On-orbit identification method and system for testing mass bias force of gravitational wave spacecraft
By constructing a relative motion dynamics model and applying periodic excitation signals, combined with spacecraft attitude control system data, the problem of on-orbit bias force identification was solved, achieving high-precision bias force monitoring and identification, and supporting spacecraft state assessment and compensation strategies.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- INNOVATION ACAD FOR MICROSATELLITES OF CAS
- Filing Date
- 2026-05-21
- Publication Date
- 2026-07-31
AI Technical Summary
Existing technologies lack methods for directly measuring and verifying mass bias forces in orbit. Ground models cannot accurately reflect the actual bias forces in orbit, and the bias forces are difficult to separate from disturbance noise coupling, affecting the sensitivity of gravitational wave detection.
A relative motion dynamic model for quality inspection was constructed, a sun-pointing measurement condition was designed, a periodic excitation signal was applied, and the equations were processed by the least squares method to identify the bias force, in conjunction with the spacecraft attitude control system data.
It achieves accurate on-orbit identification of bias force, improves the reliability and accuracy of identification results, and enables monitoring of bias force changes throughout the spacecraft's entire life cycle, supporting state assessment and compensation strategy optimization.
Smart Images
Figure CN122490833A_ABST
Abstract
Description
Technical Field
[0001] This invention belongs to the field of space gravitational wave detection technology, specifically relating to an on-orbit identification method and system for the test mass bias force of a gravitational wave spacecraft, and also relating to a corresponding on-board computer and computer-readable storage medium. Background Technology
[0002] The core principle of space-based gravitational wave detection is to establish a high-precision laser interferometric link between two or more spacecraft (such as satellites) and measure the minute changes in the distance between the test masses of the two spacecraft caused by gravitational waves. To achieve this, each spacecraft has a test mass suspended inside as an inertial reference, and a drag-free control system controls the spacecraft platform to always follow the movement of the test mass, thereby suppressing the influence of external non-conservative forces on the test mass. However, even under ideal drag-free control, the test mass itself is still subject to a perturbation force generated by multiple physical fields, including the spacecraft's own gravitational field, electric field, and magnetic field. This perturbation is a major factor affecting drag-free control and the sensitivity of gravitational wave detection. The main components of this perturbation include the bias force experienced by the test mass at the center of the electrode cage and the stiffness generated by the motion of the test mass relative to the electrode cage. According to the requirements of space-based gravitational wave detection, the bias force perturbation generally does not exceed 10. -9 m / s 2 Accurately identifying the bias forces acting on the spacecraft's in-orbit mass is crucial for assessing the spacecraft's self-gravity design level, validating ground-based calculation models, guiding the development of in-orbit compensation strategies, and monitoring changes in the spacecraft's in-orbit status.
[0003] Existing technologies for identifying bias forces mainly suffer from the following problems: First, there is a lack of direct on-orbit measurement methods. The test mass is suspended in a vacuum environment and has no physical contact with the spacecraft, making it impossible to install traditional force sensors. There are no instruments or equipment that can directly measure the bias force experienced by the test mass. Second, indirect identification methods rely on models pre-established during the design of ground-based spacecraft, which cannot accurately reflect the true value of the actual bias force in orbit. They also struggle to capture changes in bias force caused by factors such as propellant consumption and thermal deformation. Furthermore, the bias force is coupled with various disturbances and noises, making effective separation difficult. Third, ground-based measurement methods are greatly affected by external interference such as gravity, making it difficult to obtain the true force situation experienced by the test mass in a free-floating state. Summary of the Invention
[0004] To address the aforementioned shortcomings in the prior art, this invention provides an on-orbit identification method and system for the mass bias force of a gravitational wave spacecraft, along with a corresponding onboard computer and computer-readable storage medium.
[0005] According to a first aspect of the present invention, an on-orbit identification method for a gravitational wave spacecraft's test mass bias force is provided, comprising: Construct a relative motion dynamic model for the inspection quality; The design of the solar pointing measurement condition is such that the solar radiation pressure, which is the main non-conservative force disturbance of the spacecraft, is along the Z-axis of the spacecraft. Under the aforementioned sun-pointing measurement condition, a periodic excitation signal is applied to the relative motion dynamics model of the test quality, causing the spacecraft to generate a small AC component in the orthogonal direction; Under the action of the excitation signal, based on the attitude dynamics parameter data in the spacecraft attitude control system and combined with the relative motion dynamics model of the test mass, a set of equations containing the parameters to be identified is constructed. Data processing is performed on the set of equations containing the parameters to be identified to obtain the optimal solution for the bias force of the test mass of the gravitational wave spacecraft, thereby realizing the identification of the bias force.
