CN110816898A - Large-angle momentum compensation satellite three-stage instability judgment and control design method - Google Patents
Large-angle momentum compensation satellite three-stage instability judgment and control design method Download PDFInfo
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- 239000013598 vector Substances 0.000 claims description 11
- 230000036544 posture Effects 0.000 claims description 6
- 230000009977 dual effect Effects 0.000 claims description 3
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- 230000009286 beneficial effect Effects 0.000 description 1
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- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/244—Spacecraft control systems
- B64G1/245—Attitude control algorithms for spacecraft attitude control
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/26—Guiding or controlling apparatus, e.g. for attitude control using jets
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Abstract
The invention provides a three-stage instability judgment and control design method for a large-angle momentum compensation satellite, wherein the satellite is in ground three-axis flywheel steady state control, the attitude of the satellite is out of tolerance, a first-stage instability judgment condition is achieved, and the satellite autonomously enters a magnetic control mode; the satellite is in the steady state control of the earth triaxial flywheel, the attitude of the satellite is out of tolerance, the satellite reaches the second-stage instability judgment condition, the satellite autonomously shifts into a steady state open loop mode, and further, the satellite enters a sun-oriented mode after being in an illumination area and losing the sun; the satellite is in the steady state control of the ground triaxial flywheel, the attitude of the satellite is out of tolerance, the three-level instability judgment condition is achieved, the satellite automatically enters an air injection control mode, and the control returns to the control of the ground triaxial flywheel after the control is stable. The invention provides a specific solution for the phenomenon that the satellite attitude rolls over due to the fluctuation of the rotating speed of an on-orbit large rotating part or abnormal clamping stagnation from the characteristic of a large-angle compensation satellite, and the satellite can keep three-axis ground control through the invention, thereby ensuring that the continuous work of the satellite service is not influenced.
Description
Technical Field
The invention relates to a satellite overall optimization design method, in particular to a large-angle momentum compensation satellite three-level instability judgment and control design method.
Background
The space passive microwave remote sensing is one of important ways for earth observation, and the full-polarization microwave radiometer is used as a novel space passive microwave remote sensing instrument and provides an important means for observing atmospheric marine environment parameters such as sea surface wind field, sea surface temperature, atmospheric water vapor content, rainfall intensity and the like. The full-polarization microwave radiometer adopts a conical scanning mode, has large rotational inertia, requires a designed scanning control system of the microwave radiometer to have good driving control capability, and simultaneously needs a satellite to compensate angular momentum generated by rotation so as to achieve stable ground attitude of three axes.
In order for the satellite to effectively compensate for the momentum generated by the antenna scan of the on-board microwave radiometer during the overall design of the satellite, it is necessary to obtain reliable compensation data by using an appropriate method. Meanwhile, the satellite still can keep a stable attitude under the condition that the days of the microwave radiometer rotate at an abnormal speed, which puts higher requirements on the design of the whole satellite.
The prior art related to the present application is patent document CN103112603A, which discloses a method for establishing a normal attitude by an under-actuated high-speed spinning satellite, comprising the following steps: (1) determining the spin axis of the under-actuated satellite by utilizing the output data of the attitude sensor; (2) determining an underactuated shaft and a normal shaft; (3) performing under-actuated despinning and precession control on the satellite until the gyroscope is desaturated; (4) after the gyroscope is desaturated, under-actuated control is carried out on the three-axis angular velocity; (5) determining and updating an initial attitude quaternion; (6) and (3) adopting the momentum wheel to capture the attitude and unload the magnetic torquer, determining the attitude of the satellite, and restoring the satellite to a normal three-axis stable attitude relative to the ground.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a large-angle momentum compensation satellite three-stage instability judgment and control design method.
The invention provides a large-angle momentum compensation satellite three-level instability judgment and control design method, which comprises the following steps:
step 1: the satellite is in the steady state control of the earth triaxial flywheel, the satellite attitude reaches a primary instability state, and the satellite autonomously enters a magnetic control mode;
step 2: the satellite is in the steady state control of the earth triaxial flywheel, the satellite attitude reaches a secondary instability state, the satellite autonomously shifts into a steady state open loop mode, the satellite is in an uncontrolled state, and the satellite is ready for ground processing;
and step 3: the satellite is in the steady state control of the ground triaxial flywheel, the attitude of the satellite reaches a three-level instability state, the satellite autonomously enters an air injection control mode, and the steady state control returns to the steady state control of the ground triaxial flywheel after the control is stable.
