CN111605733B - Spacecraft autonomous cooperative coarse-fine layering main-to-main integrated three-layer control parameter determination method - Google Patents
Spacecraft autonomous cooperative coarse-fine layering main-to-main integrated three-layer control parameter determination method Download PDFInfo
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Abstract
A spacecraft autonomous cooperative coarse-fine layered main-quilt integrated three-layer control parameter determination method is suitable for the fields with the requirements of ultrahigh-precision determination of loads, such as astronomical observation, high-resolution earth observation and the like. A control parameter design method is provided for spacecraft three-phase control with ultrahigh precision, ultrahigh stability and hypersensitive control, parameters of controllers of a spacecraft three-phase control system are designed respectively based on an index decomposition method, and design efficiency and control performance are improved. The main design idea is as follows: 1) firstly, establishing a control model for controlling three levels of a star body, a load and a quick reflector according to a three-level control system architecture; 2) deducing control loop transfer functions of all levels of three-level control according to the three-level control system model; 3) according to the noise characteristics of the selected sensor and the actuator, parameters of each level of controller are designed by a frequency domain analysis method, so that the power spectral density of each level of control loop meets the design index, and the three-level control performance of the spacecraft is realized.
Description
Technical Field
The invention relates to a three-level attitude control system for realizing a star-load-fast reflector of a spacecraft and a parameter design method, belonging to the field of attitude control of the spacecraft.
Background
At present, components containing high-speed rotors, such as flywheels, control moment gyros and the like, are generally adopted by spacecrafts as actuating mechanisms of attitude control systems. These high speed rotating components inevitably produce high frequency jitter and micro-vibrations that directly affect the performance of the load. The requirements of astronomical observation, extremely high resolution earth observation and other space missions with optical load high-performance control requirements cannot be met. The multi-stage composite control system of the spacecraft is generated according to the requirements of the optical load high-precision attitude control. The multi-stage composite control system of the spacecraft is a spacecraft platform with three-level control performances of ultra-high precision pointing, ultra-high stability control, hypersensitive agility control and the like.
In a traditional spacecraft control system, a star body is rigidly connected with a load, and the shake and the micro-vibration in the star body are directly transmitted to the load, so that the high-quality imaging performance of the optical load is influenced. The existing attitude control method is limited by the measurement accuracy and measurement bandwidth limitation of an attitude sensor and the like, and the isolation and the inhibition of high-frequency jitter cannot be realized. The prior control system has the following defects:
1. cannot realize isolation and suppression of high-frequency micro vibration of star
In the attitude control system of the existing spacecraft, a load is rigidly connected with a star body. The flexible vibration and high-frequency micro-vibration existing in the spacecraft star body are directly transmitted to the load, so that the optical load cannot further improve the imaging quality. The traditional spacecraft attitude system is limited by the bandwidth of a controller and the precision of an actuating mechanism, and cannot realize active control on flexible vibration and high-frequency micro-vibration, so that the control precision and stability of the star body are further improved and limited.
2. It is difficult to realize the ultra-high precision pointing and ultra-high stability control of the load optical axis
The simulation model of the existing spacecraft control system only has a star-body primary model and does not contain an active pointing hyperstatic platform and a quick reflector mathematical model. The pointing accuracy and stability of the load cannot be further improved due to the factors of sensor measurement bandwidth, actuator response bandwidth and the like. In a satellite control system model comprising an active pointing hyperstatic platform and a fast reflector, a spacecraft multistage composite pointing control scheme needs to be designed, and meanwhile, an index decomposition and parameter design method based on frequency domain analysis is provided for solving the problems that multistage composite control design parts are more and transmission relations of channels are complex, so that rapid optimal design of multistage composite control is realized, and three-phase control performance of the spacecraft is realized.
Disclosure of Invention
The technical problem solved by the invention is as follows: the method overcomes the defects of the prior art, and provides a spacecraft autonomous cooperative coarse-fine layering main-quilt integrated triple control parameter determination method.
