Attitude reference stable switching method during large dynamic maneuvering tracking of satellite

Abstract

The invention relates to an attitude reference stable switching method during large dynamic maneuvering tracking of a satellite, and belongs to the field of satellite control systems. The method comprises the following steps of: 1, dividing a whole large dynamic maneuvering tracking process into four control stages, namely an attitude maneuvering stage, an attitude keeping stage, a guide tracking stage and an attitude returning stage in sequence; 2, when the satellite enters an attitude maneuver stage, the attitude reference is a criterion 1; step 3, after entering the attitude keeping stage, switching the attitude reference to a criterion 2; step 4, after entering a guide tracking stage, switching the attitude reference to a criterion 3; 5, entering an attitude return stage, wherein the attitude reference of the attitude return stage is a criterion 1; according to the method, the attitude reference stable switching of the satellite during large dynamic maneuvering tracking is realized through the methods of accessing a star sensor attitude in advance, accessing a star sensor and gyroscope combined filtering attitude in a delayed manner, cutting off the star sensor attitude as the attitude reference after the star angular velocity exceeds the normal working threshold of the star sensor, and the like.

Description

Technical Field

[0001] The invention belongs to the field of satellite control systems and relates to a method for smoothly switching attitude references during large-dynamic maneuver tracking of a satellite. Background Art

[0002] When a satellite performs dynamic tracking of a ground or air target, it first needs to maneuver from its current position to the target position, then perform guidance and tracking based on the target characteristics, and then maneuver back to its initial position after the tracking task is completed. There is usually a time limit from the current position to the target position, so the maneuvering angular velocity is relatively large, which can reach more than 2° / s; during the guidance and tracking process, depending on the distance between the target and the satellite, the corresponding guidance angular velocity span is relatively large, and the maximum guidance velocity of conventional low-orbit satellites for ground or near-earth targets is about 0.4° / s to 1.1° / s. At present, the commonly used attitude sensor on the satellite is the star sensor, which has high measurement accuracy, but as an optical sensor, the data will be invalid after sunlight or ground light enters the field of view. In addition, when the angular velocity of the star is large, it is impossible to perform star imaging and extract star points well, so star sensors are usually not used during attitude maneuvers. Relatively speaking, as an inertial sensor, the gyroscope has no optical device and is not affected by the field of view. The angular velocity measurement range can reach 50° / s. Therefore, satellites usually use gyro integration in the whole process of dynamic maneuvering tracking. Since gyro integration is used throughout the process, the attitude reference is not switched, so the attitude is stable. However, gyro integration is affected by constant drift and nonlinearity. The cumulative error becomes larger and larger during large-angle attitude maneuvers and guidance tracking, which will eventually lead to poor pointing accuracy, deviation in target tracking, and even loss of tracking.

[0003] As the requirements for pointing accuracy in the guidance and tracking process become higher and higher, the traditional method of using gyro integration throughout the process can no longer meet the business needs of high-precision pointing. Some satellites assign an initial value to the gyro integral when the star sensor is valid after the attitude maneuver is in place. This method can eliminate the accumulated errors during the attitude maneuver, but the subsequent guidance and tracking process will continue to accumulate from zero. When the guidance angular velocity is large or the guidance time is long, the accumulated error is still large. Some satellites also connect the star sensor when the star sensor is valid after the attitude maneuver is in place. This approach carries the risk that the star sensor may repeatedly change between valid and invalid due to the large angular velocity of the star, causing the attitude reference to repeatedly switch, resulting in satellite jitter and ineffective tracking. Summary of the invention

[0004] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art and propose a method for smoothly switching the attitude reference during large-dynamic maneuver tracking of a satellite, by means of pre-accessing the attitude of a star sensor, delaying the access to the attitude of a star sensor and a gyroscope combined filtering, and removing the attitude of a star sensor as an attitude reference after the angular velocity of a star exceeds the normal working threshold of the star sensor, so as to achieve smooth switching of the attitude reference of a satellite during large-dynamic maneuver tracking.

[0005] The solution of the present invention is:

[0006] A method for smoothly switching attitude reference during large-dynamic maneuver tracking of a satellite, comprising:

[0007] Step 1: Divide the whole process of large dynamic maneuver tracking into four control stages, namely attitude maneuver stage, attitude holding stage, guidance and tracking stage, and attitude return stage;

[0008] Step 2: When the satellite enters the attitude maneuvering 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 holding stage, the attitude reference is switched to criterion 2; criterion 2 is to use gyro integration, single star sensor attitude or dual star sensor joint 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 joint attitude determination as the attitude reference;

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

[0012] In the above-mentioned method for smoothly switching attitude reference during large-dynamic maneuver tracking of a satellite, in step 1, before the attitude maneuver, the star sensor continuously assigns initial values ​​to the gyro integral and completes the correction of the gyro installation matrix and the scale factor.

