Satellite stabilization control analysis method for arbitrary angle of sailboard under large interference moment

Abstract

The application discloses a satellite stability control analysis method for any turning angle of a sailboard under a large interference torque and belongs to the field of satellite control. The method comprises the following steps: calculating a first transfer function matrix of a satellite flexible dynamics system along three-axis attitude control torque to three-axis angle and three-axis angular velocity; respectively calculating an angle measurement transfer function for representing frequency domain characteristics of a sensor, a second transfer function of a satellite controller along three-axis channels for representing frequency domain characteristics of the controller, and a firing control transfer function for representing frequency domain characteristics of an actuator, so as to calculate a transfer function of three-axis attitude control of the satellite; and respectively drawing Nichols charts based on the transfer function of three-axis attitude control of the satellite, so as to determine whether phase margins of the three axes meet stability design requirements, and adjusting control parameters of the satellite controller if the stability design requirements are not met. The application considers the flexible dynamics characteristics of a double-axis solar wing of the satellite, and improves the stability analysis capability of an attitude control system during apogee firing.

Description

Technical Field

[0001] This invention relates to the field of satellite control technology, and in particular to a satellite stability control analysis method for dual-axis solar panels at arbitrary rotation angles under large disturbance torque. Background Technology

[0002] Inclined geosynchronous orbit satellites typically use high-thrust engines for apogee maneuvers during the transfer orbit phase. To ensure the satellite maintains a stable attitude pointing during apogee maneuvers and has a certain margin to resist external disturbances, a stability analysis of the satellite's attitude control system is required in this state. Inclined orbit satellites are equipped with dual-axis solar panels to capture more solar energy and achieve two-dimensional rotation of the solar arrays towards the sun. This factor makes the design conditions during ignition more complex. Furthermore, the use of dual-axis solar panels limits the load-bearing torque at the root of the solar arrays, necessitating an offset of the solar array rotation angle to prevent forced rotation due to attitude fluctuations during ignition.

[0003] Currently, the stability control analysis method for satellites during apogee ignition does not consider the impact of arbitrary rotation angles of the dual-axis solar arrays. This would affect the stability design of the control system and make it difficult to guarantee the stability of the satellite during apogee ignition.

[0004] Therefore, there is an urgent need to provide a satellite stability control analysis method for arbitrary rotation angles of the solar panel under large disturbance moments. Summary of the Invention

[0005] To address the problem that traditional stability control analysis methods do not consider the impact of arbitrary rotation angles of dual-axis solar arrays, making it difficult to guarantee the stability of satellites during apogee ignition, this invention provides a satellite stability control analysis method for arbitrary rotation angles of solar arrays under large disturbance moments.

[0006] On the one hand, a satellite stability control analysis method is provided under large disturbance moment with arbitrary rotation angle of the solar panel, the method comprising:

[0007] Based on the rotation angles of the south and north solar arrays, as well as the installation positions, masses, and moments of inertia of the north and south solar arrays, the first transfer function matrix of the satellite's flexible dynamics system is calculated from the attitude control torques along the three axes to the angles and angular velocities along the three axes. The first transfer function matrix is ​​used to characterize the frequency domain characteristics of the flexible dynamics of the dual-axis solar arrays.

[0008] Calculate the angle measurement transfer function to characterize the frequency domain characteristics of the sensor, the second transfer function of the satellite controller along the three-axis channel to characterize the frequency domain characteristics of the controller, and the ignition control transfer function to characterize the frequency domain characteristics of the actuator, respectively.

[0009] Based on the first transfer function matrix, the angle measurement transfer function, the second transfer function, and the ignition control transfer function, calculate the transfer function for the satellite's three-axis attitude control;

[0010] Nichols plots were drawn based on the transfer functions of the satellite's three-axis attitude control to determine whether the phase margin of the three axes met the stability design requirements. If not, the control parameters of the satellite controller were adjusted.

[0011] On the other hand, a satellite stability control analysis device is provided for arbitrary rotation angle of the solar panel under large disturbance torque, used to implement the steps described in any method embodiment of the specification, the device comprising:

[0012] The flexible dynamics frequency domain characteristic unit is used to calculate the first transfer function matrix of the satellite flexible dynamics system from the three-axis attitude control torque to the three-axis angle and three-axis angular velocity based on the rotation angles of the south and north solar arrays, as well as the installation position, mass and moment of inertia of the north and south solar arrays; the first transfer function matrix is ​​used to characterize the flexible dynamics frequency domain characteristics of the dual-axis solar arrays;

[0013] The calculation unit is used to calculate the angle measurement transfer function, which characterizes the frequency domain characteristics of the sensor; the second transfer function of the satellite controller along the three-axis channel, which characterizes the frequency domain characteristics of the controller; and the ignition control transfer function, which characterizes the frequency domain characteristics of the actuator.

