CN116692029A - Attitude and orbit control method and system for L2 aircraft adapting to Lagrange points - Google Patents
Attitude and orbit control method and system for L2 aircraft adapting to Lagrange points Download PDFInfo
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Abstract
The invention provides a attitude and orbit control method for an L2 aircraft adapting to a Lagrange point, which comprises the following steps: s1, after a satellite and an arrow are separated, angular speed and attitude deviation caused by the separation of the satellite and the arrow are rapidly eliminated, and sun orientation of a solar cell array is realized; s2, controlling to realize the transfer of the satellite from the GTO to the L2 orbit; s3, after the scientific cabin and the propulsion cabin of the satellite are separated, the separation angular speed and the attitude deviation are eliminated, and the sun alignment of the satellite solar cell array is realized again; s4, in the satellite task stage, carrying out attitude measurement and attitude control according to load requirements; s5, during the orbit running of the satellite, orbit maintaining control is carried out according to the satellite orbit drifting condition. The method can realize accurate control of the L2 aircraft in different running states.
Description
Technical Field
The invention relates to the technical field of attitude control of an L2 aircraft, in particular to an attitude and orbit control method and system for an L2 aircraft adapting to a Lagrange point.
Background
The satellite attitude control system is an important component of the whole satellite system. The attitude control system is an embedded real-time control system. Unlike common control system, the system has complex structure and high reliability requirement. The attitude control system specifically comprises satellite body dynamics analysis, an attitude sensor, an attitude determination algorithm, an attitude control algorithm, an executing mechanism and the like. The design objective of the attitude control system software is the same as all software designs, so as to meet the demands to the greatest extent under the constraint limits of time and various resource environments. The on-orbit running environment of the attitude control system is continuously changed, the software design is not only required to complete the function and performance requirements of the system, but also is required to reach a high level in the aspects of comprehensibility, easy maintenance and expandability so as to meet the requirement of continuous change of the control of the on-board system.
The existing attitude control method generally uses a fuzzy self-adaptive PI controller to send pulse control signals to control the flywheel speed so as to control the satellite flight attitude, or designs a controller based on PD control, and only uses a magnetic torquer to complete attitude control of the pico satellite, which are simple and convenient control methods with limited precision. In the existing satellite attitude control method, the difference of the designs of the controllers of the satellites at different stages is not considered, and only a control strategy of one attitude control execution component is described, and the cooperative work among a plurality of attitude control execution components is not considered. The attitude control system has the characteristics of complex structure, severe working environment, unknown interference and more uncertain factors, and is one of the subsystems with the largest faults.
It is therefore necessary to provide a control method of a system that achieves high-precision attitude control of a satellite in a plurality of operating states and improves control stability.
Disclosure of Invention
Technical problem to be solved
Aiming at the defects in the prior art, the invention provides a attitude and orbit control method and system for an L2 aircraft adapting to a Lagrange point, and the method and the system can realize accurate control of the L2 aircraft under different running states.
Technical proposal
In order to achieve the above purpose, the invention is realized by the following technical scheme:
the invention provides a attitude and orbit control method for an L2 aircraft adapting to a Lagrange point, which comprises the following steps:
s1, after a satellite and an arrow are separated, angular speed and attitude deviation caused by the separation of the satellite and the arrow are rapidly eliminated, and sun orientation of a solar cell array is realized;
s2, controlling to realize the transfer of the satellite from the GTO to the L2 orbit;
s3, after the scientific cabin and the propulsion cabin of the satellite are separated, the separation angular speed and the attitude deviation are eliminated, and the sun alignment of the satellite solar cell array is realized again;
s4, in the satellite task stage, carrying out attitude measurement and attitude control according to load requirements;
s5, during the orbit running of the satellite, orbit maintaining control is carried out according to the satellite orbit drifting condition.