Preferably, the primary instability state refers to that the three-axis postures of the satellite in the roll axis, the pitch axis and the yaw axis within 120s are respectively greater than 3 °,4 ° and 3 °.
Preferably, the secondary instability state means that the satellite has 30s of continuous internal rolling axis, pitching axis and yawing axis three-axis postures more than 8 degrees.
Preferably, the three-stage instability state is that the three-axis attitude of the rolling axis, the pitching axis and the yawing axis of the satellite is greater than 12 degrees within 5 seconds.
Preferably, in the step 1, the attitude reference of the satellite attitude sequentially depends on the star sensor, the double vector, the gyro integral and the infrared earth sensor, wherein the star sensor has the highest priority, and the output attitude of the star sensor is preferentially used as a judgment control basis within 2 degrees/s of the satellite attitude angular velocity; when the star sensor cannot output data or has stray light interference, double vectors are adopted as attitude references; when the dual vectors are invalid, the gyro integral is used as the attitude reference.
Preferably, in the step 2, the to-be-processed ground is to judge the sun azimuth by using a sun sensor, and when no signal is output in the sun sensitivity of all quadrants, the to-be-processed ground is judged to be in a ground shadow area, and the output state without control moment is continuously kept; otherwise, when the sun-exposed surface of the satellite is not illuminated, the satellite automatically switches into a sun-oriented control mode.
Preferably, in step 3, the attitude reference of the satellite attitude is based on an infrared earth sensor or a gyro integrator.
Preferably, the rolling shaft and the yaw shaft are controlled by adopting the precession of an air injection stamp, and the pitch shaft is controlled by adopting an inclined switch wire.
Preferably, the air injection control mode is that three-axis oblique switch line control is selected, the flywheel outputs a nominal rotating speed, the PI calculated value of the flywheel is assigned with zero output, magnetic control is isolated, and the magnetic current output of the magnetic torquer is zero.
Preferably, the return to the ground triaxial flywheel steady state control after the control is stable is that when the satellite triaxial attitude is less than 6 degrees and the angular velocity is less than 0.1 degree/s within 60 seconds, the satellite exits the air injection control, accesses the flywheel control mode and accesses the magnetic torquer.
Compared with the prior art, the invention has the following beneficial effects:
the invention provides a specific solution for the phenomenon that the satellite attitude rolls over due to the fluctuation of the rotating speed of an on-orbit large rotating part or abnormal clamping stagnation from the characteristic of a large-angle compensation satellite, and the satellite can keep three-axis ground control through the invention, thereby ensuring that the continuous work of the satellite service is not influenced.
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Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
FIG. 1 is a schematic flow chart of the present invention.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.
In the description of the present application, it is to be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience in describing the present application and simplifying the description, but do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and be operated, and thus, should not be construed as limiting the present application.
Aiming at the defects in the prior art, the invention provides a large-angle momentum compensation satellite three-stage instability judgment and control design method, which comprises the following steps:
step 1: the satellite is in the steady state control of the earth triaxial flywheel, the attitude of the satellite is out of tolerance, a first-level instability judgment condition is achieved, and the satellite autonomously enters a magnetic control mode;
step 2: the satellite is in the steady state control of the earth triaxial flywheel, the attitude of the satellite is out of tolerance, the satellite reaches the second-stage instability judgment condition, the satellite autonomously shifts into a steady state open loop mode, and further, the satellite enters a sun-oriented mode after being in an illumination area and losing the sun;
and step 3: the satellite is in the steady state control of the ground triaxial flywheel, the attitude of the satellite is out of tolerance, the three-level instability judgment condition is achieved, the satellite automatically enters an air injection control mode, and the control returns to the control of the ground triaxial flywheel after the control is stable.
In the step 1, the attitude of the satellite is out of tolerance, and when the attitude of the satellite continuously rolls for 120s, pitches and drifts for three axes, is more than 3 degrees, 4 degrees and 3 degrees, the satellite automatically switches to a magnetic control mode when the first-stage instability judgment condition is met;
in the step 2, the attitude of the ① satellite is out of tolerance and reaches continuous rolling for 30 seconds, pitching is achieved, when the yaw triaxial attitude is larger than 8 degrees, namely the condition of secondary instability judgment is met, the satellite automatically shifts into a steady state open loop mode, ② shifts into the steady state open loop mode, the satellite is in an uncontrolled state, and when ground processing is carried out, ③ if the satellite is in an illumination area and loses the sun, the satellite shifts into a sun-oriented control mode to ensure energy supply.