[0014] The attitude calculation unit is used to calculate the transfer function of the satellite's three-axis attitude control based on the first transfer function matrix, the angle measurement transfer function, the second transfer function, and the ignition control transfer function.

[0015] The stability analysis unit is used to plot Nichols diagrams based on the transfer functions of the satellite's three-axis attitude control to determine whether the phase margin of the three axes meets the stability design requirements. If not, the control parameters of the satellite controller are adjusted.

[0016] On the other hand, a computer device is provided, the computer device including a memory and a processor, the memory for storing a computer program, and the processor for executing the computer program stored in the memory to implement the steps of the method described above.

[0017] On the other hand, a computer-readable storage medium is provided, wherein a computer program is stored therein, and when the computer program is executed by a processor, it implements the steps of the method described above.

[0018] On the other hand, a computer program product is provided, including a computer program that, when executed by a processor, implements the steps of the method described above.

[0019] The technical solution provided by this invention can bring at least the following beneficial effects:

[0020] Based on the rotation angles of the south and north solar arrays, as well as the installation positions, masses, and moments of inertia of the north and south solar arrays, the first transfer function matrix used to characterize the flexible dynamic frequency domain characteristics of the dual-axis solar arrays is calculated. Furthermore, considering the flexible dynamic characteristics of the satellite using dual-axis solar arrays, the frequency domain stability of the satellite attitude control system under the influence of large thrust disturbance torque and arbitrary rotation angles of the dual-axis solar arrays during apogee ignition is ensured. Attached Figure Description

[0021] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0022] Figure 1 This is a flowchart of a satellite stability control analysis method under arbitrary rotation angle of solar panels under large disturbance torque, provided by an embodiment of the present invention;

[0023] Figure 2 This is a Nichols diagram of the rolling channel of a satellite attitude control system provided in an embodiment of the present invention;

[0024] Figure 3 This is a Nichols diagram of the pitch channel of a satellite attitude control system provided in an embodiment of the present invention;

[0025] Figure 4 This is a Nichols diagram of the yaw channel of a satellite attitude control system provided in an embodiment of the present invention;

[0026] Figure 5 This is a structural diagram of a satellite stability control and analysis device for arbitrary rotation angle of a solar panel under large disturbance torque, provided in an embodiment of the present invention.

[0027] Figure 6 This is a hardware architecture diagram of a computer device provided in an embodiment of the present invention. Detailed Implementation

[0028] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some embodiments of the present invention, but not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0029] The following describes the specific implementation of the above concept.

[0030] Please refer to Figure 1 This invention provides a satellite stability control analysis method under large disturbance moment and arbitrary rotation angle of solar panels. The method includes:

[0031] Step 100: Based on the rotation angles of the south and north solar array axes, as well as the installation positions, masses, and moments of inertia of the north and south solar arrays, calculate the first transfer function matrix of the satellite's flexible dynamics system from the attitude control torque along the three axes to the three-axis angles and three-axis angular velocities; the first transfer function matrix is ​​used to characterize the frequency domain characteristics of the flexible dynamics of the dual-axis solar array.

[0032] Step 102: Calculate the angle measurement transfer function to characterize the frequency domain characteristics of the sensor, the second transfer function of the satellite controller along the three-axis channel to characterize the frequency domain characteristics of the controller, and the ignition control transfer function to characterize the frequency domain characteristics of the actuator, respectively.

[0033] Step 104: Calculate the transfer function for the satellite's three-axis attitude control based on the first transfer function matrix, the angle measurement transfer function, the second transfer function, and the ignition control transfer function;

[0034] Step 106: Plot Nichols diagrams based on the transfer functions of the satellite's three-axis attitude control to determine whether the phase margin of the three axes meets the stability design requirements. If not, adjust the control parameters of the satellite controller.

