CN110589028A - Autonomous mode switching method for abnormal satellite attitude maneuver - Google Patents
Autonomous mode switching method for abnormal satellite attitude maneuver Download PDFInfo
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- CN110589028A CN110589028A CN201910933636.4A CN201910933636A CN110589028A CN 110589028 A CN110589028 A CN 110589028A CN 201910933636 A CN201910933636 A CN 201910933636A CN 110589028 A CN110589028 A CN 110589028A
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Classifications
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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
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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
- B64G1/28—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect
- B64G1/286—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect using control momentum gyroscopes (CMGs)
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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
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
- B64G1/361—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using star sensors
Abstract
The invention discloses an autonomous mode switching method for abnormal satellite attitude maneuver, which comprises the following steps: when the satellite maneuvers from a long-term steady state to any other attitude, the satellite automatically judges whether the attitude meets the condition of entering a full attitude capture mode; if not, the satellite autonomously judges whether the staying time under the non-long-term steady state or the solar cell array state is normal, if not, the satellite is switched into a full-attitude capturing mode or maneuvered back to the long-term steady state, if so, the satellite judges whether an attitude maneuvering instruction is received, and if so, the satellite switches the state according to the maneuvering condition of the attitude maneuvering instruction; if not, the satellite remains in a non-long term steady state. The advantages are that: by setting reasonable attitude maneuver judgment conditions and autonomous mode switching conditions, on the premise of realizing the in-orbit attitude maneuver function of the satellite, the safety of a satellite power supply is ensured, and the safety of the in-orbit autonomous attitude control of the satellite can be improved.
Description
Technical Field
The invention relates to the technical field of satellite attitude control, in particular to an autonomous mode switching method during abnormal satellite attitude maneuver.
Background
For a satellite with any attitude maneuver, the safety of the attitude and the power supply of the satellite needs to be ensured while the attitude maneuver function is realized. The existing attitude agility maneuver control method relates to a mode switching processing method less when the attitude maneuver is abnormal. However, for satellite attitude control, both attitude stabilization control of the satellite and the safety of the whole satellite power supply need to be realized, so that the autonomous mode switching method for the abnormal maneuvering of the satellite attitude has certain necessity.
At present, no description or report of similar technologies in the aspect is found, and similar data at home and abroad is not collected. Therefore, an autonomous mode switching method for satellite attitude maneuver abnormity is urgently needed, which can simultaneously ensure the attitude stability control of the satellite and the safety of the whole satellite power supply.
Disclosure of Invention
The invention aims to provide an autonomous mode switching method for abnormal satellite attitude maneuver, which ensures the energy safety of a satellite and improves the safety of the satellite in-orbit autonomous attitude control on the premise of realizing the capability of the satellite in-orbit attitude maneuver by setting reasonable attitude maneuver judgment conditions and autonomous mode switching conditions.
In order to achieve the purpose, the invention is realized by the following technical scheme:
an autonomous mode switching method for satellite attitude maneuver abnormity, the method comprises:
s1, when the satellite maneuvers from a long-term steady state to any other attitude, the satellite autonomously judges whether the attitude meets the condition of entering a full attitude capture mode; if yes, the satellite is switched into a full-attitude capturing mode, and the autonomous mode switching process is exited; if not, go to step S2;
s2, after the satellite maneuvers to any other attitude, when the satellite autonomously judges that the satellite stays in the non-long-term steady state for a long time or the state of the solar cell array is abnormal, turning to the step S3;
s3, judging a first attitude maneuver condition TjdWhether or not it is satisfied; if the satellite is in the stable state, the satellite is in the stable state for a long time, and the autonomous mode switching process is exited; if not, the satellite is switched into a full attitude acquisition mode, and the autonomous mode switching process is exited.