Further, the satellite positioning control system also comprises a safety mode 1 and a safety mode 2, wherein the safety mode 1 is used for adjusting a control scheme of the satellite when the satellite fails on the satellite and the satellite attitude is stable in a normal mode, damping the angular speed of the satellite by utilizing chemical pushing and controlling the satellite to realize daily orientation; the safety mode 2 is used for adjusting a control scheme of the scientific cabin when the scientific cabin of the satellite fails and the normal mode cannot guarantee stable posture, damping the angular speed of the scientific cabin by using the cold air micro-thruster, and controlling the scientific cabin to realize sun orientation.
Further, the step S1 specifically includes: after the satellites and the arrows are separated, the satellites automatically perform a push rate damping mode, the push rate damping mode is used for damping satellite angular velocity deviation caused by the separation of the satellites and the arrows, damping the triaxial angular velocity to a set range, and utilizing the collected solar vector information, and exerting control to realize solar orientation of the solar sailboard.
Further, step S2 specifically includes: in the satellite orbit transfer process, an orbit-changing attitude maneuver mode is entered before orbit maneuver, and the orbit-changing attitude maneuver mode is used for adjusting the satellite attitude to prepare for orbit maneuver; after the satellite finishes attitude maneuver, meeting the attitude pointing requirement, igniting an orbit control engine, and performing orbit maneuver; after the satellite enters a working orbit after orbit transfer is finished, the satellite enters a gesture capturing mode, and triaxial stable pointing control is realized.
Further, the step S3 specifically includes: after the satellite enters a preset orbit, a ground instruction enters a propulsion cabin separation mode, and the separation of a propulsion cabin and a scientific cabin of the satellite is implemented; when the separation of the scientific cabin and the propulsion cabin of the satellite is finished, implementing the speed damping of the scientific cabin, and damping the angular speed deviation caused by the separation; and (5) utilizing the collected solar vector information to exert control to realize the sun orientation of the solar sailboard.
Based on the same invention conception, the invention also provides a attitude and orbit control system of the L2 aircraft adapting to the Lagrangian point, wherein the attitude and orbit control system is used for realizing the method of any one of the above, and comprises an attitude sensor, a controller and an executing mechanism, wherein the attitude sensor is used for measuring scientific load and obtaining an attitude determination scheme; the controller is integrated in the satellite computer and is used for collecting data of the attitude sensor and analyzing and processing the data to obtain satellite attitude; the executing mechanism is used for receiving a control instruction sent by the controller to realize attitude and orbit control of the satellite.
Further, the attitude sensor comprises a star sensor, an optical fiber inertial unit, an analog sun sensor and a load camera, wherein the star sensor is an attitude measurement component, the optical fiber inertial unit is used for directly measuring the attitude angular speed and acceleration information under the satellite inertial system, the analog sun sensor is used for measuring the sun vector of the position of the satellite, and the load camera is used for shooting images.
Further, the actuator includes a cold air micro-thruster, a chemical thruster, an electric thruster, and a reaction flywheel.
Advantageous effects
The invention provides a attitude orbit control method and system for an L2 aircraft adapting to a Lagrange point. Therefore, different attitude determination and control modes are provided in different task stages, and the attitude control requirements of each stage of the satellite are met; meanwhile, a multi-working mode and a software and hardware safety mode are designed, and the reconfigurability and the reliability of the software system are improved.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below. It is evident that the drawings in the following description are only some embodiments of the present invention and that other drawings may be obtained from these drawings without inventive effort for a person of ordinary skill in the art.