[0006] Preferably, the construction of the relative motion dynamics model for the inspection quality includes: Starting from the displacement of the inspection mass relative to the electrode cage, a relative motion dynamic model of the inspection mass is constructed for each inspection mass i inside the spacecraft, expressed as: , In the formula, The displacement vector representing the inspection quality i relative to the nominal position (usually the center of the electrode cage). Take the second derivative with respect to time; This represents the position vector from the spacecraft's center of mass to the center of the test mass i; Represents the angular velocity of a spacecraft; This represents the force per unit mass acting on the inspection quality i, including the stiffness effect, electrostatic actuation force, and bias force acting on the inspection quality. This refers to the force per unit mass acting on the spacecraft, including solar radiation pressure and micro-thrust acting on the spacecraft. , Representing parameters respectively , Find the first derivative with respect to time.
[0007] Preferably, the equation set containing the parameters to be identified is constructed based on the attitude dynamics parameter data in the spacecraft attitude control system and combined with the relative motion dynamics model of the test mass, including: By combining the onboard sensors and control loops in the spacecraft's attitude control system, attitude dynamics parameter data can be obtained; Based on the aforementioned attitude dynamics parameter data, the relative motion dynamics model of the inspection quality is approximated to obtain the parametric model of the inspection quality, expressed as: , In the formula, This represents the force per unit mass acting on the spacecraft. The effect of solar radiation pressure, This represents the force per unit mass of the inspected mass i. The electrostatic actuation effect in This represents the force per unit mass of the inspected mass i. The biasing force exerted by the spacecraft on the quality of the inspection; Based on the characteristics of the applied excitation signal, when the spacecraft oscillates at an angle less than a set threshold around any orthogonal axis of the Z-axis, the direction of that orthogonal axis is denoted as... The magnitude of the force exerted by sunlight on the spacecraft along its three axes is denoted as: , remember: ,because An angle less than a set threshold, i.e. Therefore, we can conclude that: , In the formula, , Since all parameters on the right side of the equation are known quantities, the right side of the equation can be simplified as: , In the formula, The x, y, and z coordinate components are represented by the sum of the terms on the right. After the spacecraft undergoes multiple cycles of motion, it obtains N sets of sampled data. Let the j-th set of sampled data be denoted as: , This yields N sets of equations containing parameters to be identified; among them, the j-th set of equations containing parameters to be identified is expressed as: , In the formula, Indicates the spacecraft's oscillation angle The j-th data set; The x-axis component of the bias force exerted by the spacecraft on the test quality; This represents the magnitude of the light pressure force when sunlight shines directly along the z-axis of the spacecraft. The y-axis component of the bias force exerted by the spacecraft on the test quality; This represents the z-axis component of the bias force exerted by the spacecraft on the test quality. express The j-th sampled data.
[0008] Preferably, the step of processing the equations containing the parameters to be identified to obtain the optimal solution for the test mass bias force of the gravitational wave spacecraft includes: The least squares method is used to process the data of the system of equations containing the parameters to be identified; wherein: remember: , , Therefore, the least squares solution to the system of equations can be obtained as follows: , This enables the gravitational wave spacecraft to achieve the test mass bias force. The identification, and simultaneously the identification of the magnitude of the light pressure experienced by the spacecraft when its normal direction is pointing towards the sun. .
[0009] According to a second aspect of the present invention, an on-orbit identification system for a gravitational wave spacecraft to test mass bias force is provided, comprising: Inspection quality model building module, which is used to build a relative motion dynamics model of inspection quality; Measurement Condition Design Module: This module is used to design the solar pointing measurement conditions so that the solar radiation pressure, which is the main non-conservative force disturbance of the spacecraft, is along the Z-axis of the spacecraft. An excitation signal application module is used to apply a periodic excitation signal to the relative motion dynamics model of the test quality under the sun pointing measurement condition, so that the spacecraft generates a small AC component in the orthogonal direction. The module for constructing parameters to be identified is used to construct a set of equations containing the parameters to be identified based on the attitude dynamics parameter data in the spacecraft attitude control system and the relative motion dynamics model of the test mass under the action of the excitation signal. The bias force identification module is used to process the equations containing the parameters to be identified to obtain the optimal solution of the bias force of the gravitational wave spacecraft for the test mass, thereby realizing the identification of the bias force.
[0010] According to a third aspect of the present invention, an on-board computer is provided, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor, when executing the computer program, can be used to perform the method described in any one of the above inventions.
[0011] According to a fourth aspect of the present invention, a computer-readable storage medium is provided having a computer program stored thereon, which, when executed by a processor, can be used to perform the method described in any one of the preceding inventions.
[0012] By adopting the above technical solution, the present invention has at least one of the following beneficial effects compared with the prior art: This invention achieves the identification of the bias force of a gravitational wave spacecraft on the quality of the test based on on-orbit measured data. It does not rely on a pre-established finite element model on the ground and can truly reflect the actual state of the spacecraft in orbit, avoiding errors caused by the difference between the design model and the actual object.
[0013] This invention, through a special design for solar pointing measurement, effectively separates the coupling between bias force and external disturbances such as solar radiation pressure, thereby improving the reliability and accuracy of the identification results.
[0014] This invention can be repeatedly executed at different stages of the spacecraft's entire life cycle (such as the initial orbital insertion, after propellant consumption, and after thermal deformation stabilization), enabling on-orbit monitoring of the long-term evolution of the bias force, and providing a basis for spacecraft condition assessment, model correction, and compensation strategy optimization. Attached Figure Description
[0015] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings: Figure 1 This is a flowchart illustrating the on-orbit identification method for the test mass bias force of a gravitational wave spacecraft in a preferred embodiment of the present invention.
[0016] Figure 2 This is a schematic diagram of the components of the on-orbit identification system for the test mass bias force of a gravitational wave spacecraft in a preferred embodiment of the present invention.
[0017] Figure 3 This is a schematic diagram of the placement configuration of two inspection masses inside a spacecraft in a specific application example of the present invention. Figure 4 This is a schematic diagram of a spacecraft configuration in a specific application example of the present invention.
[0018] Figure 5 This is a flowchart illustrating the on-orbit identification method for the test mass bias force of a gravitational wave spacecraft in a specific application example of the present invention. Detailed Implementation
[0019] The embodiments of the present invention are described in detail below: These embodiments are implemented based on the technical solution of the present invention, and provide detailed implementation methods and specific operation processes. It should be noted that those skilled in the art can make several modifications and improvements without departing from the concept of the present invention, and these all fall within the protection scope of the present invention.
[0020] The bias force disturbance experienced by the test mass has specific requirements in space gravitational wave detection. Accurately identifying the bias force experienced by the test mass in orbit is of great significance for spacecraft evaluation. Current technologies for identifying bias forces typically suffer from technical problems such as a lack of direct on-orbit measurement methods, strong limitations of indirect on-orbit identification methods, and the inability of ground measurements to reflect the actual on-orbit condition.
[0021] To address the aforementioned problems, one embodiment of the present invention provides an on-orbit identification method for the bias force on the test mass of a gravitational wave spacecraft. This method is used on a gravitational wave spacecraft to identify the bias force on the test mass on-orbit, enabling accurate identification of the bias force on the test mass and providing reliable data support for bias force control and compensation.
[0022] Specifically, such as Figure 1 As shown in this embodiment, the on-orbit identification method for the test mass bias force of a gravitational wave spacecraft may include: S1, Construct a relative motion dynamic model for the quality inspection; S2, designed for sun-pointing measurement, ensures that the solar radiation pressure, which is the main non-conservative force disturbance of the spacecraft, is along the spacecraft's Z-axis direction; S3, under the condition of sun pointing measurement, applies a periodic excitation signal to the relative motion dynamics model of the quality inspection, so that the spacecraft generates a small AC component in the orthogonal direction; S4. Under the action of the excitation signal, based on the attitude dynamics parameter data in the spacecraft attitude control system and combined with the relative motion dynamics model of the test mass, a set of equations containing the parameters to be identified is constructed. S5 processes the equations containing the parameters to be identified to obtain the optimal solution for the bias force of the test mass of the gravitational wave spacecraft, thereby realizing the identification of the bias force.
[0023] Wherein: S4 above, the j-th group of equations containing the parameters to be identified, is expressed as: In the formula, Indicates the spacecraft's oscillation angle The j-th data set; The x-axis component of the bias force exerted by the spacecraft on the test quality; This represents the magnitude of the light pressure force when sunlight shines directly along the z-axis of the spacecraft. The y-axis component of the bias force exerted by the spacecraft on the test quality; This represents the z-axis component of the bias force exerted by the spacecraft on the test quality. express The j-th sampled data.
[0024] In some preferred embodiments, S1 above, which involves constructing a relative motion dynamic model of the inspection quality, may further include: Starting from the displacement of the inspection mass relative to the electrode cage, a relative motion dynamic model of the inspection mass is constructed for each inspection mass i inside the spacecraft, expressed as: In the formula, The displacement vector representing the inspection quality i relative to the nominal position (usually the center of the electrode cage). Take the second derivative with respect to time; This represents the position vector from the spacecraft's center of mass to the center of the test mass i; Represents the angular velocity of a spacecraft; This represents the force per unit mass acting on the inspection quality i, including the stiffness effect, electrostatic actuation force, and bias force acting on the inspection quality. This refers to the force per unit mass acting on the spacecraft, including solar radiation pressure and micro-thrust acting on the spacecraft. , Representing parameters respectively , Find the first derivative with respect to time.
[0025] In some preferred embodiments, the above-mentioned S2, which designs the solar pointing measurement condition such that the solar radiation pressure, which is the main non-conservative force disturbance of the spacecraft, is along the spacecraft's Z-axis direction, may further include: Before identifying the bias force of a gravitational wave spacecraft on the test mass, the measurement conditions need to be designed. This involves using micro-thrusters to control the spacecraft's attitude only, ensuring the spacecraft's Z-axis is aligned with the sun and the Y-axis is along the orbital plane normal. Depending on the spacecraft configuration, such as... Figure 4 As shown, the solar radiation pressure is along the -Z axis at this time.
[0026] In some preferred embodiments, S3, applying a periodic excitation signal to cause the spacecraft to generate a small AC component in the orthogonal direction, may further include: When measuring the bias force of a gravitational wave spacecraft on a test mass, a periodic excitation signal is added to the relative motion dynamics model of the test mass using a micro-thruster. This causes the spacecraft to oscillate at a small angle around any orthogonal axis of the Z-axis (e.g., a sinusoidal periodic motion around the Y-axis, with a peak oscillation angle not exceeding 0.1° and an oscillation period of 50s). The oscillation angle is denoted as... At this time, the spacecraft generates a small AC component in the orthogonal direction, and the test quality is controlled at the center of the electrode cage by the electrostatic force.
[0027] In some preferred embodiments, S4 above, based on the attitude dynamics parameter data in the spacecraft attitude control system and combined with the relative motion dynamics model of the verification mass, constructs a set of equations containing the parameters to be identified, and may further include: S41, combined with the onboard sensors and control loop in the spacecraft attitude control system, attitude dynamics parameter data is obtained; among which: attitude dynamics parameter data includes: spacecraft angular velocity and angular acceleration data obtained by inertial sensors, electrostatic actuation force data obtained by electrostatic actuators, and micro-thrust data obtained by micro-thrusters; S42, based on the attitude dynamics parameter data, the relative motion dynamics model of the inspection quality is approximated to obtain the parametric model of the inspection quality, expressed as: In the formula, This represents the force per unit mass acting on the spacecraft. The effect of solar radiation pressure, This represents the force per unit mass of the inspected mass i. The electrostatic actuation effect in This represents the force per unit mass of the inspected mass i. The biasing force exerted by the spacecraft on the quality of the inspection; S43, based on the characteristics of the applied excitation signal, when the spacecraft oscillates at a small angle around any orthogonal axis of the Z-axis, the direction of that orthogonal axis is denoted as... The magnitude of the force exerted by sunlight on the spacecraft along its three axes is denoted as: remember: ,because For a small angle, that is Therefore, we can conclude that: In the formula, , Since all parameters on the right side of the equation are known quantities, the right side of the equation can be abbreviated as: In the formula, The x, y, and z coordinate components are represented by the sum of the terms on the right. S44, after the spacecraft undergoes multiple cycles of motion, N sets of sampled data can be obtained. Let the j-th set of sampled data be: This yields N sets of equations containing parameters to be identified; among them, the j-th set of equations containing parameters to be identified is expressed as: In some preferred embodiments, S5 above, which involves data processing of the equation set containing the parameters to be identified to obtain the optimal solution for the test mass bias force of the gravitational wave spacecraft, may further include: The least squares method is used to process the data of the system of equations containing parameters to be identified; where: S51, Note: S52, from which we can obtain the least squares solution to the system of equations: S53, thus enabling gravitational wave spacecraft to achieve test mass bias force. The identification, and simultaneously the identification of the magnitude of the light pressure experienced by the spacecraft when its normal direction is pointing towards the sun. .
[0028] Based on the same inventive concept, another embodiment of the present invention provides an on-orbit identification system for the mass bias force of a gravitational wave spacecraft.
[0029] Specifically, such as Figure 2 As shown, the on-orbit identification system for the gravitational wave spacecraft's test mass bias force provided in this embodiment may include: Inspection quality model building module, which is used to build a relative motion dynamics model of inspection quality; Measurement Condition Design Module: This module is used to design the solar pointing measurement conditions so that the solar radiation pressure, which is the main non-conservative force disturbance of the spacecraft, is along the Z-axis of the spacecraft. The excitation signal application module is used to apply a periodic excitation signal to the relative motion dynamics model of the test quality under the condition of sun pointing measurement, so that the spacecraft generates a small AC component in the orthogonal direction. The module for constructing parameters to be identified is used to construct a set of equations containing the parameters to be identified based on the control parameter data in the spacecraft attitude control system and the relative motion dynamics model of the test mass under the action of the excitation signal. The bias force identification module is used to process the equations containing the parameters to be identified, obtain the optimal solution of the bias force of the gravitational wave spacecraft for the test mass, and thus realize the identification of the bias force.
[0030] It should be noted that the steps in the method provided by the present invention can be implemented using corresponding modules, devices, units, etc. in the system. Those skilled in the art can refer to the technical solution of the method to realize the composition of the system. That is, the embodiments in the method can be understood as preferred examples for building the system, and will not be elaborated here.
[0031] Based on the same inventive concept, other embodiments of the present invention also provide an on-board computer, including a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it can be used to perform any of the methods described above in the present invention, or to run any of the systems described above in the present invention.
[0032] Optionally, the memory is used to store programs; the memory may include volatile memory, such as random-access memory (RAM), such as static random-access memory (SRAM), double data rate synchronous dynamic random-access memory (DDR SDRAM), etc.; the memory may also include non-volatile memory, such as flash memory. The memory is used to store computer programs (such as application programs and functional modules that implement the above methods), computer instructions, etc., and the aforementioned computer programs and computer instructions can be partitioned and stored in one or more memories. Furthermore, the aforementioned computer programs, computer instructions, data, etc., can be accessed by the processor.
[0033] A processor is used to execute computer programs stored in memory to implement the various steps of the methods or various modules of the systems involved in the above embodiments. For details, please refer to the relevant descriptions in the preceding method and system embodiments.
[0034] The processor and memory can be separate structures or integrated structures. When the processor and memory are separate structures, they can be coupled together via a bus.
[0035] Based on the same inventive concept, other embodiments of the present invention also provide a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, can be used to perform the method of any of the above-described embodiments of the present invention, or to run the system of any of the above-described embodiments of the present invention.
[0036] Computer-readable media include computer storage media and communication media, wherein communication media include any medium that facilitates the transfer of computer programs from one place to another. Storage media can be any available medium accessible to a general-purpose or special-purpose computer. An exemplary storage medium is coupled to a processor, enabling the processor to read information from and write information to the storage medium. Of course, the storage medium can also be a component of the processor. The processor and storage medium can reside in an ASIC. Alternatively, the ASIC can reside in a user device. Of course, the processor and storage medium can also exist as discrete components in a communication device.
[0037] The following detailed explanation, using a specific application example, illustrates the application process of the technical solution provided in the above embodiments of the present invention in a specific identification task.
[0038] In this specific application example, taking an equilateral triangular space gravitational wave detection spacecraft formation as an example, the spacecraft has two test masses, TM1 and TM2, arranged as follows: Figure 3 As shown; Figure 3 In the diagram, S / C represents the spacecraft, and the corresponding coordinate system is... O 0 x 0 y 0 z 0, Origin of the coordinate system O 0 is located at the spacecraft's center of mass CoM; TM1 and TM2 represent two inspection masses, corresponding to the coordinate system... O 1 x 1 y 1 z 1 and O 2 x 2 y 2 z 2. Origin of the coordinate system O 1. O 2 is located at the center of the two nominal inspection quality locations; α Indicates the included angle of the moving optical components. r 01 , r 02 This represents the position vector from the spacecraft's center of mass to the centers of TM1 and TM2.
[0039] The on-orbit identification process involved in this specific application example, such as Figure 5 As shown, it includes the following steps: First, a relative motion dynamic model for the quality inspection was constructed; Secondly, by designing specific solar pointing measurement conditions, the main non-conservative force disturbance of the spacecraft—solar radiation pressure—is made to run along the spacecraft's Z-axis. Next, by applying a periodic excitation signal, it generates a small AC component in the orthogonal direction; Then, based on data from inertial sensors, electrostatic actuators, and micro-thrusters in the spacecraft attitude control system, and combined with the relative motion dynamics model for verifying mass, a set of equations containing parameters to be identified was constructed. Finally, the data is processed using methods such as least squares to obtain the optimal solution for the bias force of the gravitational wave spacecraft, thereby realizing the identification of the bias force.
[0040] Starting from the displacement of the inspection mass relative to the electrode cage, the relative dynamic model of the inspection mass is constructed as follows: In the formula, , These represent the displacement vectors of TM1 and TM2 relative to their nominal positions (typically the center of the electrode cage), respectively. , Take the second derivative with respect to time; , These represent the position vectors from the spacecraft's center of mass to the centers of TM1 and TM2, respectively. Represents the angular velocity of a spacecraft; , These represent the forces per unit mass acting on TM1 and TM2, including the stiffness effect, electrostatic actuation force, and bias force acting on the inspection mass, respectively. , These represent the forces acting on a spacecraft per unit mass, including solar radiation pressure and micro-thrust acting on the spacecraft. , , Representing parameters respectively , , Find the first derivative with respect to time.
[0041] Before identifying the bias force of a gravitational wave spacecraft on the test mass, the measurement conditions need to be designed. First, micro-thrusters are used to control the spacecraft's attitude only, keeping the spacecraft's Z-axis pointing towards the sun and the Y-axis along the orbital plane normal. Depending on the spacecraft configuration, such as... Figure 4 As shown, the solar radiation pressure is along the -Z axis at this time. Then, when measuring the bias force of the gravitational wave spacecraft on the test mass, an excitation signal is added to the relative dynamic model of the test mass using a micro-thruster, causing the spacecraft to oscillate at a small angle around a certain axis (e.g., a sinusoidal periodic motion around the Y-axis, with a peak oscillation angle not exceeding 0.1° and an oscillation period of 500s). The oscillation angle is denoted as... At this point, the inspection mass is controlled at the center of the electrode cage by electrostatic force. Combined with onboard sensors and the control loop, parameters such as the inspection mass displacement, electrostatic force, spacecraft angular velocity, and angular acceleration can be obtained.
[0042] Taking the inspection mass TM1 as an example, since the inspection mass is locked at the center of the electrode cage, its displacement relative to the center of the electrode cage is approximately zero. Simultaneously, the disturbance force generated by the stiffness effect is approximately zero. At this point, the inspection mass TM1 is mainly subjected to electrostatic actuation force and bias force. Therefore, the relative dynamic model of the inspection mass TM1 can be approximated as: In the formula, express The effect of solar radiation pressure, express The electrostatic actuation force on TM1 in the middle, express The bias force exerted by the spacecraft on TM1. Based on the characteristics of the applied excitation signal, taking a sinusoidal periodic motion around the Y-axis as an example, the magnitude of the force exerted by solar radiation pressure on the spacecraft along its three axes is determined. It can be written as: remember ,because For a small angle, that is Therefore, we can conclude that: In the formula, If all parameters on the right side of the equation are known quantities, then the right side of the equation can be simply written as: After the spacecraft undergoes multiple cycles of motion, N sets of sampled data can be obtained. Let the j-th set of data be denoted as: This yields N sets of equations containing the parameters to be identified, as shown below: remember: Therefore, the least squares solution to the system of equations can be obtained as follows: This allows gravitational wave spacecraft to test mass bias forces. The identification, and simultaneously the identification of the magnitude of the light pressure experienced by the spacecraft when its normal direction is pointing towards the sun. .
[0043] The on-orbit identification method and system for the bias force of a gravitational wave spacecraft's test mass provided in the above embodiments of the present invention achieves the identification of the bias force of the gravitational wave spacecraft's test mass based on on-orbit measured data. It does not rely on a pre-established finite element model on the ground and can truly reflect the actual on-orbit state of the spacecraft, avoiding errors caused by the difference between the design model and the actual object. Through special working condition design, the coupling between the bias force and external disturbances such as solar radiation pressure is effectively separated, improving the reliability and accuracy of the identification results. It can be repeatedly executed at different stages of the spacecraft's entire life cycle (such as the initial stage of orbit insertion, after propellant consumption, and after thermal deformation stabilization), realizing on-orbit monitoring of the long-term evolution of the bias force, and providing a basis for spacecraft state assessment, model correction, and compensation strategy optimization.
[0044] Any matters not covered in the above embodiments of the present invention are well-known in the art.
[0045] The specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various modifications or variations within the scope of the claims, which do not affect the essence of the present invention.
Claims
1. An on-orbit identification method for the test mass bias force of a gravitational wave spacecraft, characterized in that, include: Construct a relative motion dynamic model for the inspection quality; The design of the solar pointing measurement condition is such that the solar radiation pressure, which is the main non-conservative force disturbance of the spacecraft, is along the Z-axis of the spacecraft. Under the aforementioned sun-pointing measurement condition, a periodic excitation signal is applied to the relative motion dynamics model of the test quality, causing the spacecraft to generate a small AC component in the orthogonal direction; Under the action of the excitation signal, based on the attitude dynamics parameter data in the spacecraft attitude control system and combined with the relative motion dynamics model of the test mass, a set of equations containing the parameters to be identified is constructed. Data processing is performed on the set of equations containing the parameters to be identified to obtain the optimal solution for the bias force of the test mass of the gravitational wave spacecraft, thereby realizing the identification of the bias force.
2. The on-orbit identification method for the test mass bias force of a gravitational wave spacecraft according to claim 1, characterized in that, The construction of the relative motion dynamics model for the inspection quality includes: Starting from the displacement of the inspection mass relative to the electrode cage, a relative motion dynamic model of the inspection mass is constructed for each inspection mass i inside the spacecraft, expressed as: ， In the formula, The displacement vector representing the inspection quality i relative to the nominal position (usually the center of the electrode cage). Take the second derivative with respect to time; This represents the position vector from the spacecraft's center of mass to the center of the test mass i; Represents the angular velocity of a spacecraft; This represents the force per unit mass acting on the inspection quality i, including the stiffness effect, electrostatic actuation force, and bias force acting on the inspection quality; This represents the force per unit mass acting on the spacecraft, including the pressure from sunlight and micro-thrust acting on the spacecraft. , Representing parameters respectively , Find the first derivative with respect to time.
3. The on-orbit identification method for the test mass bias force of a gravitational wave spacecraft according to claim 1, characterized in that, The design for sun-pointing measurement conditions ensures that the solar radiation pressure, which is the main non-conservative force disturbance of the spacecraft, is along the spacecraft's Z-axis, including: Micro-thrusters are used to control the attitude of the spacecraft, keeping the spacecraft's normal Z-axis pointing towards the sun and the spacecraft's Y-axis along the orbital plane normal. According to the spacecraft configuration, the solar radiation pressure is along the spacecraft's Z-axis.
4. The on-orbit identification method for the test mass bias force of a gravitational wave spacecraft according to claim 1, characterized in that, The application of the periodic excitation signal, causing the spacecraft to generate small AC components in orthogonal directions, includes: When measuring the bias force of a gravitational wave spacecraft on a test mass, a periodic excitation signal is added to the relative motion dynamics model of the test mass using a micro-thruster. This causes the spacecraft to oscillate around any orthogonal axis of the Z-axis at an angle less than a set threshold. The oscillation angle is denoted as... At this time, the spacecraft generates a small AC component in the orthogonal direction, and the test quality is controlled at the center of the electrode cage by the electrostatic force.
5. The on-orbit identification method for the test mass bias force of a gravitational wave spacecraft according to claim 1, characterized in that, Based on the attitude dynamics parameter data in the spacecraft attitude control system, and combined with the relative motion dynamics model of the test mass, a set of equations containing the parameters to be identified is constructed, including: By combining the onboard sensors and control loops in the spacecraft's attitude control system, attitude dynamics parameter data can be obtained; Based on the aforementioned attitude dynamics parameter data, the relative motion dynamics model of the inspection quality is approximated to obtain the parametric model of the inspection quality, expressed as: ， In the formula, This represents the force per unit mass acting on the spacecraft. The effect of solar radiation pressure, This represents the force per unit mass of the inspected mass i. The electrostatic actuation effect in This represents the force per unit mass of the inspected mass i. The biasing force exerted by the spacecraft on the quality of the inspection; Based on the characteristics of the applied excitation signal, when the spacecraft oscillates at an angle less than a set threshold around any orthogonal axis of the Z-axis, the direction of that orthogonal axis is denoted as... The magnitude of the force exerted by sunlight on the spacecraft along its three axes is denoted as: ， remember: ,because An angle less than a set threshold, i.e. Therefore, we can conclude that: ， In the formula, , Since all parameters on the right side of the equation are known quantities, the right side of the equation can be simplified as: ， In the formula, The x, y, and z coordinate components are represented by the summation of the terms on the right. After the spacecraft undergoes multiple cycles of motion, it obtains N sets of sampled data. Let the j-th set of sampled data be denoted as: ， This yields N sets of equations containing parameters to be identified; among them, the j-th set of equations containing parameters to be identified is expressed as: ， In the formula, Indicates the spacecraft's oscillation angle The j-th data set; The x-axis component of the bias force exerted by the spacecraft on the test quality; This represents the magnitude of the light pressure force when sunlight shines directly along the z-axis of the spacecraft. The y-axis component of the bias force exerted by the spacecraft on the test quality; This represents the z-axis component of the bias force exerted by the spacecraft on the test quality. express The j-th sampled data.
6. The on-orbit identification method for the test mass bias force of a gravitational wave spacecraft according to claim 5, characterized in that, The attitude dynamics parameter data includes: spacecraft angular velocity and angular acceleration data obtained from inertial sensors, electrostatic actuation force data obtained from electrostatic actuators, and micro-thrust data obtained from micro-thrusters.
7. The on-orbit identification method for the test mass bias force of a gravitational wave spacecraft according to claim 1, characterized in that, The process of processing the equations containing the parameters to be identified to obtain the optimal solution for the test mass bias force of the gravitational wave spacecraft includes: The least squares method is used to process the data of the system of equations containing the parameters to be identified; wherein: remember: ， ， Therefore, the least squares solution to the system of equations can be obtained as follows: ， This enables the gravitational wave spacecraft to test the mass bias force. The identification, and simultaneously the identification of the magnitude of the light pressure experienced by the spacecraft when its normal direction is pointing towards the sun. .
8. An on-orbit identification system for the test mass bias force of a gravitational wave spacecraft, characterized in that, include: Inspection quality model building module, which is used to build a relative motion dynamics model of inspection quality; Measurement Condition Design Module: This module is used to design the solar pointing measurement conditions so that the solar radiation pressure, which is the main non-conservative force disturbance of the spacecraft, is along the Z-axis of the spacecraft. An excitation signal application module is used to apply a periodic excitation signal to the relative motion dynamics model of the test quality under the sun pointing measurement condition, so that the spacecraft generates a small AC component in the orthogonal direction. The module for constructing parameters to be identified is used to construct a set of equations containing the parameters to be identified based on the attitude dynamics parameter data in the spacecraft attitude control system and the relative motion dynamics model of the test mass under the action of the excitation signal. The bias force identification module is used to process the equations containing the parameters to be identified to obtain the optimal solution of the bias force of the gravitational wave spacecraft for the test mass, thereby realizing the identification of the bias force.
9. An on-board computer, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the computer program, it can be used to perform the method of any one of claims 1-7.
10. A computer-readable storage medium having a computer program stored thereon, characterized in that, When executed by a processor, the computer program can be used to perform the method of any one of claims 1-7.