CN108897239B - Spacecraft two-stage attitude control simulation system - Google Patents
Spacecraft two-stage attitude control simulation system Download PDFInfo
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Abstract
A two-stage attitude control simulation system of a spacecraft is used for verifying three-level control technologies of 'ultra-high precision pointing', 'ultra-high stability control', 'hypersensitive agility control' and the like of the spacecraft. The verification system includes: the system comprises a star body, a load simulator, an active pointing platform, a star body primary control loop and a load simulator secondary control loop; the star primary control circuit and the load simulator secondary control circuit both comprise: the device comprises a control unit, an actuating mechanism and a measuring unit; the star primary control loop is connected with the load simulator secondary control loop through a platform; the active pointing platform provides active control force for the secondary control loop of the load simulator; the load simulator transmits the reaction force of the active control force to the primary control loop of the star body through the active pointing platform. The spacecraft two-stage attitude control simulation system constructed by the invention can verify the three-super platform spacecraft multi-stage composite control technology and control performance indexes.
Description
Technical Field
The invention belongs to the field of spacecraft attitude control, and relates to a spacecraft two-stage attitude control simulation system.
Background
The existing aerospace attitude control simulation system is mainly used for controlling the primary attitude of a star body, the secondary control of a load cannot be realized, the star body is fixedly connected with the load, the attitude of the load is adjusted along with the adjustment of the attitude of the star body, a star body executing mechanism is a control moment gyro, the accuracy of attitude control moment output by the control moment gyro is low, disturbance moment is large, and the attitude error between the attitude of the star body and the target attitude is large.
The ground full-physical simulation system of the single-stage spacecraft in the prior art has the following defects:
1. ultrahigh precision pointing and ultrahigh stability control of load cannot be realized
At present, components containing high-speed rotors, such as a flywheel, a control moment gyro and the like, are generally adopted in a ground full-physical test system of a spacecraft as an actuating mechanism of an attitude control system. The high-speed rotating part of the actuating mechanism inevitably generates high-frequency jitter and micro vibration, which directly influences the working performance of the load and can not realize the ultrahigh precision pointing and ultrahigh stability control performance of the load.
2. The load gravity unloading and the load secondary control can not be realized
The load of the three-super platform is connected with the star body by adopting an active pointing super-static platform. Due to the existence of ground gravity interference factors, zero-rigidity gravity unloading of the load is required, so that a good control environment is provided for secondary control of the load. In the traditional spacecraft ground full-physical simulation system, load gravity unloading cannot be realized, and load secondary control cannot be realized at the same time, so that the load control performance index is difficult to improve.
Disclosure of Invention
The technical problem solved by the invention is as follows: the defects of the prior art are overcome, the spacecraft two-stage attitude control simulation system is provided, and the problem that the spacecraft load cannot be controlled with high precision and high stability in the prior art is solved.
The technical solution of the invention is as follows:
a spacecraft two-stage attitude control simulation system comprising: a star body primary control loop, a load simulator secondary control loop, an active pointing platform, a star body and a load simulator; the active pointing platform is fixed on the star body, and the load simulator is fixed on the active pointing platform; the active pointing platform comprises 3 actuator groups, each actuator group comprising 2 actuators.
Star first-level control loop: judging the attitude error between the attitude of the star body and the target attitude of the load simulator, and when the attitude error is greater than or equal to a critical value, calculating the control moment of the star body according to the attitude error to reduce the attitude error between the star body and the target attitude; when the attitude error is smaller than a critical value, controlling the star body to maintain the attitude unchanged; the critical value is the maximum control angle of the actuator.
A second-level control loop of the load simulator: judging the attitude error between the load simulator and the target attitude of the load simulator, and controlling the relative attitude between the load simulator and the star body to be kept unchanged through the active pointing platform when the attitude error is greater than or equal to a critical value; and when the attitude error is smaller than a critical value, calculating the control moment of the load simulator according to the attitude error, and controlling the active pointing platform to reduce the attitude error between the load simulator and the target attitude of the load simulator according to the control moment of the load simulator.