Preferably, in the step S2, after the satellite maneuvers to any other attitude, when the satellite autonomously determines that the satellite stays in the non-long-term steady state or the solar cell array state is normal, the process goes to step S4;
s4, judging whether the satellite receives an attitude maneuver instruction; if the attitude maneuver instruction is received, the satellite switches the state according to the maneuver condition of the attitude maneuver instruction; if not, the satellite remains in a non-long-term steady state and exits the autonomous mode switching process.
Preferably, the maneuvering condition switching state of the satellite according to the attitude maneuvering instruction in step S4 is specifically:
s41, judging second attitude maneuver condition SjdWhether or not it is satisfied; if so, the satellite carries out attitude maneuver according to the received specific attitude maneuver instruction and exits the autonomous switching process; if not, go to step S42;
s42, judging a first attitude maneuver condition TjdWhether or not it is satisfied; if so, the satellite autonomously maneuvers back to a long-term steady state and exits the autonomous mode switching process; if not, the satellite is switched to a full attitude acquisition mode, and the autonomous mode switching process is exited.
Preferably, the normality in step S2 is specifically: the satellite reaches one of three conditions as a normal state under the non-long-term steady state, and the three conditions are specifically as follows: and the staying time exceeds a set value, the solar cell array faults or the solar cell array driving circuits are all cut off.
Preferably, the first attitude maneuver condition TjdThe determination conditions were as follows:
1) the gyroscope combination can measure the three-axis inertial angular velocity;
2) the track parameters are valid;
3) a star sensor correction mode;
if the conditions 1) to 3) are satisfied simultaneously, T is setjd1, the first attitude maneuver condition is satisfied, otherwise, T is setjdThe first attitude maneuver condition is not satisfied.
Preferably, the second attitude maneuver condition SjdThe determination conditions of (1) are as follows:
1) judging by an actuating mechanism, controlling the moment gyros of the access system to at least perform attitude maneuver, and controlling the moment gyros to synthesize angular momentum abs (H) less than or equal to Hjd，HjdSynthesizing the maximum value of angular momentum according to the control moment gyro group in a steady state;
2) the battery array driving circuit is normal in state, and the driving mechanism is normal;
3) at least one star sensor is in a normal state;
4) the gyroscope combination can measure the three-axis inertial angular velocity;
5) the track parameters are valid;
6) a star sensor correction mode;
when the above 6 conditions are simultaneously satisfied, S is setjdIf the second attitude maneuver condition is satisfied, otherwise, S is setjdAnd 0, the second attitude maneuver condition is not satisfied.
Compared with the prior art, the invention has the following advantages: according to the satellite power supply system and the satellite power supply method, the satellite autonomously realizes the maximum maneuvering efficiency of the satellite according to constraint conditions such as the side-looking working state of the satellite, the sensor and the actuating mechanism, the working state of the solar cell array, the on-satellite task requirement and the like, the satellite can autonomously enter a full-attitude capturing mode to realize the safety of the on-satellite power supply when the attitude of the satellite is abnormal, the beneficial effect of improving the on-orbit safety of the satellite under the condition that the satellite frequently performs attitude maneuvering is achieved, and the satellite power supply system and the satellite power supply method can be applied to the.
Drawings
Fig. 1 is a flowchart of an autonomous mode switching method for abnormal satellite attitude maneuver according to the present invention.
Detailed Description
The present invention will now be further described by way of the following detailed description of a preferred embodiment thereof, taken in conjunction with the accompanying drawings.
For the satellites which fly obliquely from the left side view to the right side view, for example, the satellites are in a right side view state for a long time (namely, in a long-term steady state) when in orbit operation, but any attitude maneuvering capability is required. When the satellite is moved from the right side view state to any other non-right side view state, the energy safety problem of the satellite in the non-right side view state needs to be considered. As shown in fig. 1, the autonomous mode switching method for abnormal satellite attitude maneuver adopted by the present invention can realize the safety of satellite attitude and energy. The specific method comprises the following steps:
s1, when the satellite maneuvers from the right side view state to any other attitude, in order to ensure the safety of a power supply, the satellite needs to autonomously judge whether the attitude of the satellite meets the condition of entering a full attitude capture mode; if yes, the satellite is switched into a full-attitude capturing mode, and the autonomous mode switching process is exited; if not, go to step S2.