The technical solution of the invention is as follows:
the method for determining the autonomous cooperative coarse-fine layering main-to-main integrated three-layer control parameters of the spacecraft comprises the following steps:
(1) establishing a three-level control model for a spacecraft multilevel system, wherein the three-level control model comprises a control model of a satellite control loop, a control model of a load control loop and a control model of a fast reflector control loop;
(2) according to the three-level control model and the sensor type adopted by the spacecraft multilevel system, the transfer function of the CMG, the star sensor and the star gyroscope transferred to the star attitude angle is calculated;
(3) calculating the transfer functions of the angular speed of the star body, the load controller, the load gyroscope, the star sensor and the load relative attitude measurement which are respectively transferred to the load attitude angle;
(4) and the star inertia and the load inertia of the spacecraft are brought into the transfer functions to obtain a multi-stage system actual model of the spacecraft, the actual characteristic parameters of each sensor and each actuating mechanism are used as the input of each transfer function, and the square sum of the power spectral density of the transfer function from each factor to the load attitude angle is minimized by adjusting the control parameters and the filter coefficients, so that the control parameters with the optimal load attitude precision and stability under the actual spacecraft characteristics can be obtained, and the three-phase super control of the spacecraft is realized.
Further, the control model of the star control loop is as follows:
wherein, IbIs the rotational inertia of a star, HvscmgGyro angular momentum, omega, being the control moment on the starbIs the angular velocity of the star, TpbFor actively pointing to the reaction moment, T, of the hyperstatic platform on the satellite platformbdExternal disturbance torque, T, to the starbThe control moment of the star body is expressed as follows:
in the formula IsatIs the total inertia, k, of the multi-stage system of the spacecraftsatp、ksatd、ksatiIs a controller parameter; delta thetaberr、ΔωberrThe attitude control error and the angular velocity control error of the star body, omegabrThe angular velocity is desired for the stars;
representing the derivative of the vector r, r×An antisymmetric matrix representing a vector r, where r can be any vector.
Further, the control model of the fast reflector control loop is:
wherein, Jfx、JfyRotation of the fast mirror and the x, y axes of the rotating part of the support structureInertia, thetafx、θfyIs the attitude angle of the x and y axes of the fast mirror, thetapx、θpyRoll and pitch angles, k, of the load, respectivelyθTorsional rigidity of the support structure for the fast mirror, cθDamping coefficient, m, for linear motors in fast mirror support structures and actuatorsθFor the mass of the linear motor mover in the actuator, lθIs the distance between the acting point of the linear motor and the rotating shaft of the fast reflector, Tdfx、TdfyIs the external disturbance torque, T, applied to the fast reflecting mirrorfx、TfyThe control moments of the x and y axes of the fast reflector are respectively calculated according to the following formula:
in the formula, kfpx、kfpy、kfdx、kfdy、kfix、kfiyAs a controller parameter, θfxr、θfyrIs the desired attitude angle of the fast mirror x, y axes.
Further, the control model of the load control loop is as follows:
wherein, IpIs the inertia of the load, ωpFor angular velocity of the load, TpFor active control of moment, T, for loaddpDisturbance torque, T, imposed on the loadfpFor the load to be subjected to the active and passive reaction torque of the quick reflecting mirror actuating mechanism, the calculation formulas are respectively as follows:
Tfp＝[Tfpx,Tfpy,0]T
in the formula IpcIs the expression of the inertia of the load relative to the whole star centroid in a load coordinate system, kpp、kpd、kpiIs a load controller parameter; delta thetaperr、ΔωperrRespectively, a load attitude control error and an angular velocity control error, omegaprAngular velocity is desired for the load.
Further, the transfer functions of the CMG, the star sensor and the star gyroscope transferred to the star attitude angle are respectively as follows:
wherein, ω iscc、ωcp、ωcgRespectively is a CMG noise inertia characteristic coefficient, a star sensor filter coefficient and a star gyro filter coefficient; WNcmgθ、WNSPθ、WNSGθThe transfer functions are respectively the transfer functions of the CMG, the star sensor and the star gyroscope to the star attitude angle, and s is the independent variable of the complex frequency domain in the transfer functions.
Further, the transfer functions of the star angular velocity, the load controller, the load gyroscope, the star sensor and the load relative attitude measurement transferred to the load attitude angle are respectively as follows:
wherein, ω iscT、ωcR、ωcLRespectively a load controller, a load gyro and a filter coefficient measured by the relative attitude of the load, k and c respectively are an integral rigidity coefficient and a damping coefficient of a passive vibration isolation link of the load, krpThe control coefficient is compensated for load.