The star body primary control loop comprises a star body control unit, a star body actuating mechanism and a star body measuring unit which are all fixed on a star body;
a star measuring unit: measuring the angular velocity and the posture of the star body and feeding back to the star body control unit;
a star body control unit: determining an attitude error between the star attitude and the target attitude of the load simulator according to the target attitude of the load simulator and the star angular velocity and the star attitude sent by the star measuring unit, and calculating the control moment of the star according to the attitude error;
the star body actuating mechanism: the star control device comprises a plurality of control moment gyroscopes, and the star postures are adjusted according to the star control moment calculated by the star control unit.
The star measurement unit includes: 3M optical fiber gyroscopes, wherein M is a positive integer larger than 3, and the measuring directions of at least 3 optical fiber gyroscopes are pairwise orthogonal.
The number of the control moment gyroscopes is 4, the low-speed frame rotating shaft of each control moment gyroscope is in a standard pyramid configuration, and the low-speed frame rotating shaft of each control moment gyroscope is perpendicular to the center lines of two edges of the pyramid configuration.
The load simulator secondary control loop comprises: the device comprises a load control unit, a load executing mechanism, a load measuring unit and a load gravity unloading device;
a load measuring unit: measuring the angular speed and the attitude of the load simulator and feeding back to the load control unit;
a load control unit: determining an attitude error between the attitude of the load simulator and a target attitude of the load simulator according to the angular velocity of the load simulator and the attitude of the load simulator, and calculating a control moment of the load simulator; controlling the active pointing platform to reduce the attitude error between the load simulator and the target attitude according to the control moment of the load simulator;
load gravity unloading device: the load simulator is arranged on the star body, and the load simulator is hoisted at the top of the star body and unloads the gravity of the load simulator;
a load executing mechanism: adjusting the attitude of the load simulator through active control force provided by an actuator of the active pointing platform;
a load control unit: the load simulator is arranged on the load simulator and used for controlling the posture of the load simulator;
a load measuring unit: and the device is fixed on the load simulator and used for measuring the angular speed and the attitude of the load simulator.
The actuator includes: a flexible hinge and a motor stator; mounting point p for fixing one end of each of 6 actuator flexible hinges on load simulatoriSix mounting points p of load simulator, i is 1,2, …,6iCoplanar; mounting point b for fixing one end of motor stator of 6 actuators on stariMounting points b of six starsiCoplanar.
The load gravity unloading device of the load simulator secondary control loop comprises: a spring, a support structure; the supporting structure comprises a flat plate, the end face of the flat plate is provided with a plurality of supporting rods, the supporting rods form a conical structure, and the other ends of the supporting rods are fixed on the star body; one end of the spring is fixed by the flat plate structure, and the other end of the spring is used for hoisting the load simulator.
The measuring unit of the load simulator secondary control loop comprises: two autocollimators, a cubic mirror and an inertial sensor; the cubic mirror is installed on a shell of the load simulator, optical axes of the two autocollimators point to two adjacent end faces of the cubic mirror, and the inertia sensor is installed on the load simulator and used for measuring the angular speed of the satellite load simulator.
Compared with the prior art, the invention has the beneficial effects that:
1) the star body primary attitude control loop and the load secondary attitude control loop are connected through the active pointing platform actuator, the actuator of the active pointing platform actively controls the attitude of the load simulator, the attitude error between the attitude of the load simulator and the target attitude is reduced, and the control precision of the simulation system is improved.
2) The load simulator of the simulation system provided by the invention is provided with an independent gravity unloading device, can realize zero-rigidity suspension of the load, solves the gravity unloading problem of a star body and load two-stage system, meets the microgravity environmental condition during on-orbit, and improves the authenticity of the simulation system for simulating the on-orbit running environment.
