Load optical axis pointing control system and method of truss type flexible spacecraft

Abstract

The invention relates to the field of spacecraft attitude control, in particular to a load optical axis pointing control system and method for a truss type flexible spacecraft, and the system comprises a truss type flexible spacecraft structure platform, an attitude control module and a load optical axis pointing control module. The attitude control module is used for central star attitude angle fine adjustment and preliminary adjustment of optical load primary mirror optical axis direction. The load optical axis direction control module is used for optical axis direction fine adjustment and micro-vibration suppression. Two independent closed loops of central star attitude angle closed-loop control and load optical axis pointing closed-loop control are constructed, input signal sources, control objects and execution mechanisms of the two control loops are completely decoupled, and the problem of control effect divergence caused by mutual interference in traditional coupling control is solved in principle.

Description

Technical Field

[0001] The present invention relates to the field of spacecraft attitude control, and in particular to a load optical axis pointing control system and method for a truss-type flexible spacecraft. Background Art

[0002] In the field of aerospace technology, truss-type flexible spacecraft, with its advantages of light structural weight and large unfolded size, demonstrates unique value when carrying high-precision optical payloads to perform space observation and remote sensing missions.

[0003] Large-aperture optical payloads are critical for satellites and other spacecraft to acquire cosmic information. Their pointing accuracy directly determines the quality of the observation data. For example, Earth remote sensing missions require high-precision pointing to the target area to obtain clear surface images for resource monitoring and environmental assessment. However, the flexible structure of truss-type flexible spacecraft poses significant challenges to controlling the pointing of optical payloads.

[0004] Existing relevant patents, such as patent document CN109002047B (application number 201810589560.3), "A method and process for multi-stage composite control of spacecraft with coarse and fine layered fast and slow integration of main and passive components," proposes the construction of a multi-stage composite control system for spacecraft, including a satellite, payload, sensors, actuators, etc., and the design of a first-stage PID controller for the satellite's attitude, a second-stage PID controller for the active pointing ultra-quiet platform's attitude, and a third-stage PID controller for the fast-reflecting mirror's attitude, to achieve multi-stage control of the spacecraft's attitude. Although this solution improves control accuracy to a certain extent, achieving approximately 20dB of attenuation of high-frequency disturbances (>10Hz) on the satellite platform through the passive link of the active pointing ultra-quiet platform, and improving payload stability to a certain extent, it is difficult to implement effective active control of flexible vibrations and high-frequency micro-vibrations due to limitations in controller bandwidth and actuator accuracy, hindering further improvements in control accuracy and stability.

[0005] Another example is patent document CN118907441A (application number 202411103618.0), "A flexible spacecraft attitude control method with dual vibration suppression and self-disturbance rejection based on shaping programming." For flexible spacecraft with external interference and model uncertainty, it introduces self-disturbance rejection technology, constructs a third-order extended observer to observe the system attitude angular velocity and total disturbance, realizes dynamic compensation linearization and disturbance suppression, and designs a singularity avoidance control law for the singularity problem of the single-frame control moment gyroscope group. However, this solution does not fully consider the combined impact of model uncertainty and the actuator singularity avoidance problem on control accuracy, and the control effect under complex working conditions needs to be further improved.

[0006] Looking at patent document CN111625010A (application number 202010350592.5) "A spacecraft three-level near-zero error tracking control method and process based on combined filtering", this patent establishes a multi-level system dynamics model of the spacecraft, adopts combined filtering and deep learning technology, and realizes the rapid solution of the position and attitude of the characteristic parts of the target spacecraft, and gradually improves the attitude control accuracy from the three-level system of the satellite, payload, and fast reflection mirror. However, in actual application, it still faces the contradiction of high-precision attitude determination and high dynamic performance. It cannot fully meet the stringent requirements of spacecraft for high-precision, high-stability and ultra-agile control in complex space environments.

[0007] In summary, considering the complex structure of truss-type flexible spacecraft and the high-precision pointing requirements of optical payloads, it is of great practical significance and urgent engineering need to study an efficient, accurate and robust truss-type flexible spacecraft optical payload pointing control method and system. Summary of the Invention

[0008] In view of the defects in the prior art, the purpose of the present invention is to provide a load optical axis pointing control system and method for a truss-type flexible spacecraft.