[0035] In this embodiment of the invention, based on the rotation angles of the south and north solar array axes, as well as the installation positions, masses, and moments of inertia of the north and south solar arrays, a first transfer function matrix is ​​calculated to characterize the flexible dynamic frequency domain characteristics of the dual-axis solar array. Furthermore, the flexible dynamic characteristics of the satellite using dual-axis solar arrays are considered to ensure the frequency domain stability of the satellite attitude control system under the influence of large thrust disturbance torque and arbitrary rotation angles of the dual-axis solar arrays during apogee ignition.

[0036] The following description Figure 1 The execution method for each step is shown.

[0037] For step 100:

[0038] In some implementations, step 100 may include S1-S5:

[0039] S1, based on the rotation angles of the A and B axes of the south solar wing and the north solar wing respectively, calculate the transfer matrix of the south solar wing installation coordinate system relative to the whole satellite mechanical coordinate system and the transfer matrix of the north solar wing installation coordinate system relative to the whole satellite mechanical coordinate system.

[0040] In this step, the transfer matrix C of the south solar wing installation coordinate system relative to the overall satellite mechanical coordinate system is... SSJ The specific calculation formula is as follows:

[0041]

[0042] In the formula, α SS and β SS These are the A-axis rotation angle and B-axis rotation angle of the south solar array, respectively.

[0043] The transition matrix C between the North Solar Panel installation coordinate system and the overall satellite mechanical coordinate system NSJ The specific calculation formula is as follows:

[0044]

[0045] In the formula, α NS and β NS These are the A-axis rotation angle and B-axis rotation angle of the north solar array, respectively.

[0046] S2, for both the south and north solar arrays, the rotational inertia of the solar array relative to the satellite's center of mass is calculated based on the transfer matrix of the solar array's installation coordinate system relative to the overall satellite mechanical coordinate system, the mass of the solar array, the moment of inertia of the solar array's mass relative to its center of mass, the position of the origin of the solar array's installation coordinate system in the overall satellite mechanical coordinate system, the position of the solar array's center of mass in the solar array's installation coordinate system, and the position of the satellite's center of mass in the overall satellite mechanical coordinate system.

[0047] In this step, the moment of inertia of the south solar array relative to the satellite's center of mass... The specific calculation formula is as follows:

[0048]

[0049] In the formula, C SSJ The transfer matrix for the south solar array's installation coordinate system relative to the entire satellite's mechanical coordinate system, m SA For the mass of the south solar wing, Let be the moment of inertia of the south solar wing relative to its center of mass. The position of the origin of the coordinate system for the south solar array within the overall satellite's mechanical coordinate system. This represents the position of the south solar wing's center of mass in the south solar wing's mounting coordinate system. Let the position of the satellite's center of mass in the whole satellite's mechanical coordinate system be (). × This refers to the vector cross product operation between matrices.

[0050] Moment of inertia of the north solar array relative to the satellite's center of mass The specific calculation formula is as follows:

[0051]

[0052] In the formula, C NSJ The transfer matrix for the north solar array's installation coordinate system relative to the entire satellite's mechanical coordinate system, m NA For the mass of the north solar array, Let be the moment of inertia of the north solar array relative to its center of mass. The position of the origin of the coordinate system for the North Solar Arm within the overall satellite's mechanical coordinate system. This represents the position of the North Solar Array's center of mass in the North Solar Array's mounting coordinate system. This represents the position of the satellite's center of mass in the overall satellite mechanical coordinate system.

[0053] S3. Based on the rotational inertia of the north and south solar arrays relative to the satellite's center of mass and the rotational inertia of the satellite's central body relative to its center of mass, calculate the satellite's rotational inertia relative to its center of mass.

[0054] In this step, the satellite's moment of inertia relative to its center of mass... The specific calculation formula is as follows:

[0055]

[0056] In the formula, Let be the moment of inertia of the south solar array relative to the satellite's center of mass. Let be the moment of inertia of the north solar array relative to the satellite's center of mass. Let be the moment of inertia of the satellite's central body relative to its center of mass.

[0057] S4. For both the south and north solar arrays, the rotational coupling coefficient of the solar array relative to the center of mass in the body coordinate system is calculated based on the transfer matrix of the solar array installation coordinate system relative to the overall satellite mechanical coordinate system, the translational and rotational coupling coefficients of the solar array relative to the connection point in the solar array installation coordinate system, the position of the origin of the solar array installation coordinate system in the overall satellite mechanical coordinate system, and the position of the satellite's center of mass in the overall satellite mechanical coordinate system.