In the step 3, when the attitude of the satellite is out of tolerance and the continuous rolling, pitching and yawing triaxial attitudes are more than 12 degrees for 5s, the satellite automatically sends a thruster opening instruction and shifts to an air injection control mode.
In specific implementation, as shown in fig. 1, the invention provides an attitude three-level instability judgment and control method based on the characteristics of a large angle compensation satellite, so as to cope with the influence on the attitude of the whole satellite when the scanning rotation speed of the microwave radiometer antenna is abnormal. The steady state control of the satellite has three-gear control, and according to the satellite attitude in the ground triaxial flywheel steady state, three conditions are distinguished respectively:
the satellite is in the earth triaxial flywheel steady state control, and the satellite attitude is out of tolerance, reaches one-level unstability judgement condition, and the satellite independently enters the magnetic control mode, and the specific implementation mode is as follows:
and if the attitude of the satellite is out of tolerance and reaches continuous 120s rolling, pitching and yawing three-axis attitudes of more than 3 degrees, 4 degrees and 3 degrees, the satellite automatically switches to a magnetic control mode if the first-level instability judgment condition is met.
Entry conditions were as follows: and continuously rolling for 120s, pitching, and yawing for three axes, wherein the postures of the three axes are more than 3 degrees, 4 degrees and 3 degrees.
Attitude reference: star sensor, double-vector, gyro integral, infrared earth sensor and three-axis magnetometer
An executing mechanism: magnetic torquer
The control mode is as follows: magnetic control
Within 120s of attitude out-of-tolerance, the attitude reference is selected to be a star sensor, a double vector, a gyro integral and an infrared earth sensor in sequence, the priority of the default star sensor is the highest, and within 2 degrees/s of the attitude angular speed of the satellite, the output attitude of the star sensor is the most judgment and control basis. When the star sensor cannot output data or has stray light interference, double vectors are used as attitude references. The gyro integral is used as the attitude reference when the dual vectors are invalid.
The satellite is in the earth triaxial flywheel steady state control, and the satellite attitude is out of tolerance, reaches the second grade unstability judgement condition, and the satellite independently shifts into steady state open loop mode, and further, the satellite is in the illumination zone and after losing the sun, enters into the directional mode of the sun, and the concrete implementation mode is as follows:
① when the attitude of the satellite is out of tolerance and reaches continuous 30s rolling, pitching and yawing three-axis attitude more than 8 degrees, namely the condition for judging secondary instability is met, the satellite automatically switches into a steady state open loop mode,
②, the satellite is in an uncontrolled state after the satellite is switched to the steady state open loop mode, and the satellite is ready for ground processing.
③ if the satellite is in the illumination area and the sun is lost, to ensure energy supply, the control mode is switched to the sun-oriented control mode.
Entry conditions were as follows: and continuously rolling for 30s, pitching and yawing for three-axis postures more than 8 degrees.
Attitude reference: star sensor, double-vector, gyro integral, infrared earth sensor and sun sensor
An executing mechanism: is free of
The control mode is as follows: open loop
When the sun is sensitive to the sun and no signal is output, the sun is considered to be in the ground shadow area, the state of the uncontrolled moment output is continuously kept, otherwise, when the sun surface of the satellite is not illuminated, the sun sailboard cannot track the sun, the energy on the satellite cannot be guaranteed, and the satellite autonomously shifts to the sun-oriented control mode.
When the attitude of the satellite is out of tolerance and reaches continuous 5s rolling, pitching and yawing three-axis attitudes are more than 12 degrees, the satellite automatically sends a thruster opening instruction and switches to an air injection control mode, and the specific implementation mode is as follows:
entry conditions were as follows: and continuously rolling for 5s, pitching and yawing for three-axis postures of more than 12 degrees.
Attitude reference: infrared earth sensor or gyro integral
Angular velocity: output angular velocity of gyroscope
An executing mechanism: thruster, magnetic torquer and flywheel connected when returning to flywheel control
The control mode is as follows: jet control
Default to a rolling shaft and yaw shaft jet seal precession control algorithm and a pitch shaft inclined switch line control algorithm;
the control of a three-axis oblique switch line can be selected, the flywheel outputs a nominal rotating speed, the PI calculated value of the flywheel is assigned to zero, magnetic control is isolated, and the magnetic current output of the magnetic torquer is zero.