3) The executing mechanism comprises a control moment gyroscope and an actuator, wherein the control moment precision of the actuator is superior to that of the control moment gyroscope output, and when the attitude error between the attitude of the star body and the target attitude is greater than or equal to a critical value, the attitude of the star body and the attitude of the load simulator are adjusted by using the control moment gyroscope; and when the attitude error between the star body attitude and the target attitude is smaller than a critical value, independently controlling the load simulator by using the actuator with higher precision, and reducing the attitude error between the load simulator and the target attitude.
Drawings
FIG. 1 is a block diagram of the system of the present invention;
FIG. 2 is a schematic diagram of a platform configuration according to the present invention;
FIG. 3 is a mounting configuration diagram of a load micrometer sensor;
FIG. 4 is a test chart of the load and star stability orientation;
fig. 5 is a test chart of agility maneuver orientation of load and star.
Detailed Description
With the continuous improvement of astronomical observation requirements, a control system is required to realize the three-super control performance of the optical load, namely the control of the optical load, namely the ultrahigh precision pointing, the ultrahigh stability control and the hypersensitive control. The three-super platform is just based on the requirement of aiming at the optical load and controlling the attitude with high precision. The emerging three-super-platform spacecraft consists of a star body and a load two-stage system, wherein the star body and the load are connected through a flexible active pointing super-static platform. The relevant control performance indexes of the three-super platform must be subjected to strict examination of ground full physical tests before model application. Aiming at large loads, how to construct a three-super-platform full-physical simulation test system is the problem which is mainly solved by engineers.
Different from the traditional single-stage spacecraft, the three-stage super platform spacecraft is provided with a star body first-stage control loop and a load simulator second-stage control loop. The design method of the three-super platform related control technology ground full-physical test system is completely different from the design method of the single-stage spacecraft ground full-physical test system.
As shown in fig. 1, the two-stage attitude control simulation system of the present invention includes: a star body primary control loop, a load simulator secondary control loop, an active pointing platform, a star body and a load simulator; the active pointing platform is fixed on the star body, and the load simulator is fixed on the active pointing platform; the active pointing platform comprises 3 actuator groups, each actuator group comprising 2 actuators.
Star first-level control loop: judging the attitude error between the attitude of the star body and the target attitude of the load simulator, and when the attitude error is greater than or equal to a critical value, calculating the control moment of the star body according to the attitude error to reduce the attitude error between the star body and the target attitude; when the attitude error is smaller than a critical value, controlling the star body to maintain the attitude unchanged;
a second-level control loop of the load simulator: judging the attitude error between the load simulator and the target attitude of the load simulator, and controlling the relative attitude between the load simulator and the star body to be kept unchanged through the active pointing platform when the attitude error is greater than or equal to a critical value; and when the attitude error is smaller than a critical value, calculating the control moment of the load simulator according to the attitude error, and controlling the active pointing platform to reduce the attitude error between the load simulator and the target attitude of the load simulator according to the control moment of the load simulator. The critical value is the maximum control angle of the actuator.
The star body primary control loop comprises a star body control unit, a star body actuating mechanism and a star body measuring unit which are all fixed on a star body;
the star primary control loop is connected with the load simulator secondary control loop through a platform; the platform provides active control force for a secondary control loop of the load simulator; the load simulator transmits the reaction force of the active control force to the primary control loop of the star body through the platform.
A star measuring unit: the satellite angular velocity and the satellite attitude are measured and fed back to the satellite control unit; the method comprises the following steps: 3M optical fiber gyroscopes, wherein M is a positive integer larger than 3, and at least 3 optical fiber gyroscopes have their measuring directions orthogonal to each other to control the nominal angular momentum h of the moment gyroscope0The following conditions are satisfied:
h0≥Isatωmax/γn
wherein gamma is a gyro group angular momentum coefficient formed by n control moment gyros. I issatIs the integral inertia of the spacecraft, omegamaxIs the maximum angular velocity of the spacecraft.