Further, the spacecraft multilevel system comprises a star platform, an active pointing hyperstatic platform, a load, a sensor system and a fast reflector;
the load is an optical system used for imaging the celestial body;
the quick reflector is arranged in the load and used for adjusting the direction of the optical axis of the load;
the sensor system is used for measuring data;
the star platform is used for supporting the active pointing hyperstatic platform and the load;
the active pointing hyperstatic platform is arranged between the load and the star platform, the upper plane of the active pointing hyperstatic platform is connected with the load, and the lower plane of the active pointing hyperstatic platform is connected with the star platform; the active pointing hyperstatic platform consists of six actuators, and each actuator comprises a displacement sensor, a spring-damping structure and a linear motor;
the displacement sensor is used for measuring the translational displacement of the linear motor; the spring-damping structure is used for isolating the high-frequency vibration of the star platform; the linear motor is used for providing main power and realizing the attitude control of the load.
Furthermore, the linear motor is used as an active link, the spring-damping structure is used as a passive link, the fast reflector realizes the passive link through the supporting structure of the fast reflector, and the spring-damping structure is installed in parallel with the linear motor; the spring-damper structure includes a spring and a damper in parallel.
Furthermore, the sensor system of the spacecraft multilevel system comprises a star gyroscope, a load star sensor, a load micrometer sensor, a guide star sensor, an active pointing hyperstatic platform displacement measurement sensor and a fast reflector displacement measurement sensor;
the star gyroscope measures the angular speed of a star; the load star sensor measures the inertia attitude of the load; the load micro-sensor measures the angular speed of the load; the guide star sensor measures the inertia attitude of the fast reflector; actively pointing to the hyperstatic platform displacement measurement sensor to measure the relative attitude of the load and the star; the fast reflector displacement measuring sensor measures the relative attitude of the fast reflector and the load.
Furthermore, the load micrometer sensor is realized by adopting a micrometer optical fiber gyroscope, and the precision is one order of magnitude higher than that of a star gyroscope.
Compared with the prior art, the invention has the advantages that:
1. the invention realizes the micro-vibration isolation and inhibition of the star platform and improves the load stability index.
The existing spacecraft attitude control system only has star-body primary attitude control. The load is rigidly connected with the star body, and the medium-high frequency micro-vibration of the star body platform is directly transmitted to the load, so that the stability index of the load is seriously reduced. The multi-stage composite control system of the spacecraft, which is designed by the invention, realizes the attenuation of 20dB of high-frequency disturbance (>10Hz) in the satellite platform through a passive link actively pointing to the hyperstatic platform. And the attenuation of 10dB of low-frequency disturbance (<10Hz) in the satellite platform is further realized through active control of the active directional hyperstatic platform. Through a fast reflector passive link, the low-frequency micro-vibration (<0.1Hz) of the star platform and the attenuation of low-frequency disturbance 20dB in load can be realized. The micro-vibration isolation and inhibition of the satellite platform are realized through two-stage passive links, and the load stability is improved.
2. The invention can realize the rapid optimal design of the parameters of the multi-stage composite control system. The multi-stage composite control system of the spacecraft has the characteristics of complex control structure, various disturbance sources and more related components, and the transfer characteristic of the multi-stage composite control system is far more complex than that of the traditional single-stage control system. According to the system index analysis and parameter design method based on index decomposition, the transfer functions of various disturbance factors to the control indexes are deduced according to the model of the multi-stage composite control system, the parameter comprehensive optimization rapid design of each stage of controllers and filters is realized through a frequency domain analysis method, the design efficiency and the control performance are greatly improved compared with the design parameters of a traditional trial and error method, and the three-super performance control of a spacecraft is realized.
Drawings
FIG. 1 is a diagram of a multi-stage compound control system of a spacecraft;
FIG. 2 is a schematic diagram of analysis of various noise influences of spacecraft star attitude accuracy;
FIG. 3 is a schematic diagram of analysis of various noise influences on the stability of a spacecraft star;
FIG. 4 is a schematic diagram of analysis of various noise influences on spacecraft load attitude accuracy;
fig. 5 is a schematic diagram of various noise influences on the load stability of the spacecraft.