The three-axis attitude of the satellite continuously meets the condition that the three-axis attitude is less than 6 degrees in 60 seconds, and when the angular speed is less than 0.1 degree/s, the satellite exits the air injection control, is connected into a flywheel control mode, and is connected into a magnetic torquer.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
Claims (10)
1. A large angular momentum compensation satellite three-level instability judgment and control design method is characterized by comprising the following steps:
step 1: the satellite is in the steady state control of the earth triaxial flywheel, the satellite attitude reaches a primary instability state, and the satellite autonomously enters a magnetic control mode;
step 2: the satellite is in the steady state control of the earth triaxial flywheel, the satellite attitude reaches a secondary instability state, the satellite autonomously shifts into a steady state open loop mode, the satellite is in an uncontrolled state, and the satellite is ready for ground processing;
and step 3: the satellite is in the steady state control of the ground triaxial flywheel, the attitude of the satellite reaches a three-level instability state, the satellite autonomously enters an air injection control mode, and the steady state control returns to the steady state control of the ground triaxial flywheel after the control is stable.
2. The large angular momentum compensation satellite three-level instability judgment and control design method according to claim 1, wherein the first-level instability state is that the three-axis postures of a rolling axis, a pitching axis and a yawing axis in 120s of a satellite are respectively greater than 3 degrees, 4 degrees and 3 degrees.
3. The large angular momentum compensation satellite three-stage instability judgment and control design method as claimed in claim 1, wherein the secondary instability state is that the satellite has a rolling axis, a pitching axis and a yawing axis three-axis attitude greater than 8 degrees within 30 s.
4. The large angular momentum compensation satellite three-stage instability judgment and control design method as claimed in claim 1, wherein the three-stage instability state is that the three-axis attitude of the rolling axis, the pitching axis and the yawing axis of the satellite is greater than 12 degrees within 5 s.
5. The three-stage instability judgment and control design method for the large angular momentum compensated satellite according to claim 1, wherein in the step 1, the attitude reference of the satellite attitude is sequentially based on a star sensor, a double vector, a gyro integral and an infrared earth sensor, wherein the star sensor has the highest priority, and the output attitude of the star sensor is preferentially used as a judgment control basis within 2 °/s of the satellite attitude angular velocity; when the star sensor cannot output data or has stray light interference, double vectors are adopted as attitude references; when the dual vectors are invalid, the gyro integral is used as the attitude reference.
6. The large-angle momentum compensation satellite three-level instability judgment and control design method according to claim 1, wherein in the step 2, the to-be-processed ground is to utilize a sun sensor to judge the sun azimuth, and when no signal is output in the sun sensitivity of all quadrants, the sun sensor is judged to be in a ground shadow area, and the uncontrolled moment output state is continuously kept; otherwise, when the sun-exposed surface of the satellite is not illuminated, the satellite automatically switches into a sun-oriented control mode.
7. The large angular momentum compensation satellite three-level instability judgment and control design method as claimed in claim 1, wherein in the step 3, the attitude reference of the satellite attitude is based on an infrared earth sensor or a gyro integrator.
8. The large-angle momentum compensation satellite three-stage instability judgment and control design method as claimed in claim 4, wherein the rolling axis and the yaw axis are controlled by precession of a jet seal, and the pitch axis is controlled by an oblique switch line.
9. The large angular momentum compensation satellite three-level instability judgment and control design method according to claim 8, wherein the air injection control mode is to select three-axis oblique switch line control, the flywheel outputs a nominal rotating speed, the PI calculated value of the flywheel gives zero output, the magnetic control is isolated, and the magnetic torquer magnetic current output is zero.
10. The large angular momentum compensation satellite three-stage instability judgment and control design method according to claim 1, wherein the return to ground three-axis flywheel steady state control after the control is stable is that when the satellite three-axis attitude is less than 6 degrees and the angular velocity is less than 0.1 degrees/s within 60 seconds, the satellite exits the air injection control, accesses the flywheel control mode and accesses the magnetic torquer.
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| CN111532456A (en) * | 2020-05-15 | 2020-08-14 | 清华大学 | Control method and device |
| CN115783312A (en) * | 2022-12-07 | 2023-03-14 | 上海航天控制技术研究所 | All-day-area sun vector autonomous capture control method of analog sun sensor |
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