A star body control unit: and the attitude error between the star attitude and the target attitude is determined according to the target attitude, the star angular velocity and the star attitude sent by the star measuring unit, and the control moment of the star is calculated according to the attitude error.
The star body actuating mechanism: the star control unit is arranged on a star and comprises 4 control moment gyros, and the attitude of the star is adjusted according to the star control moment calculated by the star control unit. The low-speed frame rotating shaft of each control moment gyroscope is in a standard pyramid configuration, and the low-speed frame rotating shaft of each control moment gyroscope is perpendicular to the center lines of two edge sides of the pyramid configuration.
The active pointing platform comprises 3 actuator groups, and the mounting points of the 3 actuator groups are uniformly distributed on a circumference. Each actuator group comprises 2 actuators. The actuator comprises a flexible hinge and a motor stator; one end of the flexible hinge of the actuator is fixed at a mounting point p of the load simulatoriSix mounting points p of load simulator, i is 1,2, …,6iCoplanar; mounting point b of actuator motor stator end and stariFixed, six star mounting points biCoplanar.
And the displacement sensor is arranged on the platform actuator and used for measuring the translational displacement of the active pointing platform actuator, and the displacement sensor is an eddy current sensor or a capacitive sensor. The star body control unit calculates the relative posture between the load simulator and the star body according to the translation displacement of the platform actuator, and the star body executing mechanism controls the posture of the star body according to the relative posture to keep the relative posture between the load simulator and the star body unchanged.
The active pointing platform actuator also comprises a flexible passive link and an active control link; the flexible passive link is used for isolation and suppression of high-frequency micro-vibration of the star, and the active control link is used for active isolation and suppression of pointing active control force of load simulation and low-frequency micro-vibration of the star. The output bandwidth of the active control link is more than 1kHz, and the accuracy of the output control force is better than 0.001N.
Minimum resolution epsilon of displacement sensor on active pointing platform actuatorl0Satisfies the following conditions:
wherein, I3×3A 3 × 3 unit array; epsilonl＝[ε1 … εl]TThe measurement precision of a displacement sensor of the platform is I.
Output force f of active pointing platform actuator0Satisfies the following conditions:
in the formula, amaxIs the maximum angular acceleration of the spacecraft; i ispcThe maximum value of the three-axis moment of inertia of the load simulator. J. the design is a squarepIs a load centroid Jacobian matrix, IpcThe largest moment of inertia in the three axes of the load simulator.
The load simulator secondary control circuit includes: the device comprises a load control unit, a load executing mechanism, a load measuring unit and a load gravity unloading device;
the load simulator is used for simulating the mass and inertia characteristics of the load;
load gravity unloading device: gravity for unloading a load simulator, comprising: a spring, a support structure; the supporting structure comprises a flat plate, the end face of the flat plate is provided with a plurality of supporting rods, the supporting rods form a conical structure, and the other ends of the supporting rods are fixed on the top layer of the star body; the flat plate structure fixes one end of the spring, the other end of the spring is hoisted on the load simulator, and the bottom of the load simulator is fixed with the platform actuator of the platform.
A load executing mechanism: adjusting the attitude of the load simulator by an active control force provided by a platform actuator of the platform;
a load control unit: the load simulator is arranged on the load simulator and used for controlling the posture of the load simulator;
a load measuring unit: and the load simulator is fixed on the load simulator and used for measuring the angular speed of the load simulator. The method comprises the following steps: two autocollimators, a cubic mirror and an inertial sensor; the cubic mirror is installed on a shell of the load simulator, optical axes of the two autocollimators point to two adjacent end faces of the cubic mirror, and the inertia sensor is installed on the load simulator and used for measuring the angular speed of the satellite load simulator.
Autocollimator measurement accuracy muθInertia sensor, namely load micrometer sensor, measurement precision muωThe following conditions are satisfied:
wherein epsilonθFor the accuracy of the directional control of the load simulator, epsilonOmega isStability, λ, of load simulatorsθMeasuring the precision coefficient for the autocollimator; lambda [ alpha ]ωThe measurement precision coefficient of the load micrometer sensor is obtained.