[0058] In this step, the rotational coupling coefficient of the south solar array relative to the center of mass in the body coordinate system is... The specific calculation formula is as follows:

[0059]

[0060] In the formula, C SSJ The installation matrix for the south solar array's coordinate system relative to the overall satellite's mechanical coordinate system. and These represent the translational and rotational coupling coefficients of the south solar array relative to the connection point in the south solar array installation coordinate system, respectively. The position of the origin of the coordinate system for the south solar array within the overall satellite's mechanical coordinate system. This represents the position of the satellite's center of mass in the overall satellite mechanical coordinate system.

[0061] In this step, the rotational coupling coefficient of the north solar array relative to the center of mass in the body coordinate system is... The calculation formula is as follows:

[0062]

[0063] In the formula, C NSJ The installation matrix for the north solar array's coordinate system relative to the overall satellite's mechanical coordinate system. and These represent the translational and rotational coupling coefficients of the north solar array relative to the connection point in the north solar array installation coordinate system. The position of the origin of the coordinate system for the North Solar Arm within the overall satellite's mechanical coordinate system. This represents the position of the satellite's center of mass in the overall satellite mechanical coordinate system.

[0064] S5. Based on the satellite's rotational inertia relative to its center of mass, the rotational coupling coefficients of the north and south solar arrays relative to the center of mass in the body coordinate system, the frequency matrices of the north and south solar arrays, and the damping matrices of the north and south solar arrays, the system dynamic equations are constructed.

[0065] In some implementations, step S5 includes:

[0066]

[0067] Y Dyn =C Dyn X Dyn +D Dyn U Dyn

[0068] Among them, X Dyn The system state vector can be represented as: θ represents the roll, pitch, and yaw attitude angles. For the roll, pitch, and yaw attitude angular rates, q S The coordinates of the south solar wing mode are as follows: Let q be the first-order rate of change of the modal coordinates of the south solar wing. N The coordinates for the northern solar wing mode are as follows: The first-order rate of change of the modal coordinates of the north solar wing. U is the derivative of the system state vector; Dyn The attitude control torque acting on the celestial body is the system input vector; A Dyn B Dyn C Dyn and D Dyn The system's first matrix, second matrix, third matrix, and fourth matrix are respectively represented as:

[0069]

[0070] In the formula, n S Let n be the order of the flexible mode of the south solar wing. N Let I be the order of the flexible mode of the north solar array, 0 be the identity matrix, and Y be the zero matrix. Dyn1 Y Dyn2 and Y Dyn3 The intermediate matrix has no real meaning and is represented as follows:

[0071]

[0072]

[0073] In the formula, Λ S The frequency matrix of the south solar wing, Λ N Here is the frequency matrix of the north solar wing, ξ N Let ξ be the damping matrix of the north solar wing. S The damping matrix of the south solar panel. Let be the moment of inertia of the satellite relative to its center of mass. and These are the rotational coupling coefficients of the south solar array relative to the center of mass in the body coordinate system and the rotational coupling coefficients of the north solar array relative to the center of mass in the body coordinate system, respectively.

[0074] S6, based on the system dynamic equations, determines the first transfer function matrix of the satellite flexible dynamic system from the three-axis attitude control torque to the three-axis angle and the three-axis angular velocity.

[0075] In some implementations, step S6 may include:

[0076]

[0077] Among them, G Dyn (s) is the first transfer function matrix, representing the transfer function of the satellite flexible dynamics system from the attitude control torques of the three axes of roll, pitch, and yaw to the angles and angular velocities of the three axes of roll, pitch, and yaw; s is a complex variable, A Dyn B DynC Dyn and D Dyn These are the system's first matrix, second matrix, third matrix, and fourth matrix, respectively, where I is the identity matrix, and n... S Let n be the order of the flexible mode of the south solar wing. N The order of the flexible mode of the North Solar Array.

[0078] Regarding step 102:

[0079] In this embodiment, the angle measurement transfer function used to characterize the frequency domain characteristics of the sensor is calculated using the following formula:

[0080]

[0081] Among them, G Mea (s) is the angle measurement transfer function, K M τ is the gain coefficient of the sensor's frequency domain characteristics. M0 τ M1 τ M2 Let be the time constant of the frequency domain characteristics of the sensor, and s be a complex variable.