Detailed Description
The spacecraft multilevel system provided by the invention comprises a satellite platform, an active pointing hyperstatic platform, a load, a sensor system and a fast reflector;
the load is an optical system used for imaging the celestial body;
the quick reflector is arranged in the load and used for adjusting the direction of the optical axis of the load;
the sensor system is used for measuring data;
the star platform is used for supporting the active pointing hyperstatic platform and the load;
the active pointing hyperstatic platform is arranged between the load and the star platform, the upper plane of the active pointing hyperstatic platform is connected with the load, and the lower plane of the active pointing hyperstatic platform is connected with the star platform; the active pointing hyperstatic platform consists of six actuators, and each actuator comprises a displacement sensor, a spring-damping structure and a linear motor;
the displacement sensor is used for measuring the translational displacement of the linear motor; the spring-damping structure is used for isolating the high-frequency vibration of the star platform; the linear motor is used for providing main power and realizing the attitude control of the load.
Preferably, the first and second liquid crystal materials are,
1. the linear motor is used as an active link;
2. the spring-damping structure is used as a passive link;
3. the fast reflector comprises a passive link, and particularly, the fast reflector realizes the passive link through a supporting structure of the fast reflector.
The spring-damping structure is installed in parallel with the linear motor; the spring-damper structure includes a spring and a damper in parallel.
The invention realizes the attenuation of 20dB of high-frequency disturbance (>10Hz) in the satellite platform by actively pointing to a passive link (a spring-damping structure) of the hyperstatic platform. The attenuation of 10dB of low-frequency disturbance (<10Hz) in the star platform is further realized through active control (linear motor) of the active directional hyperstatic platform. Through a fast reflector passive link (a supporting structure), the low-frequency micro-vibration (<0.1Hz) of the star platform and the attenuation of low-frequency disturbance 20dB in load can be realized. The micro-vibration isolation and inhibition of the satellite platform are realized through two-stage passive links, and the load stability is improved.
Preferably, the sensor system of the spacecraft multilevel system comprises a star gyroscope, a load star sensor, a load micrometering sensor, a guide star sensor, an active-pointing hyperstatic platform displacement measurement sensor and a fast reflector displacement measurement sensor;
the star gyroscope measures the angular speed of a star; the load star sensor measures the inertia attitude of the load; the load micro-sensor measures the angular speed of the load; the guide star sensor measures the inertia attitude of the fast reflector; actively pointing to the hyperstatic platform displacement measurement sensor to measure the relative attitude of the load and the star; the fast reflector displacement measuring sensor measures the relative attitude of the fast reflector and the load.
Preferably, the load micrometer sensor is realized by adopting a micrometer fiber-optic gyroscope, and the precision is one order of magnitude higher than that of a star gyroscope.
Aiming at the spacecraft multilevel system, the invention discloses a spacecraft autonomous cooperative coarse-fine layering main-passive integrated three-layer control parameter design method, as shown in figure 1, which comprises the following specific steps:
(1) establishing a three-level control model for a spacecraft multilevel system, wherein the three-level control model comprises a control model of a satellite control loop, a control model of a load control loop and a control model of a fast reflector control loop;
the control model of the star control loop is as follows:
wherein, IbIs the rotational inertia of a star, HvscmgGyro angular momentum, omega, being the control moment on the starbIs the angular velocity of the star, TpbFor actively pointing to the reaction moment, T, of the hyperstatic platform on the satellite platformbdExternal disturbance torque, T, to the starbThe control moment of the star body is expressed as follows:
in the formula IsatIs the total inertia, k, of the multi-stage system of the spacecraftsatp、ksatd、ksatiIs a controller parameter; delta thetaberr、ΔωberrThe attitude control error and the angular velocity control error of the star body, omegabrThe angular velocity is desired for the stars;
representing the derivative of the vector r, r×An antisymmetric matrix representing a vector r, where r can be any vector.
The control model of the fast reflector control loop is as follows:
wherein, Jfx、JfyThe moment of inertia, theta, of the fast mirror and the x, y axes of the rotating part of the support structurefx、θfyIs the attitude angle of the x and y axes of the fast mirror, thetapx、θpyRoll and pitch angles, k, of the load, respectivelyθTorsional rigidity of the support structure for the fast mirror, cθDamping coefficient, m, for linear motors in fast mirror support structures and actuatorsθFor the mass of the linear motor mover in the actuator, lθIs the distance between the acting point of the linear motor and the rotating shaft of the fast reflector, Tdfx、TdfyIs the external disturbance torque, T, applied to the fast reflecting mirrorfx、TfyThe control moments of the x and y axes of the fast reflector are respectively calculated according to the following formula:
in the formula, kfpx、kfpy、kfdx、kfdy、kfix、kfiyAs a controller parameter, θfxr、θfyrIs the desired attitude angle of the fast mirror x, y axes.
