Spacecraft center of mass on-orbit identification technology ground verification device and method thereof

Abstract

The application discloses a ground verification device and method for spacecraft mass center on-orbit identification technology. The ground verification device for spacecraft mass center on-orbit identification technology comprises a mass center measuring platform, a driving mechanism, an accelerometer and a mass center adjusting mechanism. The mass center measuring platform comprises a torsion pendulum platform, a torsion pendulum base and a suspension wire. The torsion pendulum platform is used for placing a spacecraft simulator. Only a rotation degree of freedom around the gravity direction is reserved between the torsion pendulum platform and the torsion pendulum base. The mass center measuring platform comprises the torsion pendulum platform and the torsion pendulum base. The torsion pendulum platform and the torsion pendulum base are connected through the suspension wire. The motion state of the spacecraft in orbit is simulated. The acceleration of the spacecraft is measured through the accelerometer, and the mass center position is calculated. The mass center position of the spacecraft simulator is adjusted through the mass center adjusting device, so that the rotation center of the spacecraft simulator is equivalent to the mass center of the spacecraft in orbit.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aerospace technology, in particular to a spacecraft mass center on-orbit identification technology ground verification device and method thereof. BACKGROUND

[0002] In related technologies, accurately establishing an earth gravity field model is of great significance to the fields of geodesy, oceanography, seismology, and resource exploration. A high-precision gravity spacecraft is one of the necessary equipment for establishing an accurate earth gravity field model. To ensure accurate measurement of the earth's gravity, it is necessary to accurately remove other non-conservative forces other than the earth's gravity and ensure that the mass center position of the gravity spacecraft and the detection center position of the accelerometer test mass are within a certain range, which is a key technical challenge.

[0003] During the on-orbit period of the spacecraft, factors such as spacecraft structural deformation, material outgassing, and propellant consumption will cause the spacecraft mass center to change relative to the spacecraft frame. The accelerometer and the spacecraft frame are approximately rigidly connected, so the actual positions of the accelerometer test mass mass center and the spacecraft mass center will change over time.

[0004] However, compared with the on-orbit working environment of the spacecraft, the spacecraft is affected by gravity factors in ground experiments. This is because the earth's gravity will have a significant impact on the movement and balance of the spacecraft. In ground experiments, the spacecraft is difficult to reproduce the real running state on the orbit, and is affected by the earth's gravity, resulting in differences in its behavior on the orbit. This situation affects the accurate identification and adjustment of the mass center position.

[0005] Therefore, when performing ground equivalent verification schemes for mass center identification and adjustment, it is difficult to simulate the state of the spacecraft on the orbit in the ground verification test. SUMMARY

[0006] The present application aims to at least solve one of the technical problems existing in the prior art. To this end, the present application proposes a spacecraft mass center on-orbit identification technology ground verification method, which can simulate the state of the spacecraft on the orbit when performing ground verification experiments.

[0007] The present application also proposes a spacecraft mass center on-orbit technology ground verification device for implementing the above-mentioned spacecraft mass center on-orbit technology ground verification method.

[0008] The spacecraft mass center on-orbit identification technology ground verification method according to the first aspect embodiment of the present application comprises the following steps:

[0009] A mass center measurement platform is set up, and the degrees of freedom of the mass center measurement platform are limited so that the mass center measurement platform only retains the degree of freedom of rotation around the gravity direction;

[0010] The spacecraft simulator is placed on the center-of-mass measuring platform, the center-of-mass position of the spacecraft simulator is measured by using the center-of-mass inertia measuring instrument, an accelerometer and an attitude measuring instrument are arranged above the center-of-mass measuring platform, the center-of-mass position of the spacecraft simulator is arranged in a staggered manner with the position of the measuring center of the accelerometer, that is, the center-of-mass position of the spacecraft simulator has a deviation from the position of the measuring center;

[0011] A torque excitation is applied to the spacecraft simulator, so that the spacecraft simulator generates a periodic swing on the center-of-mass measuring platform;

[0012] The attitude of the spacecraft simulator in swing is measured by using the attitude measuring instrument, and the linear acceleration of the spacecraft simulator at the center of the accelerometer is measured by using the accelerometer;

[0013] The center-of-mass position of the spacecraft simulator at this time is calculated by measuring the attitude and acceleration of the spacecraft simulator;

[0014] According to the calculated center-of-mass position, the center-of-mass position of the spacecraft simulator is adjusted, and the deviation of the center-of-mass of the spacecraft simulator from the center of the accelerometer is controlled, that is, the distance between the center-of-mass of the spacecraft simulator and the center of the accelerometer is controlled.

[0015] The spacecraft center-of-mass on-orbit identification technology ground verification method according to the embodiment of the application has at least the following beneficial effects:

[0016] By arranging the center-of-mass measuring platform, the spacecraft simulator is placed on the center-of-mass measuring platform, the spacecraft simulator can perform a torsional pendulum motion on the center-of-mass measuring platform, the center-of-mass position of the spacecraft simulator can be calculated by measuring various parameters of the spacecraft in the torsional pendulum motion, in the ground verification test of the spacecraft simulator, the center-of-mass of the spacecraft simulator can also be measured by other methods, so as to verify the accuracy of the above calculation method, and the calculation method can be applied to the measurement of the spacecraft on-orbit.