[0082] The specific formula for calculating the second transfer function of the satellite controller along the three-axis channel, which is used to characterize the frequency domain characteristics of the controller, is as follows:

[0083]

[0084] Among them, G Ctrlx (s), G Ctrly (s), G Ctrlz (s) represent the second transfer functions of the satellite controller along the roll, pitch, and yaw channels, respectively, and K... Px K Dx K Ix For the proportional, derivative, and integral control parameters of the satellite controller's rolling channel; K Py K Dy K Iy For the proportional, derivative, and integral control parameters of the satellite controller's pitch channel; K Pz K Dz K Iz These are the proportional, derivative, and integral control parameters for the yaw channel of the satellite controller.

[0085] Due to the large disturbance torque conditions during satellite ignition at apogee, the ignition control is approximately a phase-leading element. The ignition control transfer function G, used to characterize the frequency domain characteristics of the actuator, is... Thr (s) can be expressed as:

[0086]

[0087] In the formula, KT For the gain of ignition control, T T0 For the thruster output torque delay; T T1 The ignition control is approximated by the time constant of the phase lead element, where s is a complex variable.

[0088] Regarding step 104:

[0089] In some implementations, the transfer function for satellite three-axis attitude control is calculated using the following formula:

[0090] G x (s)=G Dyn1,1 (s)G Thr (s)G Ctrlx (s)G Mea (s)

[0091] G y (s)=G Dyn2,2 (s)G Thr (s)G Ctrly (s)G Mea (s)

[0092] G z (s)=G Dyn3,3 (s)G Thr (s)G Ctrlz (s)G Mea (s)

[0093] In the formula, G x (s), G y (s) and G z (s) represent the transfer functions for attitude control of the satellite in the roll, pitch, and yaw channels, respectively, and G Dyn1,1 (s), G Dyn2,2 (s), G Dyn3,3 (s) is the element in the i-th row and j-th column of the first transfer function matrix, G Dyn1,1 (s) is the transfer function from the rolling channel attitude control torque to the angle, G Dyn2,2 (s) is the transfer function from pitch channel attitude control torque to angle, G Dyn3,3 (s) is the transfer function from the yaw channel attitude control torque to the angle, G Thr (s) is the ignition control transfer function, G Ctrlx (s), G Ctrly (s), G Ctrlz (s) represent the second transfer functions of the satellite controller along the roll, pitch, and yaw channels, respectively, G Mea (s) is the angle measurement transfer function.

[0094] Regarding step 106:

[0095] Based on the transfer function G of the satellite's three-axis attitude control x (s), G y (s), G z (s) Plot the three-axis Nichols plots respectively. The control system should meet the stability design requirement that the phase margin of each of the three-axis channels is greater than 15°. If it meets the requirement, it means that the stability design of the satellite controller is up to standard; if it does not meet the requirement, adjust the control parameters of the satellite controller until the phase margin of each of the three-axis channels is greater than 15°.

[0096] The experimental verification results of this scheme will be described below. Figure 2 The image shows the Nichols plot of the rolling channel of the satellite attitude control system under the condition that the A-axis rotation angle is 45° and the B-axis rotation angle is 30°. The phase stability margin is 20.7° (the phase when the curve first crosses the X-axis is -159.3°). Figure 3 The image shows the Nichols plot of the pitch channel of the satellite attitude control system under the condition that the A-axis rotation angle is 45° and the B-axis rotation angle is 30°. The phase stability margin is 43.1° (the phase when the curve first crosses the X-axis is -136.9°). Figure 4 The Nichols plot for the yaw channel of the satellite attitude control system is shown for a dual-axis solar array with an A-axis rotation angle of 45° and a B-axis rotation angle of 30°. The phase stability margin is 35.6° (the phase when the curve first crosses the X-axis is -144.4°). The calculation results indicate that the phase stability margins of the roll, pitch, and yaw channels of the control system are greater than 15°, meeting the design requirements. Therefore, based on the rotation angles of the south and north solar arrays, as well as the installation positions, masses, and moments of inertia of the north and south solar arrays, this scheme calculates the first transfer function matrix to characterize the flexible dynamic frequency domain characteristics of the dual-axis solar array. This further considers the flexible dynamic characteristics of the satellite using a dual-axis solar array, which can improve the stability of the satellite's attitude control system during apogee ignition.

[0097] Please refer to Figure 5 This invention provides a satellite stability control analysis device for arbitrary rotation angle of a solar panel under large disturbance torque, used to implement the steps of any method embodiment in the specification. The device includes:

[0098] The flexible dynamics frequency domain characteristic unit 501 is used to calculate the first transfer function matrix of the satellite flexible dynamics system from the three-axis attitude control torque to the three-axis angle and three-axis angular velocity based on the rotation angle of the south solar array dual-axis and the north solar array dual-axis, as well as the installation position, mass and rotational inertia of the north and south solar arrays; the first transfer function matrix is ​​used to characterize the flexible dynamics frequency domain characteristics of the dual-axis solar array;

[0099] The calculation unit 502 is used to calculate the angle measurement transfer function for characterizing the frequency domain characteristics of the sensor, the second transfer function of the satellite controller along the three-axis channel for characterizing the frequency domain characteristics of the controller, and the ignition control transfer function for characterizing the frequency domain characteristics of the actuator, respectively.

[0100] The attitude calculation unit 503 is used to calculate the transfer function of the satellite's three-axis attitude control based on the first transfer function matrix, the angle measurement transfer function, the second transfer function, and the ignition control transfer function.

[0101] The stability analysis unit 504 is used to plot Nichols diagrams based on the transfer functions of the satellite's three-axis attitude control to determine whether the phase margin of the three axes meets the stability design requirements. If not, the control parameters of the satellite controller are adjusted.

[0102] It should be noted that the satellite stabilization control and analysis device for arbitrary rotation angle of the solar panel under large disturbance torque provided in the above embodiments is only an example of the division of the above functional units. In practical applications, the above functions can be assigned to different functional units as needed, that is, the internal structure of the device can be divided into different functional units to complete all or part of the functions described above. In addition, the above device embodiments and method embodiments belong to the same concept, and the specific implementation process can be found in the method embodiments, which will not be repeated here.

[0103] Embodiments of this application also provide a computer device, please refer to... Figure 6 The computer device includes a processor and a memory, the memory storing at least one instruction, at least one program, code set or instruction set, the at least one instruction, at least one program, code set or instruction set being loaded and executed by the processor to implement the satellite stability control analysis method for arbitrary angle rotation of the solar panel under large disturbance torque provided in the above-described method embodiments.

[0104] The embodiments of this application also provide a computer-readable storage medium storing at least one instruction, at least one program, code set, or instruction set, wherein the at least one instruction, at least one program, code set, or instruction set is loaded and executed by a processor to implement the satellite stability control analysis method for arbitrary rotation angle of solar panels under large disturbance torque provided in the above-described method embodiments.

[0105] Embodiments of this application also provide a computer program product, which includes a computer program. A processor of a computer device reads the computer program from a computer-readable storage medium and executes the computer program, causing the computer device to perform any of the satellite stability control analysis methods for arbitrary angle rotation of the solar panel under large disturbance torque in the above embodiments.

[0106] For ease of description, the above systems or devices are described separately as various modules or units based on their functions. Of course, in implementing this application, the functions of each unit can be implemented in one or more software and / or hardware components.

[0107] As can be seen from the above description of the embodiments, those skilled in the art can clearly understand that this application can be implemented by means of software plus necessary general-purpose hardware platforms. Based on this understanding, the technical solution of this application, in essence or the part that contributes to the prior art, can be embodied in the form of a software product. This computer software product can be stored in a storage medium, such as ROM / RAM, magnetic disk, optical disk, etc., and includes several instructions to cause a computer device (which may be a personal computer, server, or network device, etc.) to execute the methods of various embodiments or some parts of the embodiments of this application.

[0108] Finally, it should be noted that in this document, relational terms such as first, second, third, and fourth are used merely to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes the element.

[0109] The above are merely preferred embodiments of this application. It should be noted that those skilled in the art can make various improvements and modifications without departing from the principles of this application, and these improvements and modifications should also be considered within the scope of protection of this application.

Claims

1. A satellite stability control analysis method under large disturbance moment with arbitrary rotation angle of solar panels, characterized in that, include: Based on the rotation angles of the south and north solar arrays, as well as the installation positions, masses, and moments of inertia of the north and south solar arrays, the first transfer function matrix of the satellite's flexible dynamics system is calculated from the attitude control torques along the three axes to the angles and angular velocities along the three axes. The first transfer function matrix is ​​used to characterize the frequency domain characteristics of the flexible dynamics of the dual-axis solar arrays. Calculate the angle measurement transfer function to characterize the frequency domain characteristics of the sensor, the second transfer function of the satellite controller along the three-axis channel to characterize the frequency domain characteristics of the controller, and the ignition control transfer function to characterize the frequency domain characteristics of the actuator, respectively. Based on the first transfer function matrix, the angle measurement transfer function, the second transfer function, and the ignition control transfer function, calculate the transfer function for the satellite's three-axis attitude control; Nichols plots were drawn based on the transfer functions of the satellite's three-axis attitude control to determine whether the phase margin of the three axes met the stability design requirements. If not, the control parameters of the satellite controller were adjusted.

2. The method as described in claim 1, characterized in that, The calculation of the first transfer function matrix of the satellite's flexible dynamics system, based on the rotation angles of the south and north solar arrays and the installation positions, masses, and moments of inertia of the north and south solar arrays, along with the triaxial torques to triaxial angles and angular velocities, includes: Based on the rotation angles of the A and B axes of the south solar array and the north solar array, respectively, calculate the transfer matrix of the south solar array installation coordinate system relative to the overall satellite mechanical coordinate system and the transfer matrix of the north solar array installation coordinate system relative to the overall satellite mechanical coordinate system. For both the south and north solar arrays, the moment of inertia of the solar array relative to the satellite's center of mass is calculated based on the transfer matrix of the solar array's installation coordinate system relative to the overall satellite mechanical coordinate system, the mass of the solar array, the moment of inertia of the solar array's mass relative to its center of mass, the position of the origin of the solar array's installation coordinate system in the overall satellite mechanical coordinate system, the position of the solar array's center of mass in the solar array's installation coordinate system, and the position of the satellite's center of mass in the overall satellite mechanical coordinate system. The moment of inertia of the satellite relative to its center of mass is calculated based on the moment of inertia of the north and south solar arrays relative to the satellite's center of mass and the moment of inertia of the satellite's central body relative to its center of mass. For both the south and north solar arrays, the rotational coupling coefficient of the solar array relative to the center of mass in the body coordinate system is calculated based on the transfer matrix of the solar array's installation coordinate system relative to the overall satellite mechanical coordinate system, the translational and rotational coupling coefficients of the solar array relative to the connection point in the solar array's installation coordinate system, the position of the origin of the solar array's installation coordinate system in the overall satellite mechanical coordinate system, and the position of the satellite's center of mass in the overall satellite mechanical coordinate system. Based on the satellite's rotational inertia relative to its center of mass, the rotational coupling coefficients of the north and south solar arrays relative to the center of mass in the body coordinate system, the frequency matrices of the north and south solar arrays, and the damping matrices of the north and south solar arrays, the system dynamic equations are constructed. Based on the system dynamic equations, the first transfer function matrix of the satellite flexible dynamic system along the three-axis attitude control torque to the three-axis angle and three-axis angular velocity is determined.

3. The method as described in claim 2, characterized in that, The system dynamic equations are constructed based on the satellite's rotational inertia relative to its center of mass, the rotational coupling coefficients of the north and south solar arrays relative to the center of mass in the body coordinate system, the frequency matrices of the north and south solar arrays, and the damping matrices of the north and south solar arrays, including: Y Dyn =C Dyn X Dyn +D Dyn U Dyn Among them, X Dyn The system state vector can be represented as θ represents the roll, pitch, and yaw attitude angles. For the roll, pitch, and yaw attitude angular rates, q S The coordinates of the south solar wing mode are as follows: Let q be the first-order rate of change of the modal coordinates of the south solar wing. N The coordinates of the north solar wing mode are as follows: The first-order rate of change of the modal coordinates of the north solar wing. U is the derivative of the system state vector; Dyn The attitude control torque acting on the celestial body is the system input vector; A Dyn B Dyn C Dyn and D Dyn The system's first matrix, second matrix, third matrix, and fourth matrix are respectively represented as: In the formula, n S Let n be the order of the flexible mode of the south solar wing. N Let I be the order of the flexible mode of the north solar array, 0 be the identity matrix, and Y be the zero matrix. Dyn1 Y Dyn2 and Y Dyn3 The intermediate matrix has no real meaning and is represented as follows: In the formula, Λ S The frequency matrix of the south solar wing, Λ N Here is the frequency matrix of the north solar wing, ξ N Let ξ be the damping matrix of the north solar wing. S The damping matrix of the south solar panel. Let be the moment of inertia of the satellite relative to its center of mass. and These are the rotational coupling coefficients of the south solar array relative to the center of mass in the body coordinate system and the rotational coupling coefficients of the north solar array relative to the center of mass in the body coordinate system, respectively.

4. The method as described in claim 3, characterized in that, The determination of the first transfer function matrix of the satellite flexible dynamics system from the three-axis attitude control torque to the three-axis angles and three-axis angular velocities based on the system dynamic equations includes: Among them, G Dyn (s) is the first transfer function matrix, representing the transfer function of the satellite flexible dynamics system from the attitude control torques of the three axes of roll, pitch, and yaw to the angles and angular velocities of the three axes of roll, pitch, and yaw; s is a complex variable, A Dyn B Dyn C Dyn and D Dyn These are the system's first matrix, second matrix, third matrix, and fourth matrix, respectively, where I is the identity matrix, and n... S Let n be the order of the flexible mode of the south solar wing. N The order of the flexible mode of the North Solar Array.

5. The method as described in claim 1, characterized in that, The angle measurement transfer function is calculated using the following formula: Among them, G Mea (s) is the angle measurement transfer function, K M τ is the gain coefficient of the sensor's frequency domain characteristics. M0 τ M1 τ M2 Let be the time constant of the frequency domain characteristics of the sensor, and s be a complex variable.

6. The method as described in claim 1, characterized in that, The transfer function for satellite three-axis attitude control is calculated using the following formula: G x (s)=G Dyn1,1 (s)G Thr (s)G Ctrlx (s)G Mea (s) G y (s)=G Dyn2,2 (s)G Thr (s)G Ctrly (s)G Mea (s) G z (s)=G Dyn3,3 (s)G Thr (s)G Ctrlz (s)G Mea (s) In the formula, G x (s), G y (s) and G z (s) represent the transfer functions for attitude control of the satellite in the roll, pitch, and yaw channels, respectively, and G Dyn1,1 (s), G Dyn2,2 (s), G Dyn3,3 (s) is the element in the i-th row and j-th column of the first transfer function matrix, G Dyn1,1 (s) is the transfer function from the rolling channel attitude control torque to the angle, G Dyn2,2 (s) is the transfer function from pitch channel attitude control torque to angle, G Dyn3,3 (s) is the transfer function from the yaw channel attitude control torque to the angle, G Thr (s) is the ignition control transfer function, G Ctrlx (s), G Ctrly (s), G Ctrlz (s) represent the second transfer functions of the satellite controller along the roll, pitch, and yaw channels, respectively, G Mea (s) is the angle measurement transfer function.

7. A satellite stability control and analysis device for arbitrary rotation angle of solar panels under large disturbance torque, used to implement the steps of the method described in any one of claims 1-6, characterized in that, include: The flexible dynamics frequency domain characteristic unit is used to calculate the first transfer function matrix of the satellite flexible dynamics system from the three-axis attitude control torque to the three-axis angle and three-axis angular velocity based on the rotation angle of the south solar array dual axis and the north solar array dual axis, as well as the installation position, mass and rotational inertia of the north and south solar arrays. The first transfer function matrix is ​​used to characterize the flexible dynamic frequency domain characteristics of the dual-axis solar array; The calculation unit is used to calculate the angle measurement transfer function, which characterizes the frequency domain characteristics of the sensor; the second transfer function of the satellite controller along the three-axis channel, which characterizes the frequency domain characteristics of the controller; and the ignition control transfer function, which characterizes the frequency domain characteristics of the actuator. The attitude calculation unit is used to calculate the transfer function of the satellite's three-axis attitude control based on the first transfer function matrix, the angle measurement transfer function, the second transfer function, and the ignition control transfer function. The stability analysis unit is used to plot Nichols diagrams based on the transfer functions of the satellite's three-axis attitude control to determine whether the phase margin of the three axes meets the stability design requirements. If not, the control parameters of the satellite controller are adjusted.

8. A computer device, characterized in that, The computer device includes a memory and a processor. The memory is used to store computer programs, and the processor is used to execute the computer programs stored in the memory to implement the steps of the method according to any one of claims 1-6.

9. A computer-readable storage medium, characterized in that, The storage medium stores a computer program, which, when executed by a processor, implements the steps of the method described in any one of claims 1-6.

10. A computer program product, characterized in that, Includes a computer program, which, when executed by a processor, implements the steps of the method according to any one of claims 1-6.

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