The control model of the load control loop is as follows:
wherein, IpIs the inertia of the load, ωpFor angular velocity of the load, TpFor active control of moment, T, for loaddpDisturbance torque, T, imposed on the loadfpFor the load to be subjected to the active and passive reaction torque of the quick reflecting mirror actuating mechanism, the calculation formulas are respectively as follows:
Tfp＝[Tfpx,Tfpy,0]T
in the formula IpcIs the expression of the inertia of the load relative to the whole star centroid in a load coordinate system, kpp、kpd、kpiIs a load controller parameter; delta thetaperr、ΔωperrRespectively, a load attitude control error and an angular velocity control error, omegaprAngular velocity is desired for the load.
(2) According to the three-level control model and the sensor type adopted by the spacecraft multilevel system, the transfer function of the CMG, the star sensor and the star gyroscope transferred to the star attitude angle is calculated;
specifically, the transfer functions of the CMG, the star sensor and the star gyroscope, which are transferred to the star attitude angle respectively, are as follows:
wherein, ω iscc、ωcp、ωcgRespectively is a CMG noise inertia characteristic coefficient, a star sensor filter coefficient and a star gyro filter coefficient; WNcmgθ、WNSPθ、WNSGθThe transfer functions are respectively the transfer functions of the CMG, the star sensor and the star gyroscope to the star attitude angle, and s is the independent variable of the complex frequency domain in the transfer functions.
(3) Calculating the transfer functions of the angular speed of the star body, the load controller, the load gyroscope, the star sensor and the load relative attitude measurement which are respectively transferred to the load attitude angle;
specifically, the transfer functions of the star angular velocity, the load controller, the load gyroscope, the star sensor and the load relative attitude measurement transferred to the load attitude angle are respectively as follows:
wherein, ω iscT、ωcR、ωcLRespectively a load controller, a load gyro and a filter coefficient measured by the relative attitude of the load, k and c respectively are an integral rigidity coefficient and a damping coefficient of a passive vibration isolation link of the load, krpThe control coefficient is compensated for load.
(4) And the star inertia and the load inertia of the spacecraft are brought into the transfer functions to obtain a multi-stage system actual model of the spacecraft, the actual characteristic parameters of each sensor and each actuating mechanism are used as the input of each transfer function, and the square sum of the power spectral density of the transfer function from each factor to the load attitude angle is minimized by adjusting the control parameters and the filter coefficients, so that the control parameters with the optimal load attitude precision and stability under the actual spacecraft characteristics can be obtained, and the three-phase super control of the spacecraft is realized.
Examples of the present invention are given.
Bringing the star inertia, load inertia and fast mirror inertia of a spacecraft into the aboveThe transfer functions yield a system actual model, where Ib＝diag(10000,10000,8000)kgm2，Ip＝diag(140,140,130)kgm2，If＝diag(0.001,0.001)kgm2The real characteristic parameters of each sensor and each actuator are used as the input of each transfer function, wherein the noise of the star sensor is 2' (3 sigma), and the noise of the star gyro is
The noise of the load gyro is
Random noise of CMG is
The inertia characteristic coefficient of CMG noise is 0, and the noise of the load controller is 2 multiplied by 10-2N (3 sigma), the noise of the load relative attitude measurement is 1 x 10-6m (3. sigma.). By designing appropriate control parameters and filter coefficients, ksatp＝103×diag(4.07,4.07,2.42)，ksati＝diag(0.02,0.02,0.02)，ksatd＝103×diag(9.08,9.08,5.41)，k＝50000N/m，c＝200Ns/m，kpp＝104×diag(4.13,4.13,3.628)，kpi＝105×diag(1.34,1.34,1.823)，kpd＝103×diag(5923,5923,3222)，kfpx＝kfpy＝51.26，kfdx＝kfdy＝0.1362，kfix＝kfiy＝0，ωcp＝3×10-3、ωcg＝2×10-2，ωcR＝2×10-2，ωcT＝0，ωcLAnd (2) the sum of squares of the power spectral densities of the transfer functions from various factors to the load attitude is minimized, so that a control scheme with optimal load attitude precision and stability under the actual spacecraft characteristics can be obtained, and the three-phase super control of the spacecraft is realized.
And (5) checking a frequency domain analysis and parameter design method. The designed control parameters and filter coefficients are brought into each transfer function, the output power spectral density of the star body and the load attitude angle is inspected, the calculation shows that the star body attitude control precision reaches 1.6 percent, the attitude control stability reaches 22 '/s, the load attitude control precision reaches 0.084 percent, the attitude control stability reaches 3.8'/s, the proportion of each noise factor is shown in figures 2-5, and the results show that the attitude precision and the stability of the load are obviously superior to those of the star body attitude, which indicates that the control performance of the multilevel composite control provided by the invention is superior to that of single-level control.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Claims (10)
1. The method for determining the autonomous cooperative coarse-fine hierarchical main-to-main integrated triple control parameters of the spacecraft is characterized by comprising the following steps of:
(1) establishing a three-level control model for a spacecraft multilevel system, wherein the three-level control model comprises a control model of a satellite control loop, a control model of a load control loop and a control model of a fast reflector control loop;
(2) according to the three-level control model and the sensor type adopted by the spacecraft multilevel system, the transfer function of the CMG, the star sensor and the star gyroscope transferred to the star attitude angle is calculated;
(3) calculating the transfer functions of the angular speed of the star body, the load controller, the load gyroscope, the star sensor and the load relative attitude measurement which are respectively transferred to the load attitude angle;
(4) and substituting the star inertia and the load inertia of the spacecraft into the transfer functions to obtain an actual model of the multi-stage system of the spacecraft, taking the actual characteristic parameters of each sensor and each actuating mechanism as the input of each transfer function, and adjusting the control parameters and the filter coefficients to minimize the square sum of the power spectral densities of the transfer functions from each factor to the load attitude angle, so that the control parameters with the optimal load attitude precision and stability under the actual spacecraft characteristics can be obtained, and the control of the spacecraft is realized.
2. The spacecraft autonomous cooperative coarse-fine layering main-integrated three-layer control parameter determination method according to claim 1, characterized in that: the control model of the star control loop is as follows:
wherein, IbIs the rotational inertia of a star, HvscmgGyro angular momentum, omega, being the control moment on the starbIs the angular velocity of the star, TpbFor actively pointing to the reaction moment, T, of the hyperstatic platform on the satellite platformbdExternal disturbance torque, T, to the starbThe control moment of the star body is expressed as follows:
in the formula IsatIs the total inertia, k, of the multi-stage system of the spacecraftsatp、ksatd、ksatiIs a controller parameter; delta thetaberr、ΔωberrThe attitude control error and the angular velocity control error of the star body, omegabrThe angular velocity is desired for the stars;
representing the derivative of the vector r, r×An antisymmetric matrix representing a vector r, where r can be any vector.
3. The spacecraft autonomous cooperative coarse-fine layering main-integrated three-layer control parameter determination method according to claim 2, characterized in that: the control model of the fast reflector control loop is as follows:
wherein, Jfx、JfyThe moment of inertia, theta, of the fast mirror and the x, y axes of the rotating part of the support structurefx、θfyIs the attitude angle of the x and y axes of the fast mirror, thetapx、θpyRoll and pitch angles, k, of the load, respectivelyθTorsional rigidity of the support structure for the fast mirror, cθDamping coefficient, m, for linear motors in fast mirror support structures and actuatorsθFor the mass of the linear motor mover in the actuator, lθIs the distance between the acting point of the linear motor and the rotating shaft of the fast reflector, Tdfx、TdfyIs the external disturbance torque, T, applied to the fast reflecting mirrorfx、TfyThe control moments of the x and y axes of the fast reflector are respectively calculated according to the following formula:
in the formula, kfpx、kfpy、kfdx、kfdy、kfix、kfiyAs a controller parameter, θfxr、θfyrIs the desired attitude angle of the fast mirror x, y axes.