[0009] According to the present invention, a load optical axis pointing control system for a truss-type flexible spacecraft includes:

[0010] A truss-type flexible spacecraft structure platform is used to support the core structure of multiple modules. The truss-type flexible spacecraft structure platform includes an optical payload secondary mirror, a lightweight truss, a central star body, and an optical payload primary mirror.

[0011] An attitude control module is used for fine-tuning the attitude angle of the central star and preliminary adjustment of the optical axis direction of the optical payload primary mirror. The attitude control module includes an attitude control flywheel, a central star autocollimator prism, a central star autocollimator and an attitude angle controller.

[0012] The load optical axis pointing control module is used for fine adjustment of the optical axis pointing and micro-vibration suppression. The load optical axis pointing control module includes a macro-micro actuator, an acceleration sensor, a displacement sensor, a primary mirror autocollimator prism, a primary mirror autocollimator, a macro-micro actuator controller and a truss heating belt.

[0013] Furthermore, the central star body includes three cabins. The central star body and the optical payload primary mirror are connected by a lightweight truss to form a triangular prism configuration. Through the symmetrical support of the three sides, the optical payload primary mirror is uniformly stressed in the three-dimensional direction of space, effectively suppressing the flexible deformation of the optical system.

[0014] Furthermore, the central star body and the optical payload secondary mirror are connected by a lightweight truss to form a triangular pyramid configuration, with the central star body as the vertex base, and the optical payload secondary mirror forms a spatial triangular stable structure with the vertex through three truss legs, which significantly improves the vibration and impact resistance of the optical payload secondary mirror installation.

[0015] Furthermore, the input signal of the attitude angle controller comes from the central star body, and the input signal of the macro-micro actuator controller comes from the lightweight truss and the optical payload primary mirror. The two types of closed-loop control loops, the attitude angle of the central star body and the direction of the payload optical axis, are independently decoupled.

[0016] Furthermore, the macro-micro actuator is installed on the lightweight truss, and the macro-micro actuator includes a macro unit and a micro unit. The macro unit is used to adjust the displacement of the optical load primary mirror, and the micro unit is used to suppress the micro vibration of the lightweight truss.

[0017] Furthermore, the truss heating belt is arranged on the three trusses of the optical load primary mirror section, and the truss heating belt is used to simulate the structural deformation caused by the space thermal environment.

[0018] The present invention also provides a method for controlling the pointing direction of a load optical axis of a truss-type flexible spacecraft, comprising the following steps:

[0019] Step S1: Fine-tuning the attitude angle of the central star: A closed loop of the attitude angle of the central star is formed through the attitude control flywheel, the central star autocollimator prism, the central star autocollimator, and the attitude angle controller. The attitude angle optical signal of the central star is obtained through the central star autocollimator prism and the central star autocollimator, and the attitude angle measurement signal is output as the input signal of the attitude angle controller. The attitude angle controller outputs a control signal of the attitude control flywheel to drive the attitude control flywheel to rotate, control the central star to rotate to the target attitude angle, and the attitude control flywheel rotates and drives the central star to gradually maneuver to the target attitude angle, preliminarily adjusting the direction of the payload optical axis;

[0020] Step S2: Coarse adjustment of the optical axis pointing of the payload: Based on the result of the attitude angle adjustment of the central star, the optical axis pointing is roughly adjusted through the macro-micro actuator so that the optical axis of the optical payload is preliminarily adjusted to the target pointing direction;

[0021] Step S3: Fine-tuning the load optical axis pointing: A closed-loop control of the load optical axis pointing is formed through the displacement sensor, the primary mirror autocollimator prism, the primary mirror autocollimator, the acceleration sensor, the macro-micro actuator controller, and the macro-micro actuator. The displacement sensor, the primary mirror autocollimator prism, the primary mirror autocollimator, and the acceleration sensor are used to monitor the primary mirror displacement measurement signal, the optical axis pointing measurement signal, and the acceleration measurement signal of the lightweight truss in real time during the spacecraft attitude maneuver. The above signals are then input into the macro-micro actuator controller, which outputs a control signal to control the actuation of the macro-micro actuator, suppress vibration, and accurately adjust the optical axis to the target angle.