FIG. 1 is a schematic diagram of a method for attitude and orbit control for an L2 aircraft adapted to a Lagrangian point according to one embodiment of the present invention;
FIG. 2 is a schematic diagram illustrating an embodiment of a attitude and orbit control system for an aircraft adapted to the Lagrange point L2;
fig. 3 is a schematic diagram of a attitude and orbit control system adapted to a lagrangian point L2 aircraft according to an embodiment of the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present invention more clear, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. It will be apparent that the described embodiments are some, but not all, embodiments of the invention. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
Referring to fig. 1, an embodiment of the present invention provides a attitude and orbit control method for an aircraft adapted to a lagrangian point L2, comprising the steps of:
s1, after a satellite and an arrow are separated, angular speed and attitude deviation caused by the separation of the satellite and the arrow are rapidly eliminated, and sun orientation of a solar cell array is realized;
s2, controlling to realize the transfer of the satellite from the GTO to the L2 orbit;
s3, after the scientific cabin and the propulsion cabin of the satellite are separated, the separation angular speed and the attitude deviation are eliminated, and the sun alignment of the satellite solar cell array is realized again;
s4, in the satellite task stage, carrying out attitude measurement and attitude control according to load requirements;
s5, during the orbit running of the satellite, orbit maintaining control is carried out according to the satellite orbit drifting condition.
Specifically, the working mode of the method is divided into an orbit entering stage (after a satellite and an arrow are separated), an orbit transferring stage, a scientific cabin orbit entering stage, a task stage and a safety mode.
In this embodiment, step S1 specifically includes: after the satellites and the arrows are separated, the satellites automatically perform a push rate damping mode, the push rate damping mode is used for damping satellite angular velocity deviation caused by the separation of the satellites and the arrows, damping the triaxial angular velocity to a set range, and utilizing the collected solar vector information, and exerting control to realize solar orientation of the solar sailboard.
In this embodiment, step S2 specifically includes: in the satellite orbit transfer process, an orbit-changing attitude maneuver mode is entered before orbit maneuver, and the orbit-changing attitude maneuver mode is used for adjusting the satellite attitude to prepare for orbit maneuver; after the satellite finishes attitude maneuver, meeting the attitude pointing requirement, igniting an orbit control engine, and performing orbit maneuver; after the satellite enters a working orbit after orbit transfer is finished, the satellite enters a gesture capturing mode, and triaxial stable pointing control is realized.
In this embodiment, step S3 specifically includes: after the satellite enters a preset orbit, a ground instruction enters a propulsion cabin separation mode, and the separation of a propulsion cabin and a scientific cabin of the satellite is implemented; when the separation of the scientific cabin and the propulsion cabin of the satellite is finished, implementing the speed damping of the scientific cabin, and damping the angular speed deviation caused by the separation; and (5) utilizing the collected solar vector information to exert control to realize the sun orientation of the solar sailboard.
Based on the same invention conception, the invention also provides a attitude and orbit control system of the L2 aircraft adapting to the Lagrangian point, wherein the attitude and orbit control system is used for realizing the method of any one of the above, and comprises an attitude sensor, a controller and an executing mechanism, wherein the attitude sensor is used for measuring scientific load and obtaining an attitude determination scheme; the controller is integrated in the satellite computer and is used for collecting data of the attitude sensor and analyzing and processing the data to obtain satellite attitude; the executing mechanism is used for receiving a control instruction sent by the controller to realize attitude and orbit control of the satellite.
In this embodiment, the attitude sensor includes a star sensor, an optical fiber inertial unit, an analog sun sensor and a load camera, where the star sensor is an attitude measurement component, the optical fiber inertial unit is used for directly measuring attitude angular velocity and acceleration information under a satellite inertial system, the analog sun sensor is used for measuring a sun vector of a position where a satellite is located, and the load camera is used for capturing an image.
In this embodiment, the actuator includes a cold gas micro-thruster, a chemical thruster, an electric thruster, and a reaction flywheel.
Referring to fig. 2, the system operates in the following mode:
1) Chemical push rate damping: after the satellites and the satellites are separated, the satellites autonomously perform a push rate damping mode, and the function of the push rate damping mode is to damp satellite angular velocity deviation caused by the separation of the satellites and the satellites, so that triaxial angular velocity is damped to a smaller range, and a foundation is established for sun orientation.
2) Sun orientation 1: the sun orientation 1 is mainly to determine a sun vector by utilizing sun sensor information or star sensor information, and the sun orientation of a solar sailboard is realized by using a chemical thruster to exert control, so that energy supply is ensured.
