Satellite large dynamic maneuver tracking time attitude reference smooth switching method

Abstract

The present application relates to a kind of satellite large dynamic maneuver tracking time attitude reference smooth switching method, belong to satellite control system field;Step one, the whole process of large dynamic maneuver tracking is divided into 4 control stages, in turn, attitude maneuver stage, attitude keeping stage, guide tracking stage, attitude return stage;Step two, when satellite enters attitude maneuver stage, attitude reference is criterion 1;Step three, after entering attitude keeping stage, attitude reference switches to criterion 2;Step four, after entering guide tracking stage, attitude reference switches to criterion 3;Step five, enter attitude return stage, the attitude reference of attitude return stage is criterion 1;The present application realizes the smooth switching of satellite in large dynamic maneuver tracking time attitude reference by the method such as early access star sensor attitude, delay access star sensor and gyro combined filter attitude, star angular velocity exceeds star sensor normal working threshold and removes star sensor attitude as attitude reference, etc.

Description

Technical Field

[0001] This invention belongs to the field of satellite control systems and relates to a method for smooth switching of attitude reference during satellite large dynamic maneuver tracking. Background Technology

[0002] When a satellite performs dynamic tracking of ground or air targets, it first needs to maneuver from its current position to the target position. Then, it performs guided tracking based on the target's characteristics. After the tracking mission is completed, it maneuvers back to its initial position. There is usually a time limit from the current position to the target position, therefore the maneuvering angular velocity is relatively large, reaching over 2° / s. During guided tracking, the corresponding guidance angular velocity varies greatly depending on the distance between the target and the satellite. The maximum guidance velocity for conventional low-Earth orbit satellites tracking ground or near-Earth targets is approximately 0.4° / s to 1.1° / s. Currently, the commonly used attitude sensor on satellites is the star sensor, which has high measurement accuracy. However, as an optical sensor, its data becomes invalid when sunlight or atmospheric light enters the field of view. Furthermore, when the angular velocity of the star is high, it cannot effectively perform star image imaging and star point extraction. Therefore, star sensors are usually not used during attitude maneuvers. In contrast, gyroscopes, as inertial sensors, have no optical components and are unaffected by the field of view. Their angular velocity measurement range can reach 50° / s. Therefore, satellites have traditionally used gyro integration throughout the entire dynamic maneuvering tracking process. Since gyro integration is used throughout the process, the attitude reference does not change, thus ensuring the stability of the attitude. However, gyro integration is affected by constant drift and nonlinearity. During large-angle attitude maneuvers and guidance tracking, the accumulated error becomes larger and larger, which eventually leads to poor pointing accuracy, target tracking deviation, or even loss of tracking.

[0003] As the requirements for pointing accuracy in the guidance and tracking process become increasingly stringent, the traditional method of using gyro integration throughout the entire process can no longer meet the operational demands for high-precision pointing. Some satellites assign an initial value to the gyro integration once the star sensor becomes active after the attitude maneuver is complete. This method can eliminate errors accumulated during the attitude maneuver, but subsequent guidance and tracking processes will continue to accumulate errors from zero. When the guidance angular velocity is high or the guidance time is long, the accumulated error remains significant. Other satellites connect the star sensor only when it becomes active after the attitude maneuver is complete. This approach carries the risk that the star sensor may repeatedly change between active and inactive states due to high celestial angular velocities, causing repeated switching of the attitude reference and resulting in satellite jitter that makes effective tracking impossible. Summary of the Invention

[0004] The technical problem solved by this invention is to overcome the shortcomings of the prior art and propose a method for smooth switching of attitude reference during satellite large dynamic maneuver tracking. By using methods such as pre-accessing the attitude of the star sensor, delaying the access of the attitude of the star sensor and gyroscope for joint filtering, and cutting off the attitude of the star sensor as the attitude reference after the angular velocity of the satellite exceeds the normal operating threshold of the star sensor, the smooth switching of the attitude reference of the satellite during large dynamic maneuver tracking is achieved.