Measurement range l of the eddy current sensor of the load attitude measurement unit0Satisfies the following conditions:
wherein, thetabp＝[θbpx θbpy θbpz]TThe three-axis relative attitude angle of the star body and the load simulator is obtained; j. the design is a squarepA load simulator centroid Jacobian matrix; δ L ═ δ L1 … δLl]TIs the translational displacement of the actuators of the platform.
Star actuating mechanism CMGs: the maximum inertia of the three-super platform whole satellite three-axis is Isat＝8000kgm2The maximum angular velocity of the whole-satellite agile motor is omegamaxThe number of CMGs is 4, and the coefficient of angular momentum gamma of the gyro group formed by the CMGs is 1.2, so that the nominal angular momentum h of each CMGs is 4 (DEG/s)0It should satisfy:
h0≥Isatωmax/γn≈116(Nms)
selecting the angular momentum h of CMG0＝125Nms。
The three-axis maximum inertia of the active pointing platform load is Ipc＝140kgm2The maximum relative angular acceleration of the load and the star platform is amax＝0.5(°/s2). Each voice coil motor outputs a force f0It should satisfy:
the load pointing control precision of the load micrometer sensor is epsilonθ＝0.1″，λθ0.1; stability of epsilonω＝1×10-4(°/s)，λ ω1. Then the measurement accuracy mu of the load autocollimatorθ<0.01', load micrometer sensor measuring precision muω<1×10-4(° s). Selecting a photoelectric autocollimator with measurement noise better than 0.01' and selecting an angle random walk better than 1 x 10-4Load micrometer sensor(s).
The maximum relative attitude between the load and the star platform is θpb0.3 °, the measurement range l of the eddy current sensor0Due to the satisfaction of0>2 mm. The measuring range of the eddy current sensor is selected to be 3 mm.
As shown in fig. 2, the load simulator mounting surface radius rp0.53m, star mounting surface radius rB0.53m, and the positioning angle theta of the mounting surface of the load simulatorp5 degrees, star mounting surface positioning angle thetaB36 ° and the height H of the active pointing platform 0.14 m. Active pointing platform and load simulator mounting point piAnd a star mounting point biThe calculation is as follows:
p1＝[-rP cos(θP/2),rP sin(θP/2),H]T
p2＝[-rP cos(θP/2),-rP sin(θP/2),H]T
p3＝[rP sin(30-θP/2),-rP cos(30-θP/2),H]T
p4＝[rP sin(30+θP/2),-rP cos(30+θP/2),H]T
p5＝[rP sin(30+θP/2),rPcos(30+θP/2),H]T
p6＝[rP sin(30-θP/2),rP cos(30-θP/2),H]T
b1＝[-rB cos(θB/2),rB sin(θB/2),0]T
b2＝[-rB cos(θB/2),-rB sin(θB/2),0]T
b3＝[rB sin(30-θB/2),-rB cos(30-θB/2),0]T
b4＝[rB sin(30+θB/2),-rB cos(30+θB/2),0]T
b5＝[rB sin(30+θB/2),rB cos(30+θB/2),0]T
b5＝[rB sin(30-θB/2),rB cos(30-θB/2),0]T
jacobian matrix J of load centroidpAnd the star centroid Jacobian matrix JbThe calculation is as follows:
substituting the specific numerical value to calculate a load centroid Jacobian matrix JpAnd the star centroid Jacobian matrix JbThe calculation is as follows:
mounting array C of star measuring unitgbExpressed in terms of star centroid coordinates as:
as shown in FIG. 3, the measuring axes of the load simulator micrometer sensor of the present invention are arranged in a triangular pyramid structure (OG)1,OG2,OG3Three measuring shafts of the load simulator micrometer sensor respectively), and a mounting array CgpExpressed in the load centroid coordinate system as:
in the formula, θ is 54.7356 ° and represents three measurement axes OG1,OG2,OG3With the load coordinate zpThe included angle of the axes; beta-30 deg. is Gp2、Gp3Under load xp-op-ypProjection of plane and ypThe included angle of (a).