[0013] In the above-mentioned method for smoothly switching the attitude reference during large-dynamic maneuver tracking of a satellite, in step 2, the maneuvering angular velocity is relatively large during the attitude maneuvering stage, and the gyro integral is used as the attitude reference to avoid the attitude reference switching back and forth due to the invalidation of the star sensor during the maneuvering process.

[0014] In the above-mentioned method for smoothly switching the attitude reference during large-dynamic maneuver tracking of a satellite, in step 3, the attitude holding phase is entered, and the angular velocity of the satellite tends to be stable. The specific content of criterion 2 is:

[0015] Determine the number of effective star sensors;

[0016] When the number of effective star sensors is 0, continue to use the gyro 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 dual star sensor joint attitude determination is used as the attitude reference.

[0019] In the above-mentioned method for smoothly switching the attitude reference during large-dynamic maneuver tracking of a satellite, when the attitude of a single star sensor is used as the attitude reference, the joint attitude determination of the star sensor and the gyroscope starts at the same time. After the joint attitude determination has been completed for 1 minute, the joint attitude determination of the star sensor and the gyroscope is used as the attitude reference.

[0020] In the above-mentioned method for smoothly switching the attitude reference during large-dynamic maneuver tracking of a satellite, the attitude deviation caused by the gyro integral during the attitude maneuver is repaired by using the star sensor as the attitude reference; at the same time, the guidance and tracking task has not yet started in this stage, so the attitude jitter caused by the gyro integral switching to the star sensor attitude will not affect the smooth operation of the business.

[0021] In the above-mentioned method for smoothly switching attitude reference during large-dynamic maneuver tracking of a satellite, in step 4, the specific content of criterion 3 is:

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

[0023] When the normal working threshold is exceeded, the attitude reference is switched to the gyro integral.

[0024] When the normal operating threshold is not exceeded, the attitude reference process is consistent with step 3.

[0025] In the above-mentioned method for smoothly switching the attitude reference during large-dynamic maneuver tracking of a satellite, when the attitude reference is switched to the gyro integration because the angular velocity of the satellite exceeds the normal working threshold of the star sensor, the attitude reference is no longer switched in the subsequent process and is always the gyro integration, thereby avoiding repeated switching of the attitude reference during the guidance and tracking process to affect business operations.

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

[0027] The beneficial effects of the present invention compared with the prior art are:

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

[0029] (2) The attitude reference of the present invention in the attitude maneuvering stage is criterion 1, that is, the gyro integral is used as the attitude reference. At this time, the maneuvering angular velocity is relatively large. Using the gyro integral as the attitude reference can avoid the ineffectiveness of the star sensor during the maneuvering process, which causes the attitude reference to switch back and forth;

[0030] (3) After entering the attitude holding stage, the present invention can repair the attitude deviation caused by the gyro integration during the attitude maneuver because the star sensor has been used as the attitude reference. At the same time, the guidance and tracking task has not yet started in 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 business;

[0031] (4) When the present invention enters the guidance and tracking stage, the attitude reference is switched to the gyro integration because the angular velocity of the star exceeds the normal working threshold of the star sensor. In the subsequent process, the attitude reference is no longer switched and is always the gyro integration, thereby avoiding repeated switching of the attitude reference during the guidance and tracking process to affect business operations. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] Figure 1 This is a flow chart of the smooth switching of the attitude reference of the present invention. DETAILED DESCRIPTION

[0033] The present invention will be further described below in conjunction with the embodiments.

[0034] The present invention provides a method for smoothly switching the attitude reference during large-dynamic maneuver tracking of a satellite. By means of methods such as early access to a star sensor attitude, delayed access to a star sensor and a gyroscope joint filtering attitude, and cutting off the star sensor attitude as an attitude reference after the angular velocity of a star exceeds a normal working threshold of the star sensor, the method achieves smooth switching of the attitude reference during large-dynamic maneuver tracking of a satellite.

[0035] The method of smoothly switching the attitude reference during large dynamic maneuver tracking of satellites, such as Figure 1 As shown, the specific steps include:

[0036] Step 1: Divide the entire large dynamic maneuver tracking process into four control stages, namely attitude maneuver stage, attitude holding stage, guidance and tracking stage, and attitude return stage.