[0005] The solution of the present invention is:

[0006] A method for smooth attitude reference switching during satellite large dynamic maneuvering tracking includes:

[0007] Step 1: Divide the entire process of large dynamic maneuver tracking into 4 control phases, namely, attitude maneuvering phase, attitude holding phase, guidance and tracking phase, and attitude return phase.

[0008] Step 2: When the satellite enters the attitude maneuver phase, the attitude reference is Criterion 1; Criterion 1 is to use the gyro integral as the attitude reference.

[0009] Step 3: After entering the attitude maintenance phase, the attitude reference is switched to Criterion 2; Criterion 2 is to use gyro integration, single-star sensor attitude, or dual-star sensor combined attitude determination as the attitude reference.

[0010] Step 4: After entering the guidance and tracking phase, the attitude reference is switched to Criterion 3; Criterion 3 is to use gyro integration, single-star sensor attitude, or dual-star sensor combined attitude determination as the attitude reference.

[0011] Step 5: Enter the attitude return phase. The attitude reference for the attitude return phase is Criterion 1.

[0012] In the above-mentioned method for smooth switching of attitude reference during satellite large dynamic maneuver tracking, in step one, before the attitude maneuver, the star sensor continuously assigns initial values ​​to the gyroscope integral and completes the correction of the gyroscope mounting matrix and scaling factor.

[0013] In the above-mentioned method for smooth switching of attitude reference during satellite large dynamic maneuver tracking, in step two, the maneuver angular velocity is relatively large during the attitude maneuver phase. The gyroscope integral is used as the attitude reference to avoid the star sensor becoming ineffective during the maneuver, which would cause the attitude reference to switch back and forth.

[0014] In the above-mentioned method for smooth attitude reference switching during satellite large dynamic maneuvering tracking, in step three, during the attitude maintenance phase, the angular velocity of the satellite tends to stabilize. The specific content of criterion 2 is as follows:

[0015] Determine the number of valid star sensors;

[0016] When the number of effective star sensors is 0, continue to use the gyroscope integral as the attitude reference;

[0017] When the number of effective star sensors is 1, the attitude of a single star sensor is used as the attitude reference.

[0018] When the number of effective star sensors is not less than 2, the attitude determination of the dual star sensors is used as the attitude reference.

[0019] In the above-mentioned method for smooth switching of attitude reference during large dynamic maneuvering tracking of satellites, when the attitude of a single satellite sensor is used as the attitude reference, the attitude determination of the satellite sensor and the gyroscope starts simultaneously. After 1 minute of joint attitude determination, the attitude determination of the satellite sensor and the gyroscope is used as the attitude reference.

[0020] The above-mentioned method for smooth attitude reference switching during satellite large dynamic maneuvering and tracking uses the star sensor as the attitude reference to correct attitude deviations caused by gyro integration during attitude maneuvering. At the same time, the guidance and tracking task has not yet started at this stage, so the attitude jitter caused by the gyro integration switching to the star sensor attitude will not affect the smooth operation of the service.

[0021] In the above-mentioned method for smooth attitude reference switching during satellite large dynamic maneuvering tracking, the specific content of criterion 3 in step four is as follows:

[0022] Determine if the angular velocity of a star exceeds the normal operating threshold of the star sensor;

[0023] When the normal operating threshold is exceeded, the system switches to use gyroscope integral as the attitude reference.

[0024] When the normal operating threshold is not exceeded, the attitude reference process is consistent with that in step three.

[0025] In the aforementioned method for smoothly switching the attitude reference during satellite large dynamic maneuver tracking, once the attitude reference is switched to gyro integral due to the star's angular velocity exceeding the normal operating threshold of the star sensor, the attitude reference will not be switched again in subsequent processes and will always remain at the gyro integral, thus avoiding repeated switching of the attitude reference during guidance and tracking that could affect operational status.

[0026] In the above-mentioned method for smooth switching of attitude reference during satellite large dynamic maneuver tracking, the normal operating threshold of the star sensor is related to the individual characteristics of the star sensor.

[0027] The advantages of this invention compared to the prior art are:

[0028] (1) By dividing the entire process of maneuvering tracking into appropriate control stages, and by using methods such as early access to the star sensor, delayed access to the star sensor and gyroscope joint filtering of attitude, this invention achieves the requirement that the satellite can quickly and accurately point to the target and smoothly switch the attitude reference in different control stages during large dynamic maneuvering tracking.