[0022] Step S4: After the central star attitude angle, primary mirror displacement deviation, truss vibration response level and optical axis pointing deviation all reach the target values, the control process ends.

[0023] Furthermore, the attitude angle controller adopts a PID algorithm with an additional saturation constraint, and the algorithm expression is:

[0024]

[0025] Where J is the moment of inertia matrix of the controlled object; k corresponds to the control parameter of the proportional link of the PID algorithm; T corresponds to the control parameter of the integral link of the PID algorithm; c is the control parameter corresponding to the differential link of the PID algorithm, k = 9.54, T = 10 and c = 0.55; is the control torque vector; L i is the saturation threshold of the sub-dimension, λ is the attitude angle error; ω is the rotation angular velocity vector of the spacecraft; sat(*) is the saturation state function, which is defined as follows:

[0026]

[0027] Among them, e represents the current error value of the saturation state of the attitude angle controller. By comparing it with the saturation threshold L of the sub-dimension under the saturation state of the attitude angle controller, it is determined whether the output enters the saturation state.

[0028] Furthermore, the control algorithm of the macro-micro actuator controller includes:

[0029] Based on the relative distance S of 15 observation points ij Observations of:

[0030] The state equation of the fourth-order Bessel low-pass filter with a cutoff frequency of 0.5Hz is:

[0031]

[0032] Where x is the state vector of the low filter; is the rate of change of the state vector of the low-pass filter; y is the input signal of the low-pass filter;

[0033] Force calculation formula

[0034] in, is the measurement of relative deformation after passing through the low-pass filter; The time differential of the relative deformation measurement after passing through the low-pass filter; A y and A v is the control rate matrix related to the modal characteristics of the flexible spacecraft;

[0035] Furthermore, the closed-loop control of the central star attitude angle and the closed-loop control of the payload optical axis pointing are independently decoupled from each other to avoid control coupling divergence.

[0036] Compared with the prior art, the present invention has the following beneficial effects:

[0037] 1. The present invention constructs two independent closed loops: the central satellite attitude angle closed loop control and the payload optical axis pointing closed loop control. The input signal sources, control objects and actuators of the two control loops are completely decoupled, which in principle avoids the problem of control effect divergence caused by mutual interference in traditional coupling control.

[0038] 2. The present invention adopts a staged control process of "fine adjustment of the central star body attitude angle → coarse adjustment of the payload optical axis pointing → fine adjustment of the payload optical axis pointing", first realizing global calibration of the entire satellite attitude, and then performing local fine adjustment of the optical payload, which significantly improves the control efficiency and convergence speed. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] Other features, objects and advantages of the present invention will become more apparent upon reading the detailed description of non-limiting embodiments with reference to the following drawings:

[0040] Figure 1 Schematic diagram of the load optical axis pointing control system of the truss-type flexible spacecraft described in the present invention.

[0041] Figure 2 This is a flow chart of the load optical axis pointing control method of the truss-type flexible spacecraft described in the present invention.

[0042] Description of reference numerals:

[0043] DETAILED DESCRIPTION

[0044] The present invention will be described in detail below with reference to specific embodiments. The following examples will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those skilled in the art, several changes and improvements can be made without departing from the scope of the present invention. These all fall within the scope of protection of the present invention.