[0037] Before the attitude maneuver, the star sensor continuously assigns initial values ​​to the gyro integral and completes the gyro installation matrix and scale factor correction.

[0038] Step 2: When the satellite enters the attitude maneuvering stage, the attitude reference is criterion 1; criterion 1 is to use the gyro integral as the attitude reference. During the attitude maneuvering stage, the maneuvering angular velocity is relatively large, and using the gyro integral as the attitude reference can avoid the star sensor being invalid during the maneuvering process, resulting in the attitude reference switching back and forth.

[0039] Step 3: After entering the attitude holding stage, the attitude reference is switched to criterion 2; criterion 2 is to use gyro integration, single star sensor attitude or dual star sensor joint attitude determination as the attitude reference.

[0040] Entering the attitude holding stage, the angular velocity of the star tends to be stable. The specific content of criterion 2 is:

[0041] Determine the number of effective star sensors;

[0042] When the number of effective star sensors is 0, continue to use the gyro 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 dual star sensor joint attitude determination is used as the attitude reference.

[0045] Among them, when the attitude of a single star sensor is used as the attitude reference, the joint attitude determination of the star sensor and the gyroscope starts at the same time. After the joint attitude determination has been completed for 1 minute, the joint 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 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 stage, so the attitude jitter caused by the gyro integration switching to the star sensor attitude will not affect the smooth operation of the business.

[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 joint attitude determination as the attitude reference;

[0048] The specific content of Guideline 3 is:

[0049] Determine whether the angular velocity of the star exceeds the normal working threshold of the star sensor. The normal working threshold of the star sensor is related to the single machine characteristics of the star sensor.

[0050] When the normal operating threshold is exceeded, the attitude reference is switched to the gyro integral.

[0051] When the attitude reference is switched to gyro integration because the angular velocity of a star exceeds the normal operating threshold of the star sensor, the attitude reference will no longer be switched in the subsequent process and will always be gyro integration, avoiding repeated switching of the attitude reference during the guidance and tracking process to affect business operations.

[0052] When the normal operating threshold is not exceeded, the attitude reference process is consistent with step 3.

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

[0054] Example

[0055] Step S1: The whole process of large dynamic maneuver tracking is divided into four control stages, namely attitude maneuver stage, attitude hold stage, guidance and tracking stage, and attitude return stage. Before the attitude maneuver, the star sensor continuously assigns initial values ​​to the gyro integral and completes the gyro installation matrix and scale factor correction.

[0056] Step S2: The attitude reference in the attitude maneuvering stage is criterion 1. That is, the gyro integral is used as the attitude reference. At this time, the maneuvering angular velocity is relatively large, and using the gyro integral as the attitude reference can avoid the star sensor being invalid during the maneuvering process, resulting in the attitude reference switching back and forth.

[0057] Step S3: After entering the attitude holding stage, the attitude reference is switched to criterion 2. At this time, the angular velocity of the star body has stabilized, and the star sensor is most likely valid. If the star sensor is valid, the attitude of the star sensor is used as the attitude reference, and the joint attitude determination of the star sensor and the gyro begins. After 1 minute, the joint attitude determination of the star sensor and the gyro is used as the attitude reference; if the star sensor is invalid, the gyro integration continues to be used. If multiple star sensors are valid, the dual star sensor joint attitude determination is used as the star sensor attitude, otherwise the single star sensor solution value is used as the star sensor attitude. At this stage, 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 repaired. 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 business.

[0058] Step S4: After entering the guidance and tracking stage, the attitude reference switches to criterion 3. The difference from criterion 2 is that this stage needs to determine whether the angular velocity of the star exceeds the normal working threshold of the star sensor. If it exceeds, AttGU is set to true (the initial value is false), and the attitude reference switches to gyro integration; otherwise, it is consistent with the attitude reference process in step S3. If the attitude reference is switched to gyro integration because the angular velocity of the star exceeds the normal working threshold of the star sensor, the attitude reference will no longer be switched in the subsequent process and will always be gyro integration, avoiding repeated switching of the attitude reference during the guidance and tracking process to affect business operations. The normal working threshold of the star sensor is related to the characteristics of a single machine.

[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] The present invention achieves the requirement that the satellite can point to the target quickly and with high precision during large dynamic maneuver tracking and can smoothly switch the attitude reference in different control stages by dividing the whole process of maneuver tracking into appropriate control stages, and using methods such as early access to star sensors, delayed access to star sensors and gyro joint filtering attitude.