4. The spacecraft autonomous cooperative coarse-fine layering main-integrated three-layer control parameter determination method according to claim 3, characterized in that: the control model of the load control loop is as follows:
wherein, IpIs the inertia of the load, ωpFor angular velocity of the load, TpFor active control of moment, T, for loaddpDisturbance torque, T, imposed on the loadfpFor the load to be subjected to the active and passive reaction torque of the quick reflecting mirror actuating mechanism, the calculation formulas are respectively as follows:
Tfp＝[Tfpx,Tfpy,0]T
in the formula IpcIs the expression of the inertia of the load relative to the whole star centroid in a load coordinate system, kpp、kpd、kpiIs a load controller parameter; delta thetaperr、ΔωperrRespectively, a load attitude control error and an angular velocity control error, omegaprAngular velocity is desired for the load.
5. The spacecraft autonomous cooperative coarse-fine layering main-integrated three-layer control parameter determination method according to claim 4, characterized in that: the transfer functions of the CMG, the star sensor and the star gyroscope transferred to the star attitude angle are respectively as follows:
wherein, ω iscc、ωcp、ωcgRespectively is a CMG noise inertia characteristic coefficient, a star sensor filter coefficient and a star gyro filter coefficient; WNcmgθ、WNSPθ、WNSGθThe transfer functions are respectively the transfer functions of the CMG, the star sensor and the star gyroscope to the star attitude angle, and s is the independent variable of the complex frequency domain in the transfer functions.
6. The spacecraft autonomous cooperative coarse-fine layering main-integrated three-layer control parameter determination method according to claim 5, characterized in that: the transfer functions of the star angular velocity, the load controller, the load gyroscope, the star sensor and the load relative attitude measurement transferred to the load attitude angle are respectively as follows:
wherein, ω iscT、ωcR、ωcLRespectively a load controller, a load gyro and a filter coefficient measured by the relative attitude of the load, k and c respectively are an integral rigidity coefficient and a damping coefficient of a passive vibration isolation link of the load, krpThe control coefficient is compensated for load.
7. The spacecraft autonomous cooperative coarse-fine layering main-integrated three-layer control parameter determination method according to claim 3, characterized in that: the spacecraft multilevel system comprises a star platform, an active pointing hyperstatic platform, a load, a sensor system and a fast reflector;
the load is an optical system used for imaging the celestial body;
the quick reflector is arranged in the load and used for adjusting the direction of the optical axis of the load;
the sensor system is used for measuring data;
the star platform is used for supporting the active pointing hyperstatic platform and the load;
the active pointing hyperstatic platform is arranged between the load and the star platform, the upper plane of the active pointing hyperstatic platform is connected with the load, and the lower plane of the active pointing hyperstatic platform is connected with the star platform; the active pointing hyperstatic platform consists of six actuators, and each actuator comprises a displacement sensor, a spring-damping structure and a linear motor;
the displacement sensor is used for measuring the translational displacement of the linear motor; the spring-damping structure is used for isolating the high-frequency vibration of the star platform; the linear motor is used for providing main power and realizing the attitude control of the load.
8. The spacecraft autonomous cooperative coarse-fine layering main-integrated three-layer control parameter determination method according to claim 7, characterized in that: the linear motor is used as an active link, the spring-damping structure is used as a passive link, the fast reflector realizes the passive link through the supporting structure of the fast reflector, and the spring-damping structure is installed in parallel with the linear motor; the spring-damper structure includes a spring and a damper in parallel.
9. The spacecraft autonomous cooperative coarse-fine layering main-integrated three-layer control parameter determination method according to claim 7, characterized in that: the sensor system of the spacecraft multilevel system comprises a star gyroscope, a load star sensor, a load micrometering sensor, a guide star sensor, an active pointing hyperstatic platform displacement measuring sensor and a fast reflector displacement measuring sensor;
the star gyroscope measures the angular speed of a star; the load star sensor measures the inertia attitude of the load; the load micro-sensor measures the angular speed of the load; the guide star sensor measures the inertia attitude of the fast reflector; actively pointing to the hyperstatic platform displacement measurement sensor to measure the relative attitude of the load and the star; the fast reflector displacement measuring sensor measures the relative attitude of the fast reflector and the load.
10. The spacecraft autonomous cooperative coarse and fine layering main integrated three-layer control parameter determination method according to claim 9, characterized in that: the load micrometer sensor is realized by adopting a micrometer optical fiber gyroscope, and the precision is one order of magnitude higher than that of a star gyroscope.
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