The load control unit runs a load controller program and is core hardware of a closed-loop control system of the active pointing hyperstatic platform actuator. The device mainly comprises a 16-bit A/D converter, a 16-bit D/A converter, a processor and a power management module. The 16-bit A/D converter collects signals of an eddy current sensor in the actuator, transmits measurement signals to the processor, transmits the signals to the 16-bit D/A converter after operation processing, and outputs the signals to a driver in the intelligent flexible actuator, and the power management module provides power required by collection of the 16-bit A/D converter, operation of the processor and output of the 16-bit D/A converter.
And the star controller adopts an industrial personal computer to operate a control program. The industrial personal computer is provided with a plurality of serial ports and is used for communicating with the load controller, receiving load information, vortex information and the like, and sending star target attitude information and the like.
The actuator voice coil motor driver is used for receiving the control instruction of the load controller and driving the voice coil motor to realize control output. The driver in the active pointing hyperstatic platform actuator is equivalent to a voltage/current conversion link, and outputs voltage by receiving a controller in the intelligent flexible actuator, converts the voltage into current and outputs the current to the voice coil motor.
And a star gyroscope (a gyroscope coordinate system is parallel to the star coordinate system) and a control moment gyroscope group are arranged on the star platform. According to load mounting point piAnd (4) mounting the active directional hyperstatic platform and the load simulator. According to the star body mounting point biAnd carrying out active directional hyperstatic platform/load integral and star body installation. And installing a load and load gravity unloading bracket. And the load zero-gravity unloading is realized by adjusting the length of the gravity unloading support spring.
The simulation system is adopted to carry out a three-super platform full-physical simulation test and verify the control performance index of the three-super platform. Figure 4 gives the star and load stability test results. During stable control, the three-axis attitude error of the star is less than 2 ', the three-axis attitude error of the load simulator is less than 0.05', and therefore the attitude error of the load simulator and the target attitude is greatly superior to that of the star and the target attitude. Fig. 5 shows the results of the star and load agile maneuver tests. The star and the load can realize large-angle quick maneuvering of 0-20 degrees within 408 s-418.3 s, and the postures of the star and the load can realize agile maneuvering of 4 degrees/s. After the agile maneuver of the whole satellite attitude, the steady-state time of the agile maneuver of the load simulator is better than 4s under the control of the 'three-super' platform.
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 (8)
1. A spacecraft two-stage attitude control simulation system, comprising: a star body primary control loop, a load simulator secondary control loop, an active pointing platform, a star body and a load simulator; the active pointing platform is fixed on the star body, and the load simulator is fixed on the active pointing platform;
star first-level control loop: judging the attitude error between the attitude of the star body and the target attitude of the load simulator, and when the attitude error is greater than or equal to a critical value, calculating the control moment of the star body according to the attitude error to reduce the attitude error between the star body and the target attitude; when the attitude error is smaller than a critical value, controlling the star body to maintain the attitude unchanged;
a second-level control loop of the load simulator: judging the attitude error between the load simulator and the target attitude of the load simulator, and controlling the relative attitude between the load simulator and the star body to be kept unchanged through the active pointing platform when the attitude error is greater than or equal to a critical value; when the attitude error is smaller than a critical value, calculating the control moment of the load simulator according to the attitude error, and controlling the active pointing platform to reduce the attitude error between the load simulator and the target attitude of the load simulator according to the control moment of the load simulator;
the star body primary control loop comprises a star body control unit, a star body actuating mechanism and a star body measuring unit which are all fixed on a star body;
a star measuring unit: measuring the angular velocity and the posture of the star body and feeding back to the star body control unit;
a star body control unit: determining an attitude error between the star attitude and the target attitude of the load simulator according to the target attitude of the load simulator and the star angular velocity and the star attitude sent by the star measuring unit, and calculating the control moment of the star according to the attitude error;
the star body actuating mechanism: the star control device comprises a plurality of control moment gyroscopes, wherein star postures are adjusted according to star control moments calculated by a star control unit;
the star measurement unit includes: 3M optical fiber gyroscopes, wherein M is a positive integer larger than 3, and the measuring directions of at least 3 optical fiber gyroscopes are pairwise orthogonal.