[0045] like Figure 1 As shown, the present invention provides a load optical axis pointing control system for a truss-type flexible spacecraft, comprising:

[0046] The truss-type flexible spacecraft structure platform is used for the core structure support of multiple modules. The truss-type flexible spacecraft structure platform includes an optical payload secondary mirror 1, a lightweight truss 2, a central star body 3 and an optical payload primary mirror 9. The central star body 3 includes three cabins, which serve as the core hub of the entire structure and provide installation references and support foundations for various functional modules. The lightweight truss 2 is designed with high-strength lightweight materials and undertakes multiple connection functions. The lightweight truss 2 is used to connect the three cabins of the central star body 3 with rigid connections to ensure the integrity of the structure. The lightweight truss 2 connects the optical payload secondary mirror 1 and the central star body 3, the optical payload primary mirror 9 and the central star body 3 respectively. The mirror 9 and the central star body 3 form a stable force transmission path; the central star body 3 and the optical payload primary mirror 9 are connected by a lightweight truss 2 to form a triangular prism configuration. Through the symmetrical support of the three sides, the optical payload primary mirror 9 is uniformly stressed in the three-dimensional direction of space, effectively suppressing the flexible deformation of the optical system; the central star body 3 and the optical payload secondary mirror 1 are connected by a lightweight truss 2 to form a triangular pyramid configuration, with the central star body 3 as the vertex base. The optical payload secondary mirror 1 forms a spatial triangular stable structure with the vertex through three truss legs, which significantly improves the vibration and impact resistance of the installation of the optical payload secondary mirror 1.

[0047] The attitude control module is used to achieve precise adjustment of the attitude angle of the central satellite 3 of the truss-type flexible spacecraft and preliminary adjustment of the optical axis direction of the optical payload primary mirror 9. The attitude control module includes an attitude control flywheel 4, a central satellite autocollimator prism 5, a central satellite autocollimator 7 and an attitude angle controller 20. The attitude control flywheel 4 is installed on the central satellite 3 and serves as an actuator for attitude adjustment. It generates torque by rotation to drive the central satellite 3 to perform attitude angle maneuvers; the attitude angle controller 20 serves as a core control unit, receives the central satellite attitude angle measurement signal 19 and outputs control instructions to achieve closed-loop control of the attitude control flywheel 4; the central satellite autocollimator prism 5 is fixed on the surface of the central satellite 3 and serves as a reference reflection element for attitude angle measurement. It changes the reflected light path as the attitude of the central satellite 3 changes; the central satellite autocollimator 7 cooperates with the central satellite autocollimator prism 5 to obtain the real-time attitude angle deviation of the central satellite 3 through optical measurement and output the attitude angle measurement signal 19 of the central satellite 3;

[0048] The payload optical axis pointing control module is used for fine adjustment of the optical axis pointing and micro-vibration suppression. The payload optical axis pointing control module includes a macro-micro actuator 6, an acceleration sensor 8, a displacement sensor 10, a primary mirror autocollimator prism 12, a primary mirror autocollimator 13, a macro-micro actuator controller 16, and a truss heating belt 18. The displacement sensor 10 is installed on the optical payload primary mirror 9 to measure the displacement of the optical payload primary mirror 9 relative to the central star 3 in real time and output a displacement measurement signal 11. The primary mirror autocollimator prism 12 and the primary mirror autocollimator 13 serve as optical measurement components to accurately obtain the real-time pointing angle of the optical axis of the optical payload primary mirror 9 and output an optical axis pointing signal 15. The acceleration sensor 8 is installed on the lightweight truss 2 to monitor the acceleration measurement signal 14 of the structural vibration response, reflecting the micro-vibration level; the macro-micro actuator controller 16 serves as the core processing module, receives the displacement, acceleration and optical axis pointing signals, solves and generates the macro-micro actuator control signal 17, and drives the macro-micro actuator 6 to perform adjustments; the truss heating belt 18 is arranged on two of the three trusses of the optical payload primary mirror 9 section, and simulates the uneven thermal environment in space by energizing and heating, inducing thermal deformation of the lightweight truss 2, and then simulating the displacement offset of the optical payload primary mirror 9 caused by thermal stress; the load optical axis pointing control module is used to achieve precise adjustment of the optical axis pointing of the optical payload.