3) Orbital transfer attitude maneuver: in the satellite orbit transfer process, an orbit-changing attitude maneuver mode is needed to be entered before orbit maneuver, and the satellite attitude is adjusted to prepare for the orbit maneuver.
4) Track maneuvering: after the satellite finishes attitude maneuver, the requirement of attitude pointing is met, and the orbit control engine is ignited to perform orbit maneuver.
5) Gesture capturing: after the satellite orbit transfer is finished and the satellite enters the working orbit, the satellite enters a gesture capturing mode, and three-axis stable pointing control is realized.
6) Separation of the propulsion cabin: after the satellite enters a preset orbit, the satellite enters a propulsion cabin separation mode by a ground instruction, and separation of the propulsion cabin and the scientific cabin is implemented.
7) And (3) micro-thrust rate damping, namely after separation of the satellite scientific cabin and the propulsion cabin is finished, using a cold air micro-thruster to implement rate damping of the scientific cabin, and damping angular speed deviation caused by separation.
8) Sun orientation 2: the sun orientation 2 is mainly to measure sun vectors by using sun sensor information or star sensor information, and control is exerted by using a cold air micro thruster to realize sun orientation of a solar sailboard and ensure energy supply.
9) Guidance law tracking: the attitude control performs tracking control according to the guiding law of the upper stream, and provides attitude guarantee for the work of scientific load.
10 Rail controlled posture adjustment: the orbit control posture adjustment mode provides posture guarantee for maintaining the orbit of the satellite.
11 Rail maintenance: the orbit maintenance mode is to control the orbit of the satellite according to the task requirements.
12 Security mode 1): the function of the safety mode 1 is to adjust a satellite control scheme when serious faults occur on the satellite and the satellite attitude stability cannot be guaranteed in a normal mode, damp the angular velocity of the satellite by utilizing chemical pushing, and simultaneously control the satellite to realize sun orientation so as to guarantee the basic supply of energy on the satellite.
13 Security mode 2): the safety mode 2 has the functions of adjusting a control scheme of the scientific cabin when the scientific cabin has serious faults and the normal mode cannot ensure stable posture, damping the angular speed of the scientific cabin by using the cold air micro-thruster, controlling the plain scientific cabin to realize daily orientation and ensuring the basic supply of energy sources of the scientific cabin.
Referring to fig. 3, the attitude and orbit control system is composed of an attitude sensor, an actuating mechanism and a controller. The satellite adopts a star-sensitive and gyro high-precision attitude determination scheme, a sun-sensitive and sun-oriented attitude determination scheme and a scientific load measurement result-based attitude determination scheme, and the attitude sensor comprises: star sensor, light inertial measurement unit, analog sun sensor and load camera. The scheme that the satellite science cabin and the propulsion cabin adopt chemical thrusters to carry out attitude and orbit control, after the propulsion cabin is separated, the science cabin adopts the scheme that the cold air micro-thruster dampens, the reaction flywheel is motorized, the cold air thruster carries out high-precision attitude control, and the electric thruster carries out orbit control, and the executing mechanism comprises: cold air micro-thrusters, chemical thrusters, electric thrusters, and reaction flywheels. The controllers of the gestures and the orbits are integrated in a satellite computer, data of the gesture sensors are collected, and analysis and processing are carried out to obtain satellite gestures; then, according to the satellite attitude information measured by the sensor, control is applied to achieve the control purpose; sending a control instruction to a control part; and the functions of a gesture track control center such as gesture track control flow and management of working modes are realized. It mainly carries the algorithm software of the attitude and orbit control system.