[0029] (2) The attitude reference in the attitude maneuvering phase of this invention is criterion 1, that is, the gyro integral is used as the attitude reference. At this time, the maneuvering angular velocity is large. Using the gyro integral as the attitude reference can avoid the star sensor from becoming ineffective during the maneuvering process, causing the attitude reference to switch back and forth.

[0030] (3) After entering the attitude holding stage, since the star sensor has been used as the attitude reference, the attitude deviation caused by the gyroscope integration during the attitude maneuver can be repaired. At the same time, the guidance and tracking task has not yet started in this stage. Therefore, the attitude jitter caused by the gyroscope integration switching to the star sensor attitude will not affect the smooth operation of the service.

[0031] (4) When the present invention enters the guidance and tracking stage, the attitude reference is switched to gyro integral because the star angular velocity exceeds the normal working threshold of the star sensor. In the subsequent process, the attitude reference will not be switched again and will always be gyro integral, so as to avoid the repeated switching of attitude reference during the guidance and tracking process from affecting the operation of the service. Attached Figure Description

[0032] Figure 1 This is a flowchart illustrating the smooth switching of the attitude reference in this invention. Detailed Implementation

[0033] The present invention will be further described below with reference to the embodiments.

[0034] This invention provides a method for smoothly switching the attitude reference during satellite dynamic maneuvering and tracking. By pre-accessing the star sensor attitude, delaying the access of the star sensor and gyroscope-based attitude filtering, and removing the star sensor attitude as the attitude reference after the satellite's angular velocity exceeds the normal operating threshold of the star sensor, the method achieves smooth switching of the attitude reference during satellite dynamic maneuvering and tracking.

[0035] Methods for smooth attitude reference switching during satellite large dynamic maneuvering tracking, such as Figure 1 As shown, the specific steps include the following:

[0036] Step 1: Divide the entire process of large dynamic maneuver tracking into 4 control phases, namely, attitude maneuvering phase, attitude maintenance phase, guidance and tracking phase, and attitude return phase.

[0037] Before attitude maneuvering, the star sensor continuously assigns initial values ​​to the gyroscope integral and completes the correction of the gyroscope mounting matrix and scaling factor.

[0038] Step 2: When the satellite enters the attitude maneuver phase, the attitude reference is Criterion 1; Criterion 1 uses the gyro integral as the attitude reference. During the attitude maneuver phase, the maneuver angular velocity is relatively large. Using the gyro integral as the attitude reference avoids the attitude reference switching back and forth due to the star sensor becoming ineffective during the maneuver.

[0039] Step 3: After entering the attitude maintenance phase, the attitude reference is switched to Criterion 2; Criterion 2 is to use gyro integration, single-star sensor attitude, or dual-star sensor combined attitude determination as the attitude reference.

[0040] During the attitude maintenance phase, the angular velocity of the celestial body tends to stabilize, and the specific content of criterion 2 is as follows:

[0041] Determine the number of valid star sensors;

[0042] When the number of effective star sensors is 0, continue to use the gyroscope integral as the attitude reference;

[0043] When the number of effective star sensors is 1, the attitude of a single star sensor is used as the attitude reference.

[0044] When the number of effective star sensors is not less than 2, the attitude determination of the dual star sensors is used as the attitude reference.

[0045] When the attitude of a single star sensor is used as the attitude reference, the attitude determination of the star sensor and the gyroscope starts simultaneously. After 1 minute of joint attitude determination, the attitude determination of the star sensor and the gyroscope is used as the attitude reference.

[0046] By using the star sensor as the attitude reference, the attitude deviation caused by gyro integration during attitude maneuvering is corrected; at the same time, the guidance and tracking task has not yet started at this stage, so the attitude jitter caused by the gyro integration switching to the star sensor attitude will not affect the smooth operation of the service.

[0047] Step 4: After entering the guidance and tracking phase, the attitude reference is switched to Criterion 3; Criterion 3 is to use gyro integration, single-star sensor attitude, or dual-star sensor combined attitude determination as the attitude reference.