S2, when the satellite maneuvers to any other attitude, the satellite autonomously judges whether the stay time or the solar cell array state is normal under the non-right side view state; if not, go to step S3; if so, go to step S4.
Wherein the step S2 is normal specifically: the satellite is normal when reaching one of three conditions under the non-right side view state, wherein the three conditions are as follows: and the staying time exceeds a set value, the solar cell array faults or the solar cell array driving circuits are all cut off.
S3, judging a first attitude maneuver condition TjdWhether or not it is satisfied; if the current state is met, the satellite flexibly returns to the right side view state and exits the autonomous mode switching process; if not, the satellite is switched into a full attitude acquisition mode, and the autonomous mode switching process is exited.
S4, judging whether the satellite receives an attitude maneuver instruction; if the attitude maneuver instruction is received, the satellite switches the state according to the maneuver condition of the attitude maneuver instruction; if not, the satellite remains in the non-right-side view state and exits the autonomous mode switching process.
The maneuvering condition conversion state of the satellite according to the attitude maneuvering instruction is specifically as follows:
s41, judging second attitude maneuver condition SjdWhether or not it is satisfied; if so, the satellite carries out attitude maneuver according to the received specific attitude maneuver instruction and exits the autonomous switching process; if not, go to step S42;
s42, judging a first attitude maneuver condition TjdWhether or not it is satisfied; if so, the satellite autonomously moves back to the right side view state and exits the autonomous mode switching process; if not, the satellite is switched to a full attitude acquisition mode, and the autonomous mode switching process is exited.
The first attitude maneuver condition TjdAnd a second attitude maneuver condition SjdThe determination method of (2) is as follows:
1) judging by an actuating mechanism, in the embodiment, the satellite is in a pentagonal pyramid control moment gyro group, at least 4 control moment gyros are accessed into the system, the control moment gyro of the accessed system can perform attitude maneuver at least, and the synthetic angular momentum abs (H) of the control moment gyro group is less than or equal to Hjd，HjdFor synthesizing the maximum value of angular momentum, H, from the group of control moment gyros at steady statejdThe range is 0 to 50 Nms;
2) the battery array driving circuit is normal in state, and the driving mechanism is normal;
3) at least one star sensor is in a normal state (power-up and non-fault state);
4) the gyroscope combination can measure the three-axis inertial angular velocity;
5) the track parameters are valid;
6) a star sensor correction mode;
when the above 6 conditions are simultaneously satisfied, S is setjdThat is, the second attitude maneuver condition is satisfied, otherwise S is setjdIf not, the second attitude maneuver condition is not satisfied;
if the conditions 4) to 6) are satisfied simultaneously, T is setjdIf the first attitude maneuver condition is satisfied, otherwise, T is setjdAnd 0, namely the first attitude maneuver condition is not satisfied.
In summary, according to the autonomous mode switching method for the satellite attitude maneuver abnormity, by setting the reasonable attitude maneuver judgment condition and the autonomous mode switching condition, on the premise of realizing the satellite in-orbit attitude maneuver capability, the energy safety of the satellite is ensured, the safety of the satellite in-orbit autonomous attitude control is improved, the beneficial effect of improving the in-orbit safety of the satellite under the condition that the satellite frequently performs the attitude maneuver is achieved, and the method can be applied to the satellite with the agile maneuver function.
While the present invention has been described in detail with reference to the preferred embodiments, it should be understood that the above description should not be taken as limiting the invention. Various modifications and alterations to this invention will become apparent to those skilled in the art upon reading the foregoing description. Accordingly, the scope of the invention should be determined from the following claims.
Claims (6)
1. An autonomous mode switching method for satellite attitude maneuver anomaly, the method comprising:
s1, when the satellite maneuvers from a long-term steady state to any other attitude, the satellite autonomously judges whether the attitude meets the condition of entering a full attitude capture mode; if yes, the satellite is switched into a full-attitude capturing mode, and the autonomous mode switching process is exited; if not, go to step S2;
s2, after the satellite maneuvers to any other attitude, when the satellite autonomously judges that the satellite stays in the non-long-term steady state for a long time or the state of the solar cell array is abnormal, turning to the step S3;
s3, judging a first attitude maneuver condition TjdWhether or not it is satisfied; if the satellite is in the stable state, the satellite is in the stable state for a long time, and the autonomous mode switching process is exited; if not, the satellite is switched into a full attitude acquisition mode, and the autonomous mode switching process is exited.
2. The method of autonomous mode switching in case of satellite attitude maneuver abnormality according to claim 1,
in the step S2, after the satellite maneuvers to any other attitude, when the satellite autonomously determines that the satellite stays in the non-long-term steady state for a certain time or the solar cell array state is normal, the process goes to step S4;
s4, judging whether the satellite receives an attitude maneuver instruction; if the attitude maneuver instruction is received, the satellite switches the state according to the maneuver condition of the attitude maneuver instruction; if not, the satellite remains in a non-long-term steady state and exits the autonomous mode switching process.
3. The method for autonomous mode switching during an abnormal satellite attitude maneuver according to claim 2, wherein the satellite in step S4 switches states according to the maneuver condition of the attitude maneuver instruction, specifically:
s41, judging second attitude maneuver condition SjdWhether or not it is satisfied; if so, the satellite carries out attitude maneuver according to the received specific attitude maneuver instruction and exits the autonomous switching process; if not, go to step S42;
s42, judging a first attitude maneuver condition TjdWhether or not it is satisfied; if so, the satellite autonomously maneuvers back to a long-term steady state and exits the autonomous mode switching process; if not, the satellite is switched to a full attitude acquisition mode, and the autonomous mode switching process is exited.
4. The method for autonomous mode switching at the time of an abnormality of a satellite attitude maneuver according to claim 2,
the step S2 is normally embodied as: the satellite reaches one of three conditions as a normal state under the non-long-term steady state, and the three conditions are specifically as follows: and the staying time exceeds a set value, the solar cell array faults or the solar cell array driving circuits are all cut off.
5. The autonomous mode switching method at the time of abnormality of satellite attitude maneuver according to claim 1 or 2,
the first attitude maneuver condition TjdThe determination conditions were as follows:
1) the gyroscope combination can measure the three-axis inertial angular velocity;
2) the track parameters are valid;
3) a star sensor correction mode;
if the conditions 1) to 3) are satisfied simultaneously, T is setjd1, the first attitude maneuver condition is satisfied, otherwise, T is setjdThe first attitude maneuver condition is not satisfied.
6. The method of autonomous mode switching in case of satellite attitude maneuver according to claim 3,
the second attitude maneuver condition SjdThe determination conditions of (1) are as follows:
1) judging by an actuating mechanism, controlling the moment gyros of the access system to at least perform attitude maneuver, and controlling the moment gyros to synthesize angular momentum abs (H) less than or equal to Hjd，HjdSynthesizing the maximum value of angular momentum according to the control moment gyro group in a steady state;
2) the battery array driving circuit is normal in state, and the driving mechanism is normal;
3) at least one star sensor is in a normal state;
4) the gyroscope combination can measure the three-axis inertial angular velocity;
5) the track parameters are valid;
6) a star sensor correction mode;
when the above 6 conditions are simultaneously satisfied, S is setjdIf the second attitude maneuver condition is satisfied, otherwise, S is setjdAnd 0, the second attitude maneuver condition is not satisfied.
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