In the embodiment, the star sensor is the component with the highest precision in the conventional attitude measurement components, the measurement precision can reach the order of magnitude of an angle second, and the star sensor is widely applied to high-precision attitude determination of satellites. For attitude and orbit control subsystem, the star sensor mainly comprises: 1) In a guidance law tracking mode, accurate attitude measurement is carried out, and the attitude measurement accuracy requirement required by load work is ensured; 2) And the combined gyroscope integral attitude determination meets the attitude determination/solar angle determination requirements under various working modes. According to task analysis, the measurement error of the three axes of the star sensor is required to be generally better than 20urad, the current star sensor cannot fully meet the requirement, and the measurement error around the optical axis is large, so that a double-star-sensor gesture determination mode is required to be used for enabling the gesture measurement error to meet the task requirement. According to the radiation-resistant requirement of the satellite sun orbit, the selected star sensor must be capable of bearing the radiation condition of a high orbit, and meanwhile, the maturity, flight experience, product weight reduction requirement and maturity of a product purchasing channel of the star sensor are considered, so that the star sensor used by the navigation satellite is inherited. The star sensor is applied to a satellite transfer orbit phase and a load working phase. In order to meet the requirement of load measurement precision, the optical axis of the star sensor is required to be parallel to the load optical axis, meanwhile, the arbitrary nature of the gesture in the gesture capturing process is considered, the thruster control algorithm is relatively dependent on gesture information input, 3 star sensors are configured and arranged in the XOY plane of the platform cabin in order to ensure that the star sensor is effective in gesture determination, the star sensor A optical axis points to +X axis deviation +Y axis 30 degrees, the star sensor B optical axis points to +X axis deviation-Y axis 30 degrees, and the star sensor C optical axis points to-X axis, so that at least two star sensors can be used in the load working period.
In the embodiment, the optical fiber inertial measurement unit is used for directly measuring the attitude angular speed and acceleration information under the satellite inertial system, and has the advantages of all weather and no view field constraint. For a satellite attitude and orbit control subsystem, the functions of the optical fiber inertial measurement unit mainly comprise: a. providing angular velocity information for rate damping during an on-track phase, during a track control, and in a failure mode; b. the method is used for gyro integral attitude determination, determines the three-axis attitude or solar angle of a satellite, is used as the supplement of satellite-sensitive attitude determination, and ensures the continuity of attitude measurement; c. the method is used for high-precision attitude control, and an angular velocity item is introduced to improve control performance; d. the attitude determination is combined with the star sensor, so that the attitude measurement precision is improved, and the measurement precision index requirement is ensured; e. for measuring acceleration information of the satellite during orbit determination. Aiming at the satellite orbit change characteristics, a triaxial quartz accelerometer is added in the fiber optic gyroscope. On one hand, the selection of the gyroscope considers performance indexes; on the other hand, the reliability, maturity and inheritance of the product are considered. The engineering center and the optical fiber gyro developed in the nine-yard times of aerospace have long-term cooperation, and the optical fiber gyro product is successfully and widely applied to a series of satellites in the center, such as an innovation 03 satellite, a quantum satellite, a carbon satellite, a navigation satellite and the like, and has good and stable in-orbit flight performance. The three-axis inertial measurement unit with the time-consuming photoelectric device is replaced by a high-grade component suitable for deep space aiming at the characteristic of the satellite sun orbit.