2. A spacecraft two-stage attitude control simulation system according to claim 1, wherein the active pointing platform comprises 3 actuator groups, each actuator group comprising 2 actuators.
3. A spacecraft two-stage attitude control simulation system according to claim 1, wherein: the number of the control moment gyroscopes is 4, the low-speed frame rotating shaft of each control moment gyroscope is in a standard pyramid configuration, and the low-speed frame rotating shaft of each control moment gyroscope is perpendicular to the center lines of two edges of the pyramid configuration.
4. A spacecraft two-stage attitude control simulation system according to claim 2, wherein the load simulator two-stage control loop comprises: the device comprises a load control unit, a load executing mechanism, a load measuring unit and a load gravity unloading device;
a load measuring unit: measuring the angular speed and the attitude of the load simulator and feeding back to the load control unit;
a load control unit: determining an attitude error between the attitude of the load simulator and a target attitude of the load simulator according to the angular velocity of the load simulator and the attitude of the load simulator, and calculating a control moment of the load simulator; controlling the active pointing platform to reduce the attitude error between the load simulator and the target attitude according to the control moment of the load simulator;
load gravity unloading device: the load simulator is arranged on the star body, and the load simulator is hoisted at the top of the star body and unloads the gravity of the load simulator;
a load executing mechanism: the attitude of the load simulator is adjusted by active control forces provided by actuators of the active pointing platform.
5. A spacecraft two-stage attitude control simulation system according to claim 4, wherein the actuators comprise: a flexible hinge and a motor stator; mounting point p for fixing one end of each of 6 actuator flexible hinges on load simulatoriSix mounting points p of load simulator, i is 1,2, …,6iCoplanar; mounting point b for fixing one end of motor stator of 6 actuators on stariMounting points b of six starsiCoplanar.
6. A spacecraft two-stage attitude control simulation system according to claim 5, wherein the load gravity unloading device of the load simulator two-stage control loop comprises: a spring, a support structure; the supporting structure comprises a flat plate, the end face of the flat plate is provided with a plurality of supporting rods, the supporting rods form a conical structure, and the other ends of the supporting rods are fixed on the star body; one end of the spring is fixed by the flat plate structure, and the other end of the spring is used for hoisting the load simulator.
7. A spacecraft two-stage attitude control simulation system according to claim 4, wherein the measurement unit of the load simulator two-stage control loop comprises: two autocollimators, a cubic mirror and an inertial sensor; the cubic mirror is installed on a shell of the load simulator, optical axes of the two autocollimators point to two adjacent end faces of the cubic mirror, and the inertia sensor is installed on the load simulator and used for measuring the angular speed of the satellite load simulator.
8. A spacecraft two-stage attitude control simulation system according to any of claims 2 or 4 to 7, wherein: the critical value is the maximum control angle of the actuator.
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| CN111605734B (en) * | 2020-04-28 | 2021-11-16 | 北京控制工程研究所 | Filtering timing system and method for accurate synchronization of spacecraft three-super control satellite time |
| CN111638721B (en) * | 2020-04-28 | 2023-08-11 | 北京控制工程研究所 | Spacecraft three-ultra-control full-link disturbance transfer verification system and verification method |
| CN111781943B (en) * | 2020-07-20 | 2024-04-12 | 北京控制工程研究所 | Three-override control method for distributed load pose of spacecraft |
| CN114197641A (en) * | 2022-01-07 | 2022-03-18 | 西安电子科技大学 | Active tensioning cable net truss structure |
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