[0061] The attitude reference of the present invention in the attitude maneuvering stage is criterion 1, that is, the gyro integral is used as the attitude reference. At this time, the maneuvering angular velocity is relatively large. Using the gyro integral as the attitude reference can avoid the star sensor being invalid during the maneuvering process, resulting in the attitude reference switching back and forth.

[0062] After entering the attitude holding stage, 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 repaired. At the same time, the guidance and tracking task has not yet started in 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 business.

[0063] When entering the guidance and tracking stage, the attitude reference is switched to gyro integration because the angular velocity of the star exceeds the normal working threshold of the star sensor. In the subsequent process, the attitude reference will no longer be switched and will always be gyro integration, avoiding repeated switching of the attitude reference during the guidance and tracking process to affect business operations.

[0064] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A method for smoothly switching attitude reference during large dynamic maneuver tracking of a satellite, characterized in that: include: Step 1: Divide the whole process of large dynamic maneuver tracking into four control stages, namely attitude maneuver stage, attitude holding stage, guidance and tracking stage, and attitude return stage; Step 2: When the satellite enters the attitude maneuvering phase, the attitude reference is criterion 1; criterion 1 is to use the gyro integral as the attitude reference; Step 3: After entering the posture holding stage, the posture reference is switched to criterion 2; Criterion 2 is to use gyro integration, single star sensor attitude or dual star sensor joint attitude determination as attitude reference; 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 joint attitude determination as the attitude reference; Step 5: Enter the attitude return phase. The attitude reference of the attitude return phase is criterion 1.

2. The method for smoothly switching attitude reference during large dynamic maneuver tracking of a satellite according to claim 1, characterized in that: In the step 1, before the attitude maneuver, the star sensor continuously assigns initial values ​​to the gyro integral and completes the gyro installation matrix and scale factor correction.

3. The method for smoothly switching attitude reference during large dynamic maneuver tracking of a satellite according to claim 1, characterized in that: In the step 2, the maneuvering angular velocity is relatively large during the attitude maneuvering stage, and the gyro integral is used as the attitude reference to avoid the star sensor being invalid during the maneuvering process, which causes the attitude reference to switch back and forth.

4. The method for smoothly switching attitude reference during large dynamic maneuver tracking of a satellite according to claim 3, characterized in that: In step 3, the attitude holding stage is entered, and the angular velocity of the star tends to be stable. The specific content of criterion 2 is: Determine the number of effective star sensors; When the number of effective star sensors is 0, continue to use the gyro integral as the attitude reference; When the number of effective star sensors is 1, the attitude of a single star sensor is used as the attitude reference; When the number of effective star sensors is not less than 2, the dual star sensor joint attitude determination is used as the attitude reference.

5. The method for smoothly switching attitude reference during large dynamic maneuver tracking of a satellite according to claim 4, characterized in that: When the attitude of a single star sensor is used as the attitude reference, the joint attitude determination of the star sensor and the gyro starts at the same time. After the joint attitude determination has lasted for 1 minute, the joint attitude determination of the star sensor and the gyro is used as the attitude reference.

6. The method for smoothly switching attitude reference during large dynamic maneuver tracking of a satellite according to claim 5, characterized in that: By using the star sensor as the attitude reference, the attitude deviation caused by the gyro integration during the attitude maneuver can be corrected.

7. The method for smoothly switching attitude reference during large dynamic maneuver tracking of a satellite according to claim 6, characterized in that: The guidance and tracking tasks have not yet started in the attitude holding stage, and the attitude jitter caused by the gyro integral switching to the star sensor attitude will not affect the smooth operation of the business.

8. The method for smoothly switching attitude reference during large dynamic maneuver tracking of a satellite according to claim 4, characterized in that: In step 4, the specific content of criterion 3 is: Determine whether the angular velocity of the star exceeds the normal operating threshold of the star sensor; When the normal working threshold is exceeded, the attitude reference is switched to the gyro integral. When the normal operating threshold is not exceeded, the attitude reference process is consistent with step 3.

9. The method for smoothly switching attitude reference during large dynamic maneuver tracking of a satellite according to claim 8, characterized in that: When the attitude reference is switched to gyro integration because the angular velocity of a star exceeds the normal operating threshold of the star sensor, the attitude reference will no longer be switched in the subsequent process and will always be gyro integration, avoiding repeated switching of the attitude reference during the guidance and tracking process to affect business operations.

10. The method for smoothly switching attitude reference during large dynamic maneuver tracking of a satellite according to claim 8, characterized in that: The normal operating threshold of the star sensor is related to the single-machine characteristics of the star sensor.

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