[0049] like Figure 2 As shown, the load optical axis pointing control method of the truss-type flexible spacecraft provided by the present invention specifically includes the following steps:

[0050] Step S1: Fine-tune the attitude angle of the central star:

[0051] Turn on the attitude control flywheel 4, the truss heating belt 18, and the macro-micro actuator 7. The attitude control flywheel 4, the central star autocollimator prism 5, the central star autocollimator 7, and the attitude angle controller 20 form a closed loop of the central star attitude angle. The attitude angle optical signal of the central star 3 is obtained through the central star autocollimator prism 5 and the central star autocollimator 7, and the attitude angle measurement signal 19 is output as the input signal of the attitude angle controller 20. The attitude angle controller 20 outputs the attitude control flywheel control signal 21 to drive the attitude control flywheel 4 to rotate, control the central star 3 to rotate to the target attitude angle, and the attitude control flywheel 4 rotates and drives the central star 3 to gradually maneuver to the target attitude angle, preliminarily adjusting the direction of the payload optical axis;

[0052] The attitude angle controller 20 adopts a PID algorithm with an additional saturation constraint, which is expressed as follows:

[0053]

[0054] a i =M / J i ;

[0055] Where J is the moment of inertia matrix of the controlled object; J i is the i-th component of the moment of inertia of the controlled object; k corresponds to the control parameter of the proportional link of the PID algorithm; T corresponds to the control parameter of the integral link of the PID algorithm; c is the control parameter corresponding to the differential link of the PID algorithm, k = 9.54, T = 10 and c = 0.55; are the control torque vectors corresponding to the three coordinate axes respectively; is the maximum value of the control torque vector corresponding to the three coordinate axes; L i is the saturation threshold of the sub-dimension, λ = (λ1, λ2, λ3) is the angular error corresponding to the three coordinate axes respectively; λι is the i-th component of the attitude error vector; ω = (ω1, ω2, ω3) is the angular velocity vector of the spacecraft around each coordinate axis; |ω ι | max is the maximum value of the angular velocity vector; m = (m1, m2, m3) is the attitude control torque vector corresponding to the three coordinate axes respectively; M is the maximum value of the attitude control torque allowed to be output by the attitude control flywheel actuator; a i is the maximum angular acceleration in the direction of the i-th coordinate axis; sat(*) is the saturation state function, defined as follows:

[0056]

[0057] Among them, e represents the current error value of the saturation state of the attitude angle controller. By comparing it with the saturation threshold L of the sub-dimension under the saturation state of the attitude angle controller, it is determined whether the output enters the saturation state.

[0058] The algorithm avoids control torque overload through saturation constraints and improves system robustness;

[0059] Step S2: Coarse adjustment of the load optical axis direction:

[0060] Based on the result of the attitude angle adjustment of the central star, the optical axis direction is roughly adjusted by the macro-micro actuator 16, so that the optical axis of the optical payload is preliminarily adjusted to the target direction;

[0061] Step S3: Fine-tune the load optical axis direction:

[0062] After the central star body 3 reaches the target attitude angle, a closed-loop control of the load optical axis pointing is formed through the displacement sensor 10, the primary mirror autocollimator prism 12, the primary mirror autocollimator 13, the acceleration sensor 8, the macro-micro actuator controller 16, and the macro-micro actuator 7. The displacement sensor 10, the primary mirror autocollimator prism 12, the primary mirror autocollimator 13, and the acceleration sensor 8 are used to monitor the primary mirror displacement measurement signal 11, the optical axis pointing measurement signal 15, and the acceleration measurement signal 14 of the lightweight truss in real time during the spacecraft attitude maneuvering process. The above signals are then input into the macro-micro actuator controller 16, which outputs a control signal 17 to control the actuation of the macro-micro actuator 7, suppress vibration, and accurately adjust the optical axis to the target angle;

[0063] The macro-micro actuator controller 16 adopts a fourth-order Bessel low-pass filter control algorithm based on relative distance observation, which is expressed as follows:

[0064] The control algorithm contains 15 observations, which are the relative distance s between any two observation points and are defined as follows:

[0065] s=(s 12 , s 13 , s 14 , s 15 , s 16 , s 23 , s 24 , s 25 , s 26 , s 34 , s 35 , s 36 , s 45 , s 46 , s 56 );

[0066] The relative distance between any two points i and j is:

[0067]

[0068] in, represents the initial coordinates of the i-th observation point; (x i ,y i ,z i ) represents the end point coordinates of the i-th observation point