In this embodiment, the analog solar sensor is used to measure the solar vector of the satellite, and ideally, the full celestial sphere field of view can be obtained in the sun-illuminated area through the cooperation of multiple measuring surfaces. For the attitude and orbit control subsystem, the sun sensor mainly has the following functions: the sun position is determined as control input information for the orientation of the sun. The main function of the analog sun sensor is to realize the sun orientation of the satellite, and the measurement accuracy can meet the requirement by 1 degree in consideration of the low control accuracy required for the sun orientation. The analog sun sensor is simple, reliable and mature, and has rich on-orbit experience. The method is always applied to all satellite models in the center, and the on-orbit characteristic of the method is fully known. The implementation approach of the lineage center is available, and the satellite orbit characteristic is replaced by a high-grade component suitable for deep space. The analog sun sensor is limited by a view field, and in order to achieve gesture capture under any state of a sunlight area, a 6-probe mode is adopted to respectively measure sun vectors in six directions of the star, so that a full celestial sphere view field is achieved. Because 6 probes cover all 6 directions of the three-dimensional space, the field of view of each measuring surface can be mutually compensated, and the field of view of the whole celestial sphere can be realized, and the solar sensor can be ensured to be effective all the time under any posture only by requiring that the effective field of view of each surface is larger than +/-45 degrees and +/-45 degrees. The installation mode of the propulsion cabin and the scientific cabin is considered, the propulsion cabin needs to be separated from the scientific cabin, meanwhile, the redundancy of the system is considered, two simulated solar sensors are installed on a solar sailboard of the scientific cabin and are mutually backed up, and one solar sensor is installed on the other five surfaces, so that the measurement of six surfaces can be realized. Because the installation of the propulsion cabin and the scientific cabin can cause the five sides of the scientific cabin to be shielded by the propulsion cabin, only one side of the sun sensor installed on the scientific cabin is available before the propulsion cabin is separated. In order to ensure the availability of satellite solar vectors, a solar sensor is arranged on the other five surfaces of the propulsion cabin respectively so as to meet the field requirement of all the days before separation.
In this embodiment, the reaction flywheel is mainly used for satellite attitude maneuver, the satellite moment of inertia is calculated according to 2000kgm2, the maneuver is assumed to be an acceleration-constant speed-deceleration process, the moment of control of the reaction flywheel is required to be greater than 0.0175Nm, and the angular momentum is required to be greater than 2Nms according to the maximum angular velocity of 0.03 DEG/s and the acceleration of 0.0005 DEG/s 2 in consideration of the attitude stabilization time. Considering a certain margin, 4 reaction flywheels of 4.0Nms and 0.1Nm are selected.
In this embodiment, the satellite science pod is equipped with a cold air micro-thruster as an actuator of the attitude and orbit control system, and for the satellite attitude and orbit control subsystem, the functions of the cold air micro-thruster mainly include: after the propulsion cabin is separated, the speed damping of the cold air micro thruster is carried out, and the attitude angular speed of the separated scientific cabin and propulsion cabin is damped; in the sun orientation, guidance rate tracking control and orbit maintenance stage, the three-axis attitude adjustment and stable control of the satellite are realized, and the attitude control requirements under each working mode are met. The cold air micro-thruster is used as an executing mechanism, two cold air micro-thrusters are adopted, each of the two cold air micro-thrusters is provided with 6 main and 6 standby configuration schemes, wherein the 20mN cold air micro-thruster is used for speed damping, reactive flywheel unloading and attitude control during track maintenance, and the 50 mu N cold air micro-thruster is used for high-precision attitude control.
In this embodiment, the satellite science pod is equipped with a hall electric thruster as an actuator for orbit maintenance after separation of the propulsion pod. Propellant estimation: the time required for one track maintenance was 2/1000 (40/2000/1000), 3600=27.8 hours, and the fuel consumption was 40/1000×27.8×3600/2000/9.81= 0.2040kg. The total fuel required to maintain the track over the 5-year lifetime was 0.2040 × (5×365 ∈90) = 4.1367kg, and the total fuel required to maintain the track was 6.2051kg, taking into account the 0.5-fold margin. The 40mN Hall electric thruster developed by 801 is selected according to the constraint of satellite power and the like by referring to the configuration scheme of an electric propulsion system for detecting orbit transfer in deep space at home and abroad. For orbit maintenance control of satellites in Halo orbits. The TDE40F2A Hall thruster is a 40mN thruster and mainly comprises 1 anode module and 2 cathode modules. The Hall thruster is used as the final end of the electric propulsion subsystem gas circuit, xe propellant is respectively conveyed to the anode and the cathode of the Hall thruster by an upstream flow regulating module according to a proportion, and after ionization and acceleration are carried out on a discharge channel of the thruster, electric neutral plasma plumes are sprayed outwards to generate required thrust. The product fully uses the design schemes of a XX-9A satellite 40mN Hall thruster flight prototype, an XY-4 300mN Hall thruster flight prototype and the like, and adopts mature process technology and components.
In the embodiment, the most common propulsion system configuration scheme for home and abroad deep space exploration track transfer is used for reference: "490N+10N" unified two-component constant pressure extrusion scheme. The one 490N engine is used for pushing the satellite from the initial orbit to the L2 target orbit, and the 12 10N thrusters are divided into two branches which are mutually backed up, are mainly used for attitude control of the satellite, and can be used as 490N backup to a certain extent. The 490N engine was operated in a constant pressure mode and the 10N thruster was operated in a drop pressure mode. The propellant is green dinitrogen tetroxide (MON-1) and methyl hydrazine (MMH), and is constrained by a whole star structure, the storage box layout is based on the thought of multiple domestic models, and a '2+2' structure is adopted, namely the oxidant and the combustion agent are respectively stored in 4 symmetrically-paved storage boxes, and two storage boxes of the same propellant are diagonally and symmetrically arranged.
In this embodiment, the attitude and orbit control system algorithm includes three modules, namely, working mode management, attitude determination and attitude control, which can be divided into working mode management, single machine data acquisition and preprocessing, environment model calculation, attitude determination mode selection and attitude determination algorithm, guidance law generation algorithm, control mode selection and attitude control algorithm, and executing mechanism instruction distribution.
The attitude determination mode and algorithm utilize an attitude sensor configured by a system, an applicable attitude determination algorithm is selected, and an attitude quaternion q of a satellite body relative to an inertial coordinate system is determined bi And angular velocity information omega bi . According to the satellite attitude determined by the attitude determination module and the reference system attitude calculation module, combining the current working mode and the instruction guideAnd selecting a proper control algorithm to control the 10N thruster or the cold air micro thruster to control the satellite attitude so that the actual attitude and the actual angular velocity of the satellite are consistent with the expected attitude and the expected angular velocity. The attitude sensor comprises the following components: sun sensor, star sensor and fiber optic gyroscope. The attitude determination algorithm of the satellite analyzes from two angles of attitude quaternion and attitude angular speed respectively: and the gesture quaternion gesture determining algorithm and the gesture angular speed gesture determining algorithm are two main types. From two aspects of practicality and algorithm complexity, the adopted gesture quaternion gesture determining algorithm is determined as follows: (1) a sun sensor and gyroscope attitude determination algorithm; (2) a gyro integral attitude determination algorithm; (3) a star sensor attitude determination algorithm; and (4) a star sensor double-vector attitude determination algorithm.
The invention has the advantages that the invention provides the attitude orbit control method and the attitude orbit control system for the L2 aircraft adapting to the Lagrangian points, the current attitude information of the satellite is calculated according to the measurement data acquired by the attitude sensitive device in the preset control stage, the attitude information determined by the attitude determination module is obtained, the calculation is carried out according to the attitude information and the control law corresponding to the preset control stage, the control command is generated, and the execution mechanism is driven to execute the control command. Therefore, different attitude determination and control modes are provided in different task stages, and the attitude control requirements of each stage of the satellite are met; meanwhile, a multi-working mode and a software and hardware safety mode are designed, and the reconfigurability and the reliability of the software system are improved.
The above embodiments are only for illustrating the technical solution of the present invention, and are not limiting; although the invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical scheme described in the foregoing embodiments can be modified or some technical features thereof can be replaced by equivalents; these modifications or substitutions do not depart from the essence of the corresponding technical solutions from the protection scope of the technical solutions of the embodiments of the present invention.
Claims (8)
1. The attitude and orbit control method for the L2 aircraft adapting to the Lagrangian point is characterized by comprising the following steps of:
s1, after a satellite and an arrow are separated, angular speed and attitude deviation caused by the separation of the satellite and the arrow are rapidly eliminated, and sun orientation of a solar cell array is realized;
s2, controlling to realize the transfer of the satellite from the GTO to the L2 orbit;
s3, after the scientific cabin and the propulsion cabin of the satellite are separated, the separation angular speed and the attitude deviation are eliminated, and the sun alignment of the satellite solar cell array is realized again;
s4, in the satellite task stage, carrying out attitude measurement and attitude control according to load requirements;
s5, during the orbit running of the satellite, orbit maintaining control is carried out according to the satellite orbit drifting condition.
2. The attitude and orbit control method for an adaptive earth-Lagrangian point L2 aircraft according to claim 1, further comprising a safety mode 1 and a safety mode 2, wherein the safety mode 1 is used for adjusting a control scheme of the satellite when the satellite fails and the normal mode cannot guarantee the stability of the attitude of the satellite, damping the angular velocity of the satellite by chemical pushing and controlling the satellite to realize the sun orientation; the safety mode 2 is used for adjusting a control scheme of the scientific cabin when the scientific cabin of the satellite fails and the normal mode cannot guarantee stable posture, damping the angular speed of the scientific cabin by using the cold air micro-thruster, and controlling the scientific cabin to realize sun orientation.
3. The attitude and orbit control method for an adaptive earth lagrangian point L2 aircraft according to claim 1, wherein step S1 specifically comprises: after the satellites and the arrows are separated, the satellites automatically perform a push rate damping mode, the push rate damping mode is used for damping satellite angular velocity deviation caused by the separation of the satellites and the arrows, damping the triaxial angular velocity to a set range, and utilizing the collected solar vector information, and exerting control to realize solar orientation of the solar sailboard.
4. The attitude and orbit control method for an adaptive earth lagrangian point L2 aircraft according to claim 1, wherein step S2 specifically comprises: in the satellite orbit transfer process, an orbit-changing attitude maneuver mode is entered before orbit maneuver, and the orbit-changing attitude maneuver mode is used for adjusting the satellite attitude to prepare for orbit maneuver; after the satellite finishes attitude maneuver, meeting the attitude pointing requirement, igniting an orbit control engine, and performing orbit maneuver; after the satellite enters a working orbit after orbit transfer is finished, the satellite enters a gesture capturing mode, and triaxial stable pointing control is realized.
5. The attitude and orbit control method for an adaptive earth lagrangian point L2 aircraft according to claim 1, wherein step S3 specifically comprises: after the satellite enters a preset orbit, a ground instruction enters a propulsion cabin separation mode, and the separation of a propulsion cabin and a scientific cabin of the satellite is implemented; when the separation of the scientific cabin and the propulsion cabin of the satellite is finished, implementing the speed damping of the scientific cabin, and damping the angular speed deviation caused by the separation; and (5) utilizing the collected solar vector information to exert control to realize the sun orientation of the solar sailboard.
6. A attitude and orbit control system of an L2 aircraft adapted to a solar-earth lagrangian point, wherein the attitude and orbit control system is used for realizing the method of any one of claims 1 to 5, and comprises an attitude sensor, a controller and an executing mechanism, wherein the attitude sensor is used for measuring scientific load and obtaining an attitude determination scheme; the controller is integrated in the satellite computer and is used for collecting data of the attitude sensor and analyzing and processing the data to obtain satellite attitude; the executing mechanism is used for receiving a control instruction sent by the controller to realize attitude and orbit control of the satellite.
7. The attitude and orbit control system for an adaptive earth-Lagrangian point L2 aircraft according to claim 3, wherein the attitude sensor comprises a star sensor, an optical fiber inertial sensor, an analog sun sensor and a load camera, wherein the star sensor is an attitude measurement component, the optical fiber inertial sensor is used for directly measuring attitude angular velocity and acceleration information under a satellite inertial system, the analog sun sensor is used for measuring a sun vector of a position where a satellite is located, and the load camera is used for capturing images.
8. The attitude and orbit control system for an adaptive earth-Lagrangian point L2 aircraft of claim 4, wherein the actuator comprises a cold gas micro-thruster, a chemical thruster, an electric thruster, and a reaction flywheel.
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