used for the actuator of the attitude control feedforward calculation unit and a wheel is used for the actuator of the attitude control feedback calculation unit, the driving torque required for high-speed attitude change is mainly by a thruster capable of high torque output. Since the thruster does not perform feedback control, it does not interfere with the wheel used for attitude control feedback, and the attitude can be changed at a higher speed than when only the wheel is used. The effect is that it can be set.

  In addition to the posture estimator for steady posture and the posture estimator for posture change for estimating posture angle and velocity at posture change, the posture angle for posture settling A posture estimation unit for posture setting that estimates posture angular velocity, and the posture estimation unit for posture setting uses the final value of the posture angle of the posture change unit for posture change to initialize the posture angle, and the angular velocity sensor of the low-speed mode Since the posture is estimated by integrating the output, it is possible to estimate the posture angle change with high accuracy while maintaining the posture angle estimated value continuously during operation switching from posture change to posture settling. There is an effect that it becomes possible.

  In addition, during posture setting by posture control feedback based on the output of the posture estimation unit at the time of posture setting, directivity control of devices having a pointing control mechanism such as an antenna or an observation sensor is performed based on the output of the posture estimation unit at the time of steady posture As a result, it is possible to control the attitude change of the satellite body with high accuracy and to control the pointing direction of mission equipment such as antennas and observation sensors with high accuracy immediately after the attitude change is completed. is there.

It is a block diagram which shows the structure of the attitude | position control apparatus of the artificial satellite which concerns on Embodiment 1 of this invention. It is explanatory drawing which shows operation | movement of the attitude | position angle target value generation part of the attitude | position control apparatus of the artificial satellite which concerns on Embodiment 1 of this invention. It is a block diagram which shows the structure of the attitude | position control apparatus of the artificial satellite which concerns on Embodiment 2 of this invention. It is a block diagram which shows the structure of the attitude | position control apparatus of the artificial satellite which concerns on Embodiment 3 of this invention. It is a block diagram which shows the structure of the attitude | position estimation calculating part of the attitude | position control apparatus of the artificial satellite which concerns on Embodiment 4 of this invention. It is a block diagram which shows the structure of the attitude | position estimation calculating part of the attitude | position control apparatus of the artificial satellite which concerns on Embodiment 5 of this invention. It is explanatory drawing which shows operation | movement of the attitude control apparatus of the artificial satellite using the conventional thruster. It is a block diagram which shows the structure of the attitude control apparatus of the artificial satellite which uses the conventional wheel and thruster together. It is a block diagram which shows the structure of the attitude | position estimation calculating part of the artificial satellite which uses a conventional attitude angle sensor and an angular velocity sensor together.

Explanation of symbols

  DESCRIPTION OF SYMBOLS 1 Satellite body, 3 wheel, 7 attitude control feedback calculation part, 8 attitude control feedforward calculation part, 9 deadbeat pattern generation part, 10 attitude angle target value generation part, 11 reverse characteristic filter part, 12 thruster, 14A attitude estimation Calculation unit, 15 posture sensor, 16 angular velocity sensor (low speed mode), 17 angular velocity sensor (high speed mode), 18 bias estimation unit (low speed mode), 19 bias estimation unit (high speed mode), 20 posture estimation unit during steady posture, 21 Posture change posture estimation unit, 22 posture settling posture estimation unit, 23 pointing control device.

Claims (1)

In the satellite attitude control device that changes the attitude of the satellite at high speed,
As attitude control means for attitude control of the satellite body,
An attitude control feedforward arithmetic unit that obtains a feedforward control torque based on an output from the attitude angle target value generation unit to compensate for an inertial force accompanying an attitude change of the artificial satellite, converts it into a thruster drive signal, and outputs it;
A thruster that is used as an actuator for performing attitude control of the artificial satellite with external force, and that performs attitude control of the artificial satellite based on a thruster drive signal from the attitude control feedforward calculation unit;
Feedback attitude control torque based on attitude angle feedback output and attitude angle target value to reduce attitude error from attitude angle target value from attitude angle target value generator by feedback of attitude angle of artificial satellite An attitude control feedback calculation unit that calculates and outputs a wheel drive signal,
A posture control of the artificial satellite comprising: a wheel that is used as an actuator for performing the posture control of the artificial satellite with an internal force, and that performs a posture control of the artificial satellite based on a wheel drive signal from the posture control feedback calculation unit. apparatus.

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