[0069] The observed quantity is filtered using a fourth-order Bessel low-pass filter with a cutoff frequency of 0.5 Hz. The filter is expressed as the following linear system form:

[0070]

[0071] Where x is the state vector of the low filter; is the rate of change of the state vector of the low-pass filter; y is the input signal of the low-pass filter; is the output signal of the low-pass filter;

[0072] The output force is the thrust and tension along the truss axis (thrust is defined as positive and tension is defined as negative), and the force vector T c The calculation formula is:

[0073]

[0074] in, is the measurement of relative deformation after passing through the low-pass filter; The time differential of the relative deformation measurement after passing through the low-pass filter; A y and A v is the control rate matrix related to the modal characteristics of the flexible spacecraft;

[0075] Step S4: Control end condition:

[0076] The control process ends when the central star attitude angle, primary mirror displacement deviation, truss vibration response level and optical axis pointing deviation all reach the target values.

[0077] In the description of this application, it should be understood that the terms "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", etc., indicating the orientation or position relationship, are based on the orientation or position relationship shown in the accompanying drawings, and are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on this application.

[0078] The above describes specific embodiments of the present invention. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art may make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. The embodiments of this application and the features in the embodiments may be combined with each other in any manner unless there is a conflict.

Claims

1. A load optical axis pointing control system for a truss-type flexible spacecraft, characterized in that: include: A truss-type flexible spacecraft structure platform is used for supporting a core structure of multiple modules. The truss-type flexible spacecraft structure platform comprises an optical payload secondary mirror (1), a lightweight truss (2), a central star body (3) and an optical payload primary mirror (9); An attitude control module is used for fine adjustment of the attitude angle of the central star (3) and preliminary adjustment of the optical axis direction of the optical payload primary mirror (9), the attitude control module comprising an attitude control flywheel (4), a central star autocollimator prism (5), a central star autocollimator (7) and an attitude angle controller (20); The load optical axis pointing control module is used for fine adjustment of optical axis pointing and micro-vibration suppression. The load optical axis pointing control module comprises a macro-micro actuator (6), an acceleration sensor (8), a displacement sensor (10), a primary mirror autocollimator prism (12), a primary mirror autocollimator (13), a macro-micro actuator controller (16) and a truss heating belt (18).

2. The load optical axis pointing control system of a truss-type flexible spacecraft according to claim 1, characterized in that: The central star body (3) comprises three cabins. The central star body (3) and the optical payload primary mirror (9) are connected via a lightweight truss (2) to form a triangular prism configuration. The symmetrical support of the three sides enables the optical payload primary mirror (9) to be uniformly stressed in three-dimensional spatial directions, thereby effectively suppressing the flexible deformation of the optical system.

3. The load optical axis pointing control system of a truss-type flexible spacecraft according to claim 2, characterized in that: The central star (3) and the optical payload secondary mirror (1) are connected via a lightweight truss (2) to form a triangular pyramid configuration. The central star (3) serves as the vertex base, and the optical payload secondary mirror (1) forms a spatial triangular stable structure with the vertex via three truss legs, thereby significantly improving the anti-vibration and anti-impact performance of the optical payload secondary mirror (1) during installation.

4. The load optical axis pointing control system of a truss-type flexible spacecraft according to claim 3, characterized in that: The input signal of the attitude angle controller (20) is derived from the central star body (3), the input signal of the macro-micro actuator controller (6) is derived from the lightweight truss (2) and the optical payload primary mirror (9), and the two types of closed-loop control loops, namely the attitude angle of the central star body and the direction of the optical axis of the payload, are independently decoupled.

5. The load optical axis pointing control system of a truss-type flexible spacecraft according to claim 1, characterized in that: The macro-micro actuator (6) is installed on the lightweight truss (2), and the macro-micro actuator (6) comprises a macro unit and a micro unit, the macro unit is used to adjust the displacement of the optical load primary mirror (9), and the micro unit is used to suppress micro vibrations of the lightweight truss (2).

6. The load optical axis pointing control system of a truss-type flexible spacecraft according to claim 1, characterized in that: The truss heating belt (18) is arranged on three trusses of the optical load primary mirror (9) section, and the truss heating belt (18) is used to simulate structural deformation caused by a space thermal environment.