[0048] The specific content of Guideline 3 is as follows:

[0049] Determine if the angular velocity of a star exceeds the normal operating threshold of the star sensor. The normal operating threshold of the star sensor is related to the individual characteristics of the star sensor.

[0050] When the normal operating threshold is exceeded, switch to using gyroscope integral as the attitude reference.

[0051] Once the attitude reference is switched to gyro integral due to the star's angular velocity exceeding the normal operating threshold of the star sensor, the attitude reference will not be switched again in subsequent processes and will always remain at the gyro integral, thus avoiding repeated switching of the attitude reference during guidance and tracking that could affect business operations.

[0052] When the normal operating threshold is not exceeded, the attitude reference process is consistent with that in step three.

[0053] Step 5: Enter the attitude return phase. The attitude reference for the attitude return phase is Criterion 1.

[0054] Example

[0055] Step S1: The entire large dynamic maneuver tracking process is divided into four control phases: attitude maneuvering phase, attitude holding phase, guidance and tracking phase, and attitude return phase. Before the attitude maneuver, the star sensor continuously assigns initial values ​​to the gyroscope integral and completes the gyroscope mounting matrix and scaling factor correction.

[0056] Step S2: The attitude reference during the attitude maneuver phase is Criterion 1, i.e., the gyro integral is used as the attitude reference. At this time, the maneuver angular velocity is relatively large, and using the gyro integral as the attitude reference can avoid the attitude reference switching back and forth due to the star sensor becoming ineffective during the maneuver.

[0057] Step S3: After entering the attitude maintenance phase, the attitude reference is switched to criterion 2. At this time, the angular velocity of the star has stabilized, and the star sensor is likely effective. If the star sensor is effective, its attitude is used as the attitude reference, and the joint attitude determination using the star sensor and gyroscope begins. After 1 minute, the joint attitude determination using the star sensor and gyroscope is used as the attitude reference. If the star sensor is ineffective, gyroscope integration continues to be used. If multiple star sensors are effective, the joint attitude determination using two star sensors is used as the star sensor attitude; otherwise, the solution value from a single star sensor is used as the star sensor attitude. In this phase, since the star sensor is used as the attitude reference, attitude deviations caused by gyroscope integration during attitude maneuvers can be corrected. Furthermore, since the guidance and tracking task has not yet started in this phase, attitude jitter caused by the switch from gyroscope integration to the star sensor attitude will not affect the smooth operation of the service.

[0058] Step S4: After entering the guidance and tracking phase, the attitude reference is switched to criterion 3. The difference from criterion 2 is that this stage requires determining whether the star's angular velocity exceeds the normal operating threshold of the star sensor. If it does, AttGU is set to true (initially false), and the attitude reference is switched to gyro integral; otherwise, the attitude reference process is the same as in step S3. If the attitude reference is switched to gyro integral because the star's angular velocity exceeds the normal operating threshold of the star sensor, the attitude reference will not be switched again in subsequent processes and will always remain at the gyro integral, avoiding repeated attitude reference switching during guidance and tracking that could affect operational processes. The normal operating threshold of the star sensor is related to the characteristics of the individual unit.

[0059] Step S5: The attitude return phase is consistent with the attitude maneuver phase, and the attitude reference is criterion 1, that is, the gyro integral is used as the attitude reference.

[0060] This invention achieves the requirement that satellites can quickly and accurately point to the target and smoothly switch attitude references at different control stages during high-dynamic maneuvering tracking by appropriately dividing the entire maneuvering tracking process into control stages, and by using methods such as early access to star sensors, delayed access to star sensors, and gyroscope-based attitude filtering.

[0061] In this invention, the attitude reference during the attitude maneuver phase is criterion 1, that is, the gyro integral is used as the attitude reference. At this time, the maneuver angular velocity is relatively large. Using the gyro integral as the attitude reference can avoid the star sensor becoming ineffective during the maneuver, which would cause the attitude reference to switch back and forth.

[0062] After entering the attitude holding phase, since the star sensor has been used as the attitude reference, the attitude deviation caused by the gyro integration during the attitude maneuver can be corrected. At the same time, the guidance and tracking task has not yet started in this phase. Therefore, the attitude jitter caused by the gyro integration switching to the star sensor attitude will not affect the smooth operation of the service.

[0063] When entering the guidance and tracking phase, if the angular velocity of the star exceeds the normal operating threshold of the star sensor, the attitude reference is switched to the gyro integral. In subsequent processes, the attitude reference will not be switched again and will always be the gyro integral, so as to avoid repeated switching of the attitude reference during the guidance and tracking process from affecting the operation of the business.

[0064] Although the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solutions of the present invention by utilizing the methods and techniques disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and alterations made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solutions of the present invention shall fall within the protection scope of the technical solutions of the present invention.

Claims

1. A method for smooth switching of attitude reference for large dynamic maneuvering tracking of a satellite, characterized in that: The application relates to a satellite attitude control method. Step one: the whole process of large dynamic maneuvering tracking is divided into four control stages, namely an attitude maneuvering stage, an attitude keeping stage, a guiding tracking stage and an attitude returning stage; Step two: when the satellite enters the attitude maneuvering stage, a criterion 1 is used as the attitude reference; the criterion 1 uses gyro integration as the attitude reference; Step three: after entering the attitude keeping stage, a criterion 2 is used as the attitude reference; the criterion 2 uses gyro integration, single-star sensor attitude or double-star sensor combined attitude as the attitude reference; Step four: after entering the guiding tracking stage, a criterion 3 is used as the attitude reference; the criterion 3 uses gyro integration, single-star sensor attitude or double-star sensor combined attitude as the attitude reference; Step five: after entering the attitude returning stage, the criterion 1 is used as the attitude reference. In the step one, before the attitude maneuvering, the star sensor continuously gives the initial value of the gyro integration, and the gyro installation matrix and the scale factor correction are completed.

2. The method of claim 1, wherein: In the step two, during the attitude maneuvering stage, the gyro integration is used as the attitude reference to avoid the invalidation of the star sensor during the maneuvering process.

3. The method of claim 1, wherein: In the step three, after entering the attitude keeping stage, the star angular velocity tends to be stable, and the criterion 2 is used as the attitude reference; the criterion 2 comprises the following steps:

4. The method of claim 3, wherein: judging the number of effective star sensors; when the number of effective star sensors is 0, the gyro integration is continuously used as the attitude reference; when the number of effective star sensors is 1, the single-star sensor attitude is used as the attitude reference; when the number of effective star sensors is not less than 2, the double-star sensor combined attitude is used as the attitude reference. When the single-star sensor attitude is used as the attitude reference, the star sensor and the gyro are combined to start the calculation, and the star sensor and the gyro are combined to be used as the attitude reference after 1 min.

5. The method of claim 4, wherein: The star sensor is used as the attitude reference to repair the attitude deviation caused by the gyro integration during the attitude maneuvering process.

6. The method of claim 5, wherein: The attitude keeping stage has not started to execute the guiding tracking task, and the attitude jitter caused during the process of switching the gyro integration to the star sensor attitude will not affect the smooth operation of the business.

7. The method of claim 6, wherein the method further comprises: determining a time when the satellite is in a stable attitude reference state; and switching the satellite to the stable attitude reference state at the determined time. In the step four, the criterion 3 is used as the attitude reference; the criterion 3 comprises the following steps:

8. The method for smooth attitude reference switching during satellite large dynamic maneuvering tracking according to claim 4, characterized in that: judging whether the star angular velocity exceeds the normal working threshold of the star sensor; when the star angular velocity exceeds the normal working threshold, the gyro integration is used as the attitude reference; when the star angular velocity does not exceed the normal working threshold, the process of the attitude reference is consistent with that in the step three. When the attitude reference is switched to the gyro integration due to the star angular velocity exceeding the normal working threshold of the star sensor, the attitude reference is not switched in the subsequent process, and the gyro integration is always used as the attitude reference, so that the repeated switching of the attitude reference during the guiding tracking process is avoided, and the business operation is not affected.

9. The method of claim 8, wherein: The normal working threshold of the star sensor is related to the single-machine characteristics of the star sensor.

10. The method of claim 8, wherein: ​

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