JP4509006B2 - Satellite attitude determination system - Google Patents

Description

  The present invention relates to a configuration of an attitude sensor and attitude determination system for attitude control of an artificial satellite.

  In general, a spacecraft such as an artificial satellite detects or estimates a posture angle and a posture angular velocity by combining a posture sensor and an inertial navigation device. The artificial satellite compares the detected or estimated attitude angle and attitude angular velocity with a target value, and calculates the difference so that the difference becomes zero, thereby performing attitude control. The function of detecting or estimating the attitude angle and attitude angular velocity of the artificial satellite itself is called attitude determination.

  Conventionally, a low-frequency attitude determination step for detecting the satellite angular velocity by an inertial reference device using a star sensor and a gyro, and a step for determining a broadband attitude using a high-frequency sensor, a low-pass filter, and a feedback-type attitude detection filter In the low frequency range, the entire system is made to follow the output from the low frequency attitude determination unit composed of the star sensor and the inertial reference device, and in the higher frequency range, from the high frequency attitude determination unit including the high frequency sensor. A wide-band attitude control method for artificial satellites has been proposed in which disturbances due to overlapping gain characteristics of sensors are automatically flattened by feedback that causes the entire system to follow the output. (For example, see Patent Document 1)

  In this prior art, the output from the low-frequency star sensor and gyro is always taken into the attitude determination system as a true value, and the attitude control of the artificial satellite is performed. However, the gyro output has a sampling frequency as low as a few dozen Hz, and many of the characteristics are secondary low-pass characteristics, so the frequency of minute vibrations (hereinafter referred to as disturbances) caused by satellite mounted momentum wheels and other devices. When the frequency is high and the amplitude is large, a disturbance component that is equal to or higher than the sampling frequency is actually changed to a low frequency due to a phenomenon called aliasing, and high-frequency disturbance is mixed in the output of the inertial reference device. For this reason, an error occurs in the posture estimation angle and the posture estimation angular velocity calculated by the low-frequency posture determination unit by the inertial reference device, so that there is a problem in performing high-bandwidth and high-precision control essentially.

JP 2000-142594 A

  The problem with the prior art is that the posture sensor is assumed to output accurate values for the low frequency band and the high frequency band, respectively. Since conventional satellites have low directivity accuracy requirements for observation equipment and antennas, the attitude sensor captures the effects of minute vibrations associated with antenna drive, reaction wheels, momentum wheels, and disturbances generated from driving objects inside the observation equipment. Even if it was included in the data, the attitude control of the satellite and the pointing accuracy of the observation equipment were sufficiently satisfied.

  However, in recent years, astronomical satellites and earth observation satellites have been required to have extremely high directivity stability for the onboard observation equipment, and the influence of microvibration due to disturbance of the onboard equipment that has not been a problem so far. It can no longer be ignored. The minute vibrations accompanying these disturbances also occur in the satellite structure panel where the gyro is installed. Therefore, the angular velocity (angle increment) detected by the gyro includes a true satellite attitude variation and a rotation component of the mounting panel caused by the disturbance of the driving object.

  Most of the frequency components of the disturbance are about the control frequency band of the attitude control system, that is, about 10 or more Hz. Usually, the gyro is provided with a secondary low-pass filter so that a high frequency exceeding the band of attitude control is removed. However, normally, the internal disturbance of the artificial satellite has such an amplitude that this filter cannot sufficiently remove high-frequency components contained in the signal output from the gyro.

  For this reason, a component called aliasing occurs in the components that cannot be removed by this filter, and changes to low-frequency disturbances and data is taken into the posture determination system, and posture variations of frequency components that do not actually exist are detected. Will be. When these data are taken into the posture determination system, incorrect posture angles and posture angular velocities are estimated, and when performing wideband posture control using these estimated angles, errors may occur in posture control. It becomes.

  The present invention has been made to cope with such problems, and removes the high-frequency vibration component applied to the attitude control system as described above, and outputs an attitude sensor and attitude determination that outputs only true satellite attitude fluctuations. The purpose is to provide a system.

The attitude determination system for an artificial satellite according to the present invention is for removing a high-frequency disturbance component in which a filter parameter is variable according to a frequency component of the high-frequency disturbance between a rebalance loop inside the gyro serving as an attitude sensor of the artificial satellite and an integrator. A gyro detection system having a filter, a high-frequency disturbance measurement sensor, a broadband detection system having a spectrum processing unit for detecting a high-frequency disturbance component from the sensor signal, and a disturbance component estimated from the broadband detection system A parameter determining unit that determines a filter parameter of the high-frequency disturbance component removing filter .

  Since the attitude determination system of the satellite according to the present invention is configured as described above, the attitude change of the satellite is accurately detected with a gyro in a band up to a low frequency (up to 10 Hz), which is a problem in attitude control. It becomes possible to do.

  The reason for this is that, with a normal configuration, the rotation due to the attitude variation of the satellite and the rotation angle caused by the resonance of the mounting panel are taken into the gyro without distinction. By installing a rotational angular velocity detection sensor that can measure wide-range disturbances up to the frequency range, internal disturbances can be detected regardless of before and after launch, the gyro output filter characteristics can be changed, or attitude determination system sampling can be performed. Or by providing a buffer between the gyro mounting panel and the gyro so as to remove high-frequency disturbances, it is possible to accurately sense only the attitude variation of the artificial satellite.

  In addition, accuracy of attitude control is improved by sensing attitude fluctuation with high accuracy, and it becomes possible to increase the resolution of observation images of the Earth observation satellite and the astronomical satellite by improving the attitude control accuracy.

Embodiment 1 FIG.
Embodiment 1 of the present invention will be described below with reference to the drawings. FIG. 1 is a schematic diagram showing the configuration of the attitude sensor according to the first embodiment, and more specifically, is a block diagram showing the internal configuration of the gyro 12 (FIG. 3) that is the attitude sensor. The sensor shown in FIG. 1 is provided with a high-frequency disturbance component removing filter 3 between a rebalance loop 2 and an integrator 4, which is a general gyro configuration.

  In FIG. 1, the gyro detection system 1 inputs an angle generated in the satellite structure 13 (FIG. 3), differentiates it by a differentiator 20, detects an angular velocity, and then inputs it to the rebalance loop 2. When the rebalance loop 2 has a frequency gain characteristic in FIG. 12A and a frequency phase characteristic in FIG. 12B, the frequency characteristic is a second-order low-pass filter. In this case, the gyro output is not affected by the high frequency disturbance.

However, when a momentum wheel or reaction wheel is operated, high-frequency disturbance is dominant as shown in FIG. 11 along with the rotational speed, and the frequency component also changes depending on the rotational speed. Further, when the level of the disturbance is high, the high-frequency component is affected by aliasing as shown in FIG. 10, and the actual disturbance frequency is changed to the low-frequency component and is detected. A value different from the posture variation is taken into the posture determination system.
Therefore, a high frequency disturbance component removing filter 3 is provided to output an analog output angular velocity. This angular velocity is further incorporated into the attitude determination system 5 through the integrator 4 as an angular increment. Usually, the sampling frequency of this attitude determination system 5 is set to 10 and several Hz.

  By adopting such a configuration, the influence of the high frequency component as shown in FIG. 11 is removed by the high frequency disturbance component removing filter 3, and the true low frequency satellite attitude angular velocity can be detected as shown in FIG. I can do it. Furthermore, since the true low-frequency satellite attitude angular velocity is estimated, an error in data taken into the attitude determination system 5 is reduced, so that attitude control with high accuracy is possible.

Embodiment 2. FIG.
It will now be described a second embodiment of the present invention. Since the configuration of the attitude sensor is the same as that of the first embodiment, illustration and description thereof are omitted. In the first embodiment, the high-frequency disturbance component removing filter 3 is provided as shown in FIG. 1, but in the second embodiment, this filter is variable, and the filter parameters can be changed as necessary. It is what I did.

  With this configuration, even when the frequency of the high-frequency disturbance component determined before launch based on ground test data changes on the orbit after launch, the parameters can be set appropriately. It becomes possible to remove the influence of high-frequency disturbance components.

Embodiment 3 FIG.
Next, a third embodiment of the present invention will be described. Since the configuration of the attitude sensor is the same as that of the first embodiment, illustration and description thereof are omitted. In the first embodiment, the high-frequency disturbance component removing filter 3 is provided as shown in FIG. 1, but in the third embodiment, the parameters of the high-frequency disturbance component removing filter 3 are set to various sensor data mounted on the satellite. The disturbance component is estimated and determined from the above. However, these sensors are not limited to sensors installed for detecting high-frequency disturbance components.

  By adopting such a configuration, it is not necessary to install a sensor for detecting high-frequency disturbance components, so that the configuration of the sensor can be greatly simplified.

Embodiment 4 FIG.
Next, a fourth embodiment of the present invention will be described with reference to the drawings. FIG. 2 is a schematic diagram showing the configuration of the attitude sensor according to the fourth embodiment. Specifically, FIG. 2 is a block diagram showing the internal configuration of the attitude sensor and the gyro 12. FIG. 3 is a schematic diagram showing the installation position relationship between the gyro 12 and the satellite structure 13, and the acceleration sensor 11 is provided in contact with the outside of the gyro 12.

The posture sensor shown in FIG. 2 includes a gyro detection system 1 indicated by a double broken line, an attitude determination system 5 indicated by a double solid line, and a broadband detection system 10 indicated by a broken line. Since the gyro detection system 1 is the same as that of the first embodiment, the corresponding parts are denoted by the same reference numerals and description thereof is omitted.
The broadband detection system 10 is provided for detecting the frequency component and the magnitude of the rotation due to the high-frequency panel resonance or the like that becomes noise for the gyro detection system 1 that detects the low-frequency attitude variation of the artificial satellite. It is.

  As shown in FIG. 4, the broadband detection system 10 has an acceleration sensor 11 attached to the housing of the gyro 12 so as to measure a disturbance component, specifically acceleration. The angular acceleration is calculated from the measured acceleration as follows.

  Further, the angular acceleration data is first-order integrated to calculate a rotational angular velocity including a high-frequency disturbance component generated on the mounting surface. Next, angular velocity data as shown in FIG. 11 is obtained by the angular velocity spectrum processing unit 8. Next, a high-frequency disturbance component included in the acceleration measured by the high-frequency disturbance component detection unit 9 is detected.

  In order to prevent the high-frequency disturbance component detected here from adversely affecting the output of the gyro detection system 1, the high-frequency disturbance component removal filter parameter determination unit 6 determines the parameters of the high-frequency disturbance component removal low-pass filter 3, specifically, Is the cut-off frequency and order, and this parameter is input to the high-frequency disturbance component removing low-pass filter 3 to determine the characteristics.

  At this time, the filter parameter to be set is set so as to be small enough that the influence of the high-frequency disturbance component as shown in FIG. 11 can be sufficiently ignored. Next, the filter setting method in the filter parameter determination unit 6 for removing high-frequency disturbance components will be specifically described. Here, a method for avoiding the influence of disturbance by the reaction wheel will be described as an example.

  First, an angular velocity spectrum that envelops the spectrum of the entire rotation speed range of the wheel is calculated from the data of FIG. 11, and data as shown in FIG. 5 is obtained as a result. In FIG. 5, the portion indicated by a circle is the disturbance component. However, these high-frequency disturbance components actually affect the gyro output after the high-frequency disturbance components pass through the low-pass filter inside the rebalance loop 2, and the value is This is obtained by multiplying the frequency characteristic shown in FIG. 5 by the gain obtained from FIG. 12A, for example, as shown in FIG. The filter is designed so that the response in FIG. 6 is about Ωmin or less for 1 bit of the gyro rate.

  In addition, although the case where the gyro rate used as the reference of the filter characteristic is 1 bit or less has been illustrated, this value varies depending on the directivity accuracy requirement of the artificial satellite. Therefore, the prescribed value may be determined so as to satisfy each specification. Here, the determined filter parameters are input to the high-frequency disturbance component removal filter 3 inside the gyro detection system.

  By adopting such a configuration, the influence of the high frequency component as shown in FIG. 10 is removed by the high frequency disturbance component removing filter 3, and the true low frequency satellite attitude angular velocity can be detected as shown in FIG. I can do it. Further, since the true low-frequency satellite attitude angular velocity is estimated, an error in data taken into the attitude determination system 5 is reduced, so that attitude control with high accuracy is possible.

Embodiment 5 FIG.
Next, a fifth embodiment of the present invention will be described. Since the configuration of the attitude sensor is the same as that of the fourth embodiment, illustration and description thereof are omitted. In the fifth embodiment, the high-frequency disturbance component removing low-pass filter 3 in FIG. 2 is provided with a notch filter that can remove the limited frequency component when the disturbance frequency is limited to a single one. Is.

  With such a configuration, it is possible to reduce the phase lag of the posture angular velocity output, and it is possible to improve the posture control with high accuracy.

Embodiment 6 FIG.
Next, a sixth embodiment of the present invention will be described. Since the configuration of the attitude sensor is the same as that of the fourth embodiment, illustration and description thereof are omitted. In the fourth embodiment, the four acceleration sensors 11 are installed in the gyro 12 as the broadband detection sensor, as shown in FIGS. 3 and 4, but the sixth embodiment has three acceleration sensors. 11 is installed in the gyro 12 at intervals of 120 degrees, and the biaxial component excluding the in-plane rotational component is calculated. In addition, any arrangement may be used as long as the in-plane biaxial angle variation of the structure panel surface 13 of the artificial satellite on which the gyro is installed can be calculated.

  By adopting such a configuration, the amount of data to be processed by the wideband detection system is reduced, so that it is possible to shorten the time until the parameter determination of the high frequency disturbance removal filter.

Embodiment 7 FIG.
Next, a seventh embodiment of the present invention will be described. Since the configuration of the attitude sensor is the same as that of the fourth embodiment, illustration and description thereof are omitted. In the fourth embodiment, as shown in FIG. 3, the acceleration sensor 11 is provided outside the housing of the gyro 12, but in the seventh embodiment, the acceleration sensor 11 is mounted inside the gyro 12.

  With such a configuration, it is possible to effectively use the space because there is no need to route the cable.

Embodiment 8 FIG.
Next, an eighth embodiment of the present invention will be described. Since the configuration of the attitude sensor is the same as that of the fourth embodiment, illustration and description thereof are omitted. In the fourth embodiment, as shown in FIGS. 2 to 4, the acceleration sensor 11 is used as a sensor in the broadband detection system 10, but in the eighth embodiment, the acceleration sensor 11 is not used, and a broadband two-axis A jitter sensor is installed in the gyro 12 and used as a broadband detection sensor.

  By adopting such a configuration, it is possible to reduce the load of data processing because the conversion calculation from acceleration to angular velocity described above is omitted, and it is possible to simplify the configuration of data processing. Become.

Embodiment 9 FIG.
Next, a ninth embodiment of the present invention will be described. Since the configuration of the attitude sensor is the same as that of the fourth embodiment, illustration and description thereof are omitted. In the fourth embodiment, as shown in FIG. 4, the acceleration sensor 11 is arranged so as to contact the housing of the gyro 12. A panel (not shown) having a resonance frequency higher than the disturbance frequency is installed, and the acceleration sensor 11 and the gyro 12 are arranged on the panel.

  With such a configuration, it is possible to detect the acceleration input to the gyro as in the case where the acceleration sensor 11 is attached to the housing of the gyro 12.

Embodiment 10 FIG.
Next, an embodiment 10 of the invention will be described. Since the configuration of the attitude sensor is the same as that of the fourth embodiment, illustration and description thereof are omitted. In the fourth embodiment, in the filter setting process in the filter parameter determination unit 6 for removing the high-frequency disturbance component, as shown in FIG. 5, the angular velocity spectrum is created so as to envelope the spectrum of the entire rotation speed range of the wheel. In addition, although the entire range of the number of revolutions of the momentum wheel is enveloped, the tenth embodiment is not a disturbance source such as a momentum wheel or a reaction wheel that changes the number of revolutions and changes the disturbance frequency. When the disturbance is limited to the steady disturbance, the angular velocity spectrum shown in FIG. 5 is directly obtained using the broadband inertial sensor without obtaining FIG. 5 from the data shown in FIG. The component is detected, and the filter parameter determination unit 6 for removing the high-frequency disturbance component determines the filter parameter. Than is.

  By adopting such a configuration, it is possible to obtain the angular velocity spectrum in a short time compared to the case where the rotation speed sweep of the reaction wheel and the momentum wheel is performed.

Embodiment 11 FIG.
Next, an eleventh embodiment of the present invention will be described with reference to the drawings. FIG. 8 is a block diagram showing the configuration of the attitude control system according to the eleventh embodiment. In this figure, the same or corresponding parts as in FIG. The difference from FIG. 2 is that a gyro output post-processing system 15 including a sampler 14 with a variable sampling frequency and a high-frequency disturbance removing low-pass filter 3 is provided.

  In the normal angular velocity detection, the output from the attitude sensor is captured at a constant sampling frequency of the attitude determination system 5 as in the fourth embodiment. However, the broadband detection system 10 detects a disturbance component that is equal to or higher than the sampling frequency. In this case, the angular velocity taken into the posture determination system 5 is actually affected by aliasing, and the disturbance, which is a high frequency component, becomes low frequency.

  By detecting the high-frequency disturbance component in the broadband detection system 10 and setting the sampling frequency of the variable sampler 14 that can be set to the sampling frequency so as to be less susceptible to aliasing, the disturbance is set to a frequency sufficiently higher than the disturbance frequency. The component is taken into the gyro output post-processing system 15 without changing to a low frequency.

  Further, a low-pass filter 3 for removing high-frequency disturbance components with variable characteristics is provided. The high-frequency disturbance component removing low-pass filter 3 sets filter parameters based on the high-frequency disturbance data detected by the wideband detection system 10. The parameter setting is determined by the method described in the second embodiment.

  The variable sampler 14 and the characteristic variable high-frequency disturbance component removing low-pass filter 3 can avoid a low frequency due to high-frequency disturbance aliasing, and the high-frequency disturbance component is also removed, so that the attitude angle of the satellite is detected. It becomes possible to do. When the disturbance measured in the broadband detection system 10 is a single frequency component, a notch filter may be provided.

  Note that the sampling frequency of the sampler may differ from the frequency of the attitude control system, but in this case, it can be dealt with by accumulating the angle increments and resampling to the frequency of the attitude control system.

  By adopting such a configuration, it is possible to remove high-frequency disturbances by the gyro output post-processing system 15 before taking the output of the gyro detection system 1 into the attitude determination system 5, and the angle taken into the attitude determination system Since errors due to high-frequency disturbances are not mixed in the increment, the posture angle and posture angular velocity in the posture determination system 5 can be estimated with high accuracy, and errors due to posture control can be reduced.

Embodiment 12 FIG.
Next, an embodiment 12 of the invention will be described with reference to the drawings. FIG. 9 is a block diagram showing the configuration of the attitude control system according to the twelfth embodiment. In this figure, the same or corresponding parts as in FIG.

  2 and 8 described above, the high-frequency disturbance component is removed by filtering the detection system signal of the sensor, or the sampling is varied to avoid the low-frequency disturbance component due to aliasing. Although the influence of the high-frequency disturbance component has been removed, the embodiment 12 vibrates between the satellite structure 13 and the housing of the gyro 12 so that the high-frequency disturbance component is not input to the gyro 12 as shown in FIG. An insulation system 16 is installed. The vibration isolation system 16 has a variable characteristic so that the influence of the high-frequency disturbance component detected by the broadband detection system 10 can be ignored.

  By adopting such a configuration, it is possible to remove high-frequency disturbance components mixed in the posture angle increment output from the gyroscope 12, so that the posture angle and posture angular velocity in the posture determination system can be estimated with high accuracy. This makes it possible to reduce errors due to attitude control.

Embodiment 13 FIG.
Next, a thirteenth embodiment of the present invention will be described. 2, 8, and 9, the high-frequency disturbance component is detected inside the broadband detection system 10, but the thirteenth embodiment implements up to the spectrum processing unit 8 onboard, The result is transmitted to the ground by telemetry, the parameters of the high-frequency disturbance component removing filter 3 and the sampling frequency of the sampler are determined offline, and these parameters are transmitted to the artificial satellite.

  This configuration is effective when the high-frequency disturbance component removal filter parameter determination unit 6 cannot determine the optimum sampling frequency of the high-frequency disturbance component removal filter 3 and the variable sampler 14. Also, when the volume of angular acceleration data acquired by the broadband detection system 10 is large and on-board processing takes time, this angular velocity data is transmitted to the ground as telemetry, and high-frequency disturbance components are removed by post-processing analysis. The sampling frequency of the filter 3 for use and the variable sampler 14 may be determined, and those parameters may be transmitted to the artificial satellite.

  With such a configuration, it is possible to reduce the time required for determining the sampling frequency of the high-frequency disturbance component removing filter 3 and the variable sampler 14.

It is a schematic diagram which shows the structure of the attitude | position sensor by Embodiment 1 of this invention. It is a schematic diagram which shows the structure of the attitude | position sensor by Embodiment 2 of this invention. FIG. 10 is a schematic diagram showing an installation position relationship between a gyro and a satellite structure in a second embodiment. It is explanatory drawing which shows the arrangement | positioning state of the acceleration sensor with respect to a gyro. It is a frequency characteristic figure which shows a disturbance spectrum. It is a figure which shows the disturbance spectrum after passing through the filter with a gyro frequency characteristic. FIG. 9 is a posture angular velocity response diagram detected by the second embodiment. It is a block diagram which shows the structure of Embodiment 11 of this invention. It is a block diagram which shows the structure of the attitude | position control system by Embodiment 12 of this invention. It is explanatory drawing which shows the angular velocity detection result of the gyro by the conventional attitude | position sensor. It is a characteristic view which shows an example of the disturbance input into an attitude | position sensor. It is a figure which shows an example of the transfer function of the input / output of a gyro which has a rebalance loop.

Explanation of symbols

1 Gyro detection system, 2 Rebalance loop,
3 High-frequency disturbance component removal filter, 4 integrator, 5 attitude determination system,
6 Filter parameter determination unit for high frequency disturbance component removal, 7 Second order integration,
8 spectrum processing unit, 9 high-frequency disturbance component detection unit, 10 broadband detection system,
11 Accelerometer, 12 Gyro, 13 Satellite structure, 14 Sampler,
15 Gyro output post-processing system, 16 Vibration isolation system, 20 Differentiator.

Claims (6)

A gyro detection system having a high-frequency disturbance component removal filter in which a filter parameter is variable according to the frequency component of the high-frequency disturbance between the rebalance loop inside the gyro, which is an attitude sensor of the satellite, and the integrator, and high-frequency disturbance measurement A wideband detection system having a sensor for detecting a high-frequency disturbance component from the sensor signal, and the high-frequency disturbance component removal filter so as to remove the disturbance component estimated from the broadband detection system A satellite attitude determination system characterized by comprising a parameter determination unit for determining a filter parameter of the satellite. The broadband detection system is installed so as to be in contact with the gyro internal or gyro outside, attitude determination system of the satellite according to claim 1, characterized in that it comprises a sensor for detecting the angular velocity. The data detected by the sensor of the broadband detection system is transmitted to the ground to estimate disturbance components on the ground, and the characteristics of the filter circuit can be changed based on the transmission from the ground. The attitude determination system of the artificial satellite according to 1 . A gyro detection system to be an attitude sensor for an artificial satellite, a gyro output post-processing system having a sampler and a high-frequency disturbance component removing filter, and a broadband detection system for measuring high-frequency disturbance are provided. The attitude determination system for an artificial satellite according to claim 1 . 5. The artificial satellite attitude determination system according to claim 4, wherein the sampling timing of the sampler of the gyro output post-processing system and the parameters of the high-frequency disturbance component removal filter are determined based on the data of the broadband detection system. Above for high-frequency noise component removal filter, attitude determination system of the satellite according to claim 1, characterized in that a installed vibration isolation system between the satellite body panel and a gyro.

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