7. A method for controlling the load optical axis pointing of a truss-type flexible spacecraft, characterized in that: The steps include: Step S1: fine adjustment of the attitude angle of the central star: a closed loop of the attitude angle of the central star is formed by the attitude control flywheel (4), the central star autocollimator prism (5), the central star autocollimator (7) and the attitude angle controller (20), the attitude angle optical signal of the central star (3) is obtained by the central star autocollimator prism (5) and the central star autocollimator (7), and the attitude angle measurement signal (19) is output as the input signal of the attitude angle controller (20), the attitude angle controller (20) outputs the attitude control flywheel control signal (21) to drive the attitude control flywheel (4) to rotate, control the central star (3) to rotate to the target attitude angle, the attitude control flywheel (4) rotates and drives the central star (3) to gradually maneuver to the target attitude angle, and preliminarily adjusts the direction of the load optical axis; Step S2: Coarse adjustment of the optical axis pointing of the payload: Based on the result of the attitude angle adjustment of the central star, the optical axis pointing is roughly adjusted by the macro-micro actuator (16), so that the optical axis of the optical payload is preliminarily adjusted to the target pointing; Step S3: fine adjustment of the load optical axis pointing: a closed-loop control of the load optical axis pointing is formed by the displacement sensor (10), the primary mirror autocollimator prism (12), the primary mirror autocollimator (13), the acceleration sensor (8), the macro-micro actuator controller (16), and the macro-micro actuator (7). The displacement sensor (10), the primary mirror autocollimator prism (12), the primary mirror autocollimator (13), and the acceleration sensor (8) are used to monitor the primary mirror displacement measurement signal (11), the optical axis pointing measurement signal (15), and the acceleration measurement signal (14) of the lightweight truss in real time during the spacecraft attitude maneuvering process. The above signals are then input into the macro-micro actuator controller (16), which outputs a control signal (17) to control the macro-micro actuator (7) to actuate, suppress vibration, and accurately adjust the optical axis to the target angle. Step S4: When the central star attitude angle, primary mirror displacement deviation, truss vibration response level and optical axis pointing deviation all reach the target values, the control process ends.

8. The method for controlling the load optical axis pointing of a truss-type flexible spacecraft according to claim 7, characterized in that: The attitude angle controller (20) adopts a PID algorithm with additional saturation state constraint, and the algorithm expression is: Where J is the moment of inertia matrix of the controlled object; k corresponds to the control parameter of the proportional link of the PID algorithm; T corresponds to the control parameter of the integral link of the PID algorithm; c is the control parameter corresponding to the differential link of the PID algorithm, k = 9.54, T = 10 and c = 0.55; is the control torque vector; L i is the saturation threshold of the sub-dimension, λ is the attitude angle error; ω is the rotation angular velocity vector of the spacecraft; sat(*) is the saturation state function, which is defined as follows: Among them, e represents the current error value of the saturation state of the attitude angle controller. By comparing it with the saturation threshold L of the sub-dimension under the saturation state of the attitude angle controller, it is determined whether the output enters the saturation state.

9. The method for controlling the load optical axis pointing of a truss-type flexible spacecraft according to claim 7, characterized in that: The control algorithm of the macro-micro actuator controller (16) includes: Based on the relative distance S of 15 observation points ij Observations of: The state equation of the fourth-order Bessel low-pass filter with a cutoff frequency of 0.5Hz is: Where x is the state vector of the low filter; is the rate of change of the state vector of the low-pass filter; y is the input signal of the low-pass filter; Force calculation formula in, is the measurement of relative deformation after passing through the low-pass filter; The time differential of the relative deformation measurement after passing through the low-pass filter; A y and A v is the control rate matrix related to the modal characteristics of the flexible spacecraft.

10. The load optical axis pointing control method of a truss-type flexible spacecraft according to claim 7, characterized in that: The closed-loop control of the central star body attitude angle and the closed-loop control of the payload optical axis pointing are independently decoupled from each other to avoid control coupling divergence.

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