[0017] By limiting the degrees of freedom of the center-of-mass measuring platform, the spacecraft simulator on the center-of-mass measuring platform will not move under the action of gravity, and a pure ideal rotation can be maintained, so as to simulate the state of the spacecraft on-orbit, and the accuracy of the ground verification experiment is improved.

[0018] In the measurement of the spacecraft simulator, the parameters of the spacecraft simulator are measured by using the accelerometer and the attitude measuring instrument. In order to make the measurement result more accurate, and in order to make the measurement state of the spacecraft simulator close to the state of the spacecraft in orbit, the measurement center of the accelerometer needs to keep a certain deviation from the center of mass of the spacecraft simulator. The center of mass inertia measuring instrument is used to determine the center of mass position of the spacecraft simulator before entering the torsion state, so as to facilitate the installation of the accelerometer, and make the measurement center of the accelerometer keep a certain deviation from the center of mass of the spacecraft simulator. The spacecraft simulator is excited by a torque, so that the spacecraft simulator enters a periodic torsion motion state. After the spacecraft simulator enters the torsion state, the center of mass position of the spacecraft simulator is calculated according to the parameters measured by the accelerometer and the attitude measuring instrument, and the center of mass of the spacecraft simulator is adjusted, so that the deviation between the center of mass of the spacecraft simulator and the measurement center of the acceleration measuring instrument is kept within a certain range, and the rotation center of the spacecraft simulator is equivalent to the center of mass of the spacecraft in orbit when the accelerometer is measured.

[0019] The center of mass measuring platform and the spacecraft simulator are used to simulate the state of the spacecraft in orbit, and the ground verification experiment of the center of mass measurement and the center of mass adjustment is performed, so as to verify the accuracy of the measurement method, and then the measurement method can be applied to the actual spacecraft in orbit measurement, the research efficiency is improved, and the research cost is reduced.

[0020] According to some embodiments of the application, the repeated experiments are performed according to the above steps, and the error between the center of mass measurement of the spacecraft simulator and the ideal value is calculated according to the results of the repeated experiments, so as to ensure the accuracy of the ground experiment.

[0021] According to some embodiments of the application, the center of mass measuring platform includes a torsion base and a torsion platform, and the center of mass position measurement formula of the spacecraft simulator is:

[0022]

[0023] In the formula, δ is the offset of the center of mass of the spacecraft simulator; k is the torsional stiffness coefficient of the torsion platform; m0 is the mass of the spacecraft simulator; m p is the mass of the torsion platform; ω1 and ω2 are respectively the vibration frequencies of the spacecraft simulator rotating 0° and 180° around the rotation shaft of the torsion platform; is the angle between the horizontal projection of the spacecraft simulator and the perpendicular line of the spacecraft simulator; and R is the distance between the geometric center of the spacecraft simulator and the rotation shaft of the torsion platform.

[0024] According to some embodiments of the application, the method for calculating the error includes:

[0025] Obtaining angular velocity and angular acceleration of the spacecraft simulator and attitude information of the spacecraft simulator;

[0026] Setting Asum matrix, ω matrix and dω matrix, the Asum matrix contains three-dimensional coordinates (x, y, z) of the spacecraft, the ω matrix is classification of angular velocity of the spacecraft simulator in x, y, z directions (ω x ,ω y ,ω z );The dω matrix is classification of angular acceleration of the spacecraft simulator in x, y, z directions (dω x ,dω y ,dω z );

[0027] Adding noise on ω z ,dω z ,x,y, and setting the data after adding noise as ω z’ ,dω z’ ,x', y'

[0028] Setting a 2*2 matrix A and a 2*1 matrix a, wherein:

[0029]

[0030] a=[x,y]

[0031] Calculating the deviation of the spacecraft simulator in x direction and the deviation in y direction, and the calculation formula is as follows:

[0032] l=(A×A) -1 ×A×a

[0033] error=l-[0.005;0.005]

[0034] Wherein, the first column of the matrix l is the deviation of the spacecraft simulator in x direction, and the second column of the matrix l is the deviation of the spacecraft simulator in y direction;The result 2*1 matrix error matrix is the error of the current estimated parameter and the expected value [0.005;0.005], the first column of the error matrix is the error in x direction, which is set as x error ,error matrix, the second column of the error matrix is the error in y direction, which is set as y error .

[0035] Set the error in x direction as x error , and set the error in y direction as y error ;

[0036] Repeat the experiment, record the error of each time;Set the error in x direction and the error in y direction of the i-th experiment as x error (i) and yerror (i); calculate the sum of squares of X direction and Y direction errors, and take the root mean square (RMS) error X error and Y error ; wherein

[0037]

[0038] The root mean square errors of X direction and Y direction are combined to calculate the comprehensive error, and the calculation formula is as follows:

[0039]

[0040] L error is the comprehensive error to be solved.

[0041] The spacecraft mass center on-orbit identification technology ground verification device according to the second aspect embodiment of the present application comprises:

[0042] A mass center measurement platform, the mass center measurement platform comprises a torsion pendulum platform, a torsion pendulum base and a suspension wire, the torsion pendulum platform is used for placing a spacecraft simulator, the torsion pendulum platform is arranged above the torsion pendulum base, the torsion pendulum platform and the torsion pendulum base are connected through the suspension wire, and only a degree of freedom of rotation around the gravity direction is reserved between the torsion pendulum platform and the torsion pendulum base;

[0043] A driving mechanism arranged on the mass center measurement platform is used for applying a torque excitation to the spacecraft simulator;

[0044] An accelerometer arranged above the mass center measurement platform is used for measuring the acceleration of the spacecraft simulator;

[0045] A mass center adjusting mechanism arranged on the mass center measurement platform is used for adjusting the mass center position of the spacecraft simulator.

[0046] The spacecraft mass center on-orbit identification technology ground verification device according to the embodiment of the present application has at least the following beneficial effects:

[0047] By arranging the mass center measurement platform, the mass center position of the spacecraft simulator can be measured, the mass center measurement platform is composed of a torsion pendulum platform, a torsion pendulum base and a suspension wire, the torsion pendulum platform and the torsion pendulum base are connected through the suspension wire, so that the torsion pendulum platform can perform a torsion pendulum motion on the torsion pendulum base, and the torsion pendulum realized through the suspension wire is less affected by external force and has high motion sensitivity, so that the state of the spacecraft in on-orbit motion can be more accurately simulated, only a degree of freedom of rotation around the gravity direction is reserved between the torsion pendulum platform and the torsion pendulum base, so that the spacecraft simulator on the torsion pendulum platform will not perform up-and-down swinging and other actions under the influence of gravity, and a pure ideal rotation can be maintained, so that the motion state of the spacecraft in on-orbit motion can be better simulated to perform more accurate ground verification tests.

[0048] The torsional pendulum platform is arranged above the torsional pendulum base, so that there is no shielding structure above the torsional pendulum platform, which means that the size of the spacecraft simulator does not affect the overall installation, so that a thinner suspension wire can be selected flexibly to improve the precision and sensitivity of the system.

[0049] The driving mechanism is arranged to apply torque excitation to the spacecraft simulator to make the spacecraft simulator enter a torsional pendulum state, facilitating subsequent measurement and adjustment; the accelerometer is arranged to collect various motion parameters of the spacecraft simulator when the spacecraft simulator moves, so as to calculate the center of mass position of the spacecraft simulator when the spacecraft simulator moves; and the center of mass adjusting mechanism is arranged to adjust the center of mass of the spacecraft simulator according to the center of mass position of the spacecraft simulator, so that the center of mass of the spacecraft simulator and the measurement center of the accelerometer maintain a certain deviation.

[0050] According to some embodiments of the present application, the suspension wire includes a first suspension wire, a second suspension wire, a third suspension wire, a fourth suspension wire and a fifth suspension wire, the arrangement direction of the first suspension wire is parallel to the direction of gravity, the second suspension wire and the third suspension wire intersect at a first intersection point, the fourth suspension wire and the fifth suspension wire intersect at a second intersection point, and the first intersection point and the second intersection point are arranged on the torsional pendulum platform.

[0051] According to some embodiments of the present application, the driving mechanism is a reaction flywheel.

[0052] According to some embodiments of the present application, the center of mass adjusting mechanism includes a moving rail, a mass block and a driving unit, the moving rail is installed on the spacecraft simulator, the mass block is movably arranged on the moving rail, and the driving unit is used to drive the mass block to move.

[0053] According to some embodiments of the present application, the torsional pendulum platform includes a connecting portion, the connecting portion has a bending portion, the bending portion is arranged at the bottom of the torsional pendulum base, a support wire is arranged on the bending portion, the support wire is connected with the torsional pendulum base, and the arrangement direction of the support wire is parallel to the direction of gravity.

[0054] According to some embodiments of the present application, the accelerometer is connected with a moving mechanism, and the moving mechanism is used to move the position of the accelerometer.

[0055] Additional aspects and advantages of the present application will be in part apparent and in part pointed out hereinafter. BRIEF DESCRIPTION OF DRAWINGS

[0056] The accompanying drawings are intended to provide a further understanding of the technical solutions of the present application, and constitute a part of the specification, and are used together with the embodiments disclosed in the present application to explain the technical solutions of the present application, and do not constitute a limitation on the technical solutions of the present application.

[0057] Fig. 1 A flow chart of a ground verification method for a spacecraft centroid on-orbit identification technology of an embodiment of the present application;

[0058] Fig. 2 A schematic diagram of the overall structure of a ground verification device for a spacecraft centroid on-orbit identification technology of an embodiment of the present application;

[0059] Fig. 3 A schematic diagram of the structure of a centroid measurement platform of a ground verification device for a spacecraft centroid on-orbit identification technology of an embodiment of the present application.

[0060] The accompanying drawings are intended to provide a further understanding of the technical solutions of the present application, and constitute a part of the specification, and are used together with the embodiments disclosed in the present application to explain the technical solutions of the present application, and do not constitute a limitation on the technical solutions of the present application. DETAILED DESCRIPTION

[0061] The embodiments of the present application are described in detail below, and examples of the embodiments are shown in the accompanying drawings, wherein the same or similar notations represent the same or similar elements or elements having the same or similar functions throughout. The embodiments described below by referring to the accompanying drawings are exemplary, and are only used to explain the present application, and cannot be understood as a limitation on the present application.

[0062] In the description of the present application, it should be understood that, in relation to the orientation description, for example, the orientation or position relationship indicated by up, down, front, back, left, right, etc. is based on the orientation or position relationship shown in the drawings, and is only for the purpose of facilitating the description of the present application and simplifying the description, and does not indicate or imply that the indicated device or element must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation on the present application.

[0063] In the description of the present application, if the meaning of several is more than one, the meaning of multiple is more than two, greater than, less than, more than, etc. is understood as not including the number, above, below, etc. is understood as including the number. If it is described that the first, the second is only used for the purpose of distinguishing technical features, and cannot be understood as indicating or implying relative importance or implicitly indicating the number of indicated technical features or implicitly indicating the sequence of indicated technical features.

[0064] In the description of the present application, the words such as setting, installing, connecting and the like should be understood in a broad sense unless otherwise explicitly limited, and the specific meanings of the words in the present application can be reasonably determined by those skilled in the art in combination with the specific content of the technical solutions.

[0065] In the description of the present application, the description of the terms "one embodiment", "some embodiments", "exemplary embodiment", "example", "specific example" or "some examples" means that the specific features, structures, materials or characteristics described in combination with the embodiment or example are included in at least one embodiment or example of the present application. In the present specification, the exemplary description of the above terms does not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any one or more embodiments or examples in a suitable manner.

[0066] In some alternative embodiments, the functions / operations mentioned in the block diagram can not occur in the order mentioned in the operation diagram. For example, depending on the functions / operations involved, two blocks shown in succession can actually be executed substantially simultaneously or the blocks can sometimes be executed in reverse order. In addition, the embodiments presented and described in the flowcharts of the present application are provided by way of example, with the purpose of providing a more comprehensive understanding of the technology. The disclosed method is not limited to the operations and logical flows presented herein. Alternative embodiments are contemplated in which the order of various operations is changed and in which sub-operations described as part of larger operations are independently executed.

[0067] According to the spacecraft mass center on-orbit identification technology ground verification method of the first aspect embodiment of the present application, please refer to Figs. 1-3 , which comprises the following steps:

[0068] S1: setting a mass center measurement platform 110, limiting the degrees of freedom of the mass center measurement platform 110, so that the mass center measurement platform 110 only retains the freedom of rotation around the direction of gravity;

[0069] S2: placing a spacecraft simulator on the mass center measurement platform 110, measuring the mass center position of the spacecraft simulator using a mass center inertia measuring instrument, the mass center measurement platform 110 is provided with an accelerometer 130 and an attitude measuring instrument 180, the accelerometer 130 has a measurement center, and the position of the measurement center has a deviation from the mass center position of the spacecraft simulator;

[0070] S3: applying a torque excitation to the spacecraft simulator, so that the spacecraft simulator generates a periodic swing on the mass center measurement platform 110;

[0071] S4: measuring the attitude of the spacecraft simulator in the swing state by using the attitude measuring instrument 180, and measuring the linear acceleration of the spacecraft simulator at the center of the accelerometer 130 by using the accelerometer 130;

[0072] S5: calculating the center of mass position of the spacecraft simulator at this time by using the measured attitude and acceleration of the spacecraft simulator;

[0073] S6: adjusting the center of mass position of the spacecraft simulator according to the calculated center of mass position, and controlling the deviation between the center of mass of the spacecraft simulator and the center of the accelerometer 130.

[0074] In the embodiment of the application, the spacecraft simulator is placed on the center of mass measuring platform 110, and the spacecraft simulator can perform a torsional pendulum motion on the center of mass measuring platform 110. The center of mass position of the spacecraft simulator can be calculated by measuring various parameters of the spacecraft in the torsional pendulum motion. In the ground verification test of the spacecraft simulator, the center of mass of the spacecraft simulator can also be measured by other methods, so as to verify the accuracy of the above calculation method, and thus the calculation method can be applied to the measurement of the spacecraft in orbit.

[0075] By limiting the degrees of freedom of the center of mass measuring platform 110, the spacecraft simulator on the center of mass measuring platform 110 will not move under the action of gravity, and can maintain a pure ideal rotation, so as to simulate the state of the spacecraft in orbit, thereby improving the accuracy of the ground verification test.

[0076] In the measurement of the spacecraft simulator, the accelerometer 130 and the attitude measuring instrument 180 are used to measure various parameters of the spacecraft simulator. In order to make the measurement result more accurate, and in order to make the measurement state of the spacecraft simulator close to the state of the spacecraft in orbit, the measurement center of the accelerometer 130 and the center of mass of the spacecraft simulator need to maintain a certain deviation. The embodiment of the application determines the center of mass position of the spacecraft simulator before entering the torsional pendulum state by using the center of mass inertia measuring instrument, so as to facilitate the installation of the accelerometer 130, so that the measurement center of the accelerometer 130 and the center of mass of the spacecraft simulator maintain a certain deviation. By applying a torque excitation to the spacecraft simulator, the spacecraft simulator enters a periodic torsional pendulum motion state. After the spacecraft simulator enters the torsional pendulum state, the center of mass position of the spacecraft simulator is calculated by using the parameters measured by the accelerometer 130 and the attitude measuring instrument 180, and the center of mass of the spacecraft simulator is adjusted, so that the deviation between the center of mass of the spacecraft simulator and the measurement center of the accelerometer is maintained within a certain range.

[0077] The position of the measurement center is set to be deviated from the position of the center of mass of the spacecraft simulator, which aims to solve the following problem: when the spacecraft needs to only have attitude motion in orbit, the reference point of rotation is the center of mass of the spacecraft at this time, the center of mass is regarded as (0, 0, 0), the accelerometer 130 has a certain installation distance from the center of mass, and the position of the center of mass can be determined through the distance between the two; when the spacecraft is verified on the center of mass measurement platform 110, the rotation center is not the center of mass of the spacecraft but the rotation center of the pendulum platform 150. In order to ensure that the rotation center of the spacecraft simulator is equivalent to the center of mass of the spacecraft in orbit at this time, the offset amount of the center of mass of the spacecraft simulator from the accelerometer 130 is adjusted, so as to simulate the offset amount of the center of mass of the spacecraft from the accelerometer 130 when the spacecraft is in orbit.

[0078] Specifically, in the embodiment of the present application, the center of mass measurement platform 110 includes a pendulum platform 150, a pendulum base 160 and a suspension wire, the pendulum platform 150 is used to place the spacecraft simulator, the pendulum platform 150 is arranged above the pendulum base 160, the pendulum platform 150 and the pendulum base 160 are connected through the suspension wire, and only the degree of freedom of rotation around the direction of gravity is reserved between the pendulum platform 150 and the pendulum base 160; specifically, the suspension wire includes a first suspension wire 171, a second suspension wire 172, a third suspension wire 173, a fourth suspension wire 174 and a fifth suspension wire 175, the arrangement direction of the first suspension wire 171 is parallel to the direction of gravity, the second suspension wire 172 and the third suspension wire 173 intersect at a first intersection point, the fourth suspension wire 174 and the fifth suspension wire 175 intersect at a second intersection point, the first intersection point and the second intersection point are arranged on the pendulum platform 150, and the connection point of the first suspension wire 171 and the pendulum platform 150 is a third intersection point. By arranging five suspension wires, the pendulum base 160 and the pendulum platform 150 form a five-wire pendulum structure, so as to be able to limit the degree of freedom of the pendulum platform 150, so that the pendulum platform 150 can only rotate around the direction of gravity, and when the pendulum platform 150 is subjected to an external force (such as gravity) to generate a motion trend in other directions, one or more suspension wires can generate a pulling force that counteracts the external force, so as to limit the degree of freedom of the pendulum platform 150.

[0079] Specifically, the distance from the first intersection point to the rotation center of the pendulum platform 150, the distance from the second intersection point to the rotation center of the pendulum platform 150 and the distance from the third intersection point to the rotation center of the pendulum platform 150 are equal.

[0080] Specifically, in the embodiment of the present application, the formula for measuring the center of mass of the spacecraft simulator is:

[0081]

[0082] wherein δ is the offset of the center of mass of the spacecraft simulator; k is the torsional stiffness coefficient of the torsion pendulum platform; m0 is the mass of the spacecraft simulator; m p is the mass of the torsion pendulum platform; ω1 and ω2 are the vibration frequencies of the spacecraft simulator when the spacecraft simulator rotates 0° and 180° around the rotation axis of the torsion pendulum platform, respectively; is the angle between the horizontal projection of the spacecraft simulator and the perpendicular line of the center of mass of the spacecraft simulator; and R is the distance between the geometric center of the spacecraft simulator and the rotation axis of the torsion pendulum platform.

[0083] Specifically, the horizontal projection of the spacecraft simulator is the horizontal projection of the center of mass of the spacecraft to be measured, the perpendicular line is the perpendicular line of the center of mass of the spacecraft to be measured and the rotation center axis of the torsion pendulum platform, and the vibration frequency of the spacecraft simulator can be obtained in multiple ways. In this embodiment, the vibration frequency can be extracted from the time history of the angular displacement of the spacecraft simulator, or the vibration frequency can be directly measured by an instrument, which will not be described here.

[0084] Specifically, the calculation formula of the torsional stiffness coefficient k of the torsion pendulum platform is as follows:

[0085]

[0086] wherein k is the torsional stiffness coefficient of the torsion pendulum platform, r is the distance from the rotation center to the connection of the wire, g is the acceleration of gravity, and L is the length of the wire.

[0087] The center of mass measurement platform 110 and the spacecraft simulator simulate the state of the spacecraft in orbit, and the ground verification experiment of the center of mass identification and the center of mass adjustment is performed, so as to verify the accuracy of the identification and adjustment method, and then the identification and adjustment method can be applied to the actual spacecraft in-orbit measurement, thereby improving the research efficiency and reducing the research cost.

[0088] It should be noted that the spacecraft simulator should be installed on the torsion pendulum platform 150 to perform the ground simulation experiment of the spacecraft center of mass calibration and adjustment. The center of mass of the spacecraft simulator in the initial state should be placed at the position of the rotation center of the torsion pendulum.

[0089] Specifically, in the embodiment of the present application, the method for adjusting the center of mass of the spacecraft simulator is to set the center of mass adjustment mechanism 140, which is installed on the spacecraft simulator. The center of mass adjustment mechanism 140 includes a moving rail 141, a mass block 142 and a driving unit. The mass block 142 is driven to move on the moving rail 141 by the driving unit, so as to change the mass distribution of the spacecraft simulator and the position of the center of mass of the spacecraft simulator.

[0090] According to some embodiments of the present application, the repeated experiments are performed according to the above steps, and the error between the center of mass measurement of the spacecraft simulator and the ideal value is calculated according to the results of multiple experiments, so as to ensure the accuracy of the ground verification experiment.

[0091] According to some embodiments of the present application, the method for calculating the error comprises:

[0092] S1': obtaining the angular velocity and angular acceleration of the spacecraft simulator and the attitude information of the spacecraft simulator;

[0093] S2': setting the Asum matrix, the ω matrix and the dω matrix, the Asum matrix containing the three-dimensional coordinates (x, y, z) of the spacecraft, the ω matrix being the classification of the angular velocity of the spacecraft simulator in the x, y, z directions (ω x , ω y , ω z ), and the dω matrix being the classification of the angular acceleration of the spacecraft simulator in the x, y, z directions (dω x , dω y , dω z );

[0094] S3': adding noise to ω z , dω z , x, and y, and setting the data after adding the noise as ω z’ , dω z’ , x', and y';

[0095] S4': setting a 2x2 matrix A and a 2x1 matrix a, wherein:

[0096]

[0097] a = [x, y]

[0098] S5': calculating the deviation of the spacecraft simulator in the x direction and the deviation in the y direction, and the calculation formula is as follows:

[0099] l = (AxA) -1 xAa

[0100] error = l - [0.005; 0.005]

[0101] Wherein, the first column of the matrix l is the deviation of the spacecraft simulator in the x direction, and the second column of the matrix l is the deviation of the spacecraft simulator in the y direction; the result 2x1 matrix error matrix is the error of the current estimated parameter and the expected value [0.005; 0.005], the first column of the error matrix is the error in the x direction, which is set as x error , and the second column of the error matrix is the error in the y direction, which is set as y error .

[0102] S6': setting the error in the x direction as x error , and setting the error in the y direction as y error ;

[0103] Repeat the experiment and record the error of each time. Let the error of the ith experiment in the x direction and the y direction be x error (i) and y error (i); calculate the sum of squares of the x direction and y direction errors, and calculate the root mean square (RMS) error X error and Y error ; wherein

[0104]

[0105] S7': combine the root mean square errors in the x direction and the y direction to calculate the comprehensive error, the calculation formula is as follows:

[0106]

[0107] L error is the comprehensive error to be solved.

[0108] Through the comprehensive error L error , a reference can be provided for subsequent decision-making to help us judge the accuracy of the centroid estimation, complete the ground verification of the spacecraft on-orbit identification technology, and then apply the centroid identification technology to the on-orbit measurement of the spacecraft.

[0109] The application also discloses a spacecraft centroid on-orbit identification technology ground verification device, please refer to Fig. 2 、 3 , which comprises a centroid measurement platform 110, a driving mechanism 120, an accelerometer 130 and a centroid adjusting mechanism 140, the centroid measurement platform 110 comprises a torsion pendulum platform 150, a torsion pendulum base 160 and a suspension wire, the torsion pendulum platform 150 is used for placing a spacecraft simulator (not shown in the figure), the torsion pendulum platform 150 is arranged above the torsion pendulum base 160, the torsion pendulum platform 150 and the torsion pendulum base 160 are connected through the suspension wire, and only the degree of freedom of rotation around the gravity direction is reserved between the torsion pendulum platform 150 and the torsion pendulum base 160;

[0110] The driving mechanism 120 is arranged on the centroid measurement platform 110 and is used for applying a torque excitation to the spacecraft simulator;

[0111] The accelerometer 130 is arranged on the centroid measurement platform 110 and is used for measuring the acceleration of the spacecraft simulator;

[0112] The centroid adjusting mechanism 140 is arranged on the centroid measurement platform 110 and is used for adjusting the centroid position of the spacecraft simulator.

[0113] The center of mass measuring platform 110 is arranged to measure the center of mass position of the spacecraft simulator, and the center of mass measuring platform 110 is composed of a torsion platform 150, a torsion base 160 and a suspension wire. The torsion platform 150 and the torsion base 160 are connected by the suspension wire, so that the torsion platform 150 can perform a torsion motion on the torsion base 160, and the torsion realized by the suspension wire is less affected by external force and has high motion sensitivity, so that the spacecraft simulator can more accurately simulate the state of the spacecraft in orbit motion. The torsion platform 150 and the torsion base 160 only have a rotational freedom around the gravity direction, so that the spacecraft simulator on the torsion platform 150 will not perform up-and-down swinging and other actions under the influence of gravity, and can maintain a pure ideal rotation to better simulate the motion state of the spacecraft in orbit and to perform more accurate ground verification tests.

[0114] The torsion platform 150 is arranged above the torsion base 160, so that there is no shielding structure above the torsion platform 150, which means that the size of the spacecraft simulator will not affect the overall installation, so that a thinner suspension wire can be selected to improve the precision and sensitivity of the system.

[0115] The driving mechanism 120 is arranged to apply torque excitation to the spacecraft simulator to make the spacecraft simulator enter a torsion state, so as to facilitate subsequent measurement and adjustment. The accelerometer 130 is arranged to collect various motion parameters of the spacecraft simulator in motion, so as to calculate the center of mass position of the spacecraft simulator in motion. The center of mass adjusting mechanism 140 is arranged to adjust the center of mass of the spacecraft simulator according to the center of mass position of the spacecraft simulator, so as to keep a certain deviation between the center of mass of the spacecraft simulator and the measurement center of the accelerometer 130.

[0116] Preferably, the accelerometer is a high-precision quartz flexible accelerometer, which is installed on the moving mechanism and is used to measure the linear acceleration of the measurement center of the accelerometer when the spacecraft simulator performs periodic attitude motion.

[0117] In the embodiment of the present application, the center of mass measuring platform 110 is further provided with an attitude measuring instrument 180. Specifically, the attitude measuring instrument 180 used in the embodiment of the present application is a autocollimator, which has high measurement accuracy and fast measurement speed.

[0118] The autocollimator in the embodiment of the present application has an attitude measurement accuracy of better than 5μrad; and the measurement accuracy of the accelerometer is better than 10 -6 m / s 2 .

[0119] According to some embodiments of the present application, the suspension wires include a first suspension wire 171, a second suspension wire 172, a third suspension wire 173, a fourth suspension wire 174, and a fifth suspension wire 175. The first suspension wire 171 is arranged in a direction parallel to the direction of gravity. The second suspension wire 172 intersects with the third suspension wire 173 at a first intersection point. The fourth suspension wire 174 intersects with the fifth suspension wire 175 at a second intersection point. Both the first intersection point and the second intersection point are arranged on the pendulum platform 150.

[0120] Specifically, the second suspension wire 172 and the third suspension wire 173 form a first pulling structure, and the fourth suspension wire 174 and the fifth suspension wire 175 form a second pulling structure. The first pulling structure and the second pulling structure are mutually symmetrical, with the center of symmetry being a plane formed by the rotation axis of the pendulum platform 150 and the line on which the first suspension wire 171 lies. This structure enables the suspension wires of each pulling structure to provide effective horizontal stiffness, thereby preventing the pendulum plane from producing a swinging motion and maintaining the rotational motion, thus maintaining the stability of the center-of-mass measurement platform 110.

[0121] Working principle: Without limitation, the spacecraft simulator has three rotation angles, and the five-line pendulum structure formed by the five suspension wires limits two rotation degrees of freedom and three translation degrees of freedom, so that the spacecraft simulator to be measured has only the freedom of rotation around the vertical direction, thereby achieving pure ideal rotation at a small angle. In addition, the pendulum platform 150 with the five-line pendulum structure has high flexibility and controllability. It can be adjusted and optimized according to specific requirements to adapt to the rotation requirements of different spacecraft simulators.

[0122] Specifically, in one embodiment of the present application, quartz wires with high elastic modulus and low thermal expansion coefficient are selected as the suspension wires to ensure stable performance under different environmental conditions.

[0123] Specifically, in some other embodiments of the present application, the suspension wires can also be selected from metal wires or suspension wires made of other materials.

[0124] In modern aerospace technology, precise rotation control is very important. Existing technologies use single-axis turntables and single-axis air floating tables to achieve rotation. However, with the continuous development of technology, these solutions cannot fully meet the needs of precise rotation control.

[0125] Single-axis turntables are one of the early rotation solutions used to control the rotation of spacecraft through an independent shaft. However, single-axis turntables have the problem of friction. Since the contact surfaces between the bearings produce friction during rotation, this can negatively affect the stability and accuracy of rotation. Especially at small angles, the friction force will be more obvious, causing instability in rotation.

[0126] On the other hand, the single-axis air floating table adopts the air film support to realize rotation. By forming the air film on the bearing, the contact area between the bearing and the single-axis air floating table can be reduced, thereby reducing the influence of friction. However, the single-axis air floating table also has the problems of air pressure fluctuation and air film unevenness. Since the fluctuation of the air pressure will affect the supporting effect of the air film, the instability of rotation is caused.

[0127] In contrast, the center of mass measuring platform 110 of the embodiment of the application adopts a five-wire torsion pendulum platform and adopts a completely different working principle. It limits the five-directional degrees of freedom, so that the spacecraft simulator to be measured only retains the degree of freedom around the vertical direction. In this way, within a small angle range, the rotation process is more stable, and accurate measurement and control can be realized. At a very small angle, the five-wire torsion pendulum platform can provide pure ideal rotation.

[0128] According to some embodiments of the application, the driving mechanism 120 is a reaction flywheel. The reaction flywheel can use the change of its angular momentum to exert torque excitation on the spacecraft to be measured through the law of conservation of momentum. This driving method has high stability and high response speed, and can quickly and accurately excite the spacecraft to be measured in the experiment.

[0129] According to some embodiments of the application, the center of mass adjusting mechanism 140 includes a moving rail 141, a mass block 142, and a driving unit (not shown in the figure), the moving rail 141 is installed on the spacecraft simulator, the mass block 142 is movably arranged on the moving rail 141, and the driving unit is used to drive the mass block 142 to move. By changing the position of the mass block 142, the mass distribution on the spacecraft simulator can be changed, thereby adjusting the center of mass position of the spacecraft simulator.

[0130] In the embodiment of the application, the center of mass adjusting mechanism 140 can be provided with multiple.

[0131] Preferably, according to some embodiments of the application, the center of mass adjusting mechanism 140 is provided with two, and the mass blocks of the two center of mass adjusting mechanisms 140 are orthogonal in the moving direction, so as to adjust the center of mass positions of the spacecraft simulator in two directions.

[0132] According to some embodiments of the application, the torsion pendulum platform 150 includes a connecting part 151, the connecting part 151 has a bending part 152, the bending part 152 is arranged at the bottom of the torsion pendulum base 160, a support wire 176 is arranged on the bending part 152, the support wire 176 is connected with the torsion pendulum base 160, and the support wire 176 is arranged in a direction parallel to the direction of gravity.

[0133] The embodiment of the present application, by the arrangement of the support wire 176, in combination with the five-wire torsional pendulum structure formed by the five suspension wires, forms a tensioned integral mechanism that can effectively balance the tension, further enhancing the balance of the integral mechanism. In addition, the support wire 176, through the connection with the other five suspension wires, can maintain the stability of the entire structure, ensuring the reliability and durability during the operation process.

[0134] When the gravity is collinear with the support wire 176, the tensioned integral structure is best balanced, and the stability of the structure can be maintained through the connection of the five suspension wires.

[0135] The embodiment of the present application utilizes the pre-stress of the suspension wire and the support wire 176 to maintain the shape and stability of the structure. When subjected to external forces, the suspension wire and the support wire 176 can absorb external forces and distribute these forces evenly, thereby reducing the pressure of external forces on individual points in the structure. This design effectively isolates external vibrations and reduces the impact on the structure. The natural vibration frequency of the tensioned structure is generally low, meeting the requirements of ground verification. The structure not only enhances the stability, but also improves the anti-seismic performance.

[0136] According to some embodiments of the present application, an acceleration meter 130 is connected with a data acquisition card, which is used to collect the output parameters of the acceleration meter 130 for subsequent calculation.

[0137] According to some embodiments of the present application, a moving mechanism 131 is connected with the acceleration meter 130, which is used to move the position of the acceleration meter 130. By setting the moving mechanism 131, the position of the acceleration meter 130 can be adjusted, so as to adjust the deviation between the measurement center of the acceleration meter 130 and the center of mass of the spacecraft simulator. The moving mechanism 131 includes a first moving assembly 132 and a second moving assembly 133, the acceleration meter 130 is connected to the first moving assembly 132, the first moving assembly 132 can drive the acceleration meter 130 to move, the first moving assembly 132 is connected to the second moving assembly 133, the second moving assembly 133 can drive the first moving assembly 132 to move, by setting the first moving assembly 132 and the second moving assembly 133, the acceleration meter 130 can realize movement in two directions.

[0138] The moving mechanism 131 generates the deviation between the center of mass of the spacecraft simulator and the measurement center of the acceleration meter 130, the driving mechanism (reaction flywheel) drives the spacecraft simulator on the five-wire torsional pendulum platform to generate periodic attitude driving, the acceleration meter generates the linear acceleration of the measurement center of the acceleration meter, the data acquisition card 181 collects and transmits to the upper computer, calibrates the distance between the acceleration meter center and the center of mass of the spacecraft simulator through the center of mass calibration algorithm, and then inputs the motion instruction into the center of mass adjusting mechanism according to the calibration result to adjust the center of mass of the spacecraft simulator, so as to control the deviation between the center of mass of the spacecraft simulator and the center of the acceleration meter within a certain range.

[0139] The on-orbit spacecraft centroid recognition technology ground verification device of the embodiment of the application further has a power supply 190 for providing energy required for system operation.

[0140] The embodiment of the application is described in detail above in combination with the drawings, but the application is not limited to the above-mentioned embodiments, and various changes can be made within the knowledge range of ordinary skilled in the art without departing from the purpose of the application. In addition, the embodiments of the application and the features in the embodiments can be combined with each other without conflict.

Claims

1. A ground verification method for on-orbit identification technology of spacecraft center of mass, characterized in that, Includes the following steps: A center of mass measurement platform is set up, and the degrees of freedom of the center of mass measurement platform are restricted so that the center of mass measurement platform retains only the degree of freedom of rotation about the direction of gravity; The spacecraft simulator is placed on the center of mass measurement platform, and the center of mass position of the spacecraft simulator is measured using a center of mass inertia measuring instrument. An accelerometer and an attitude measuring instrument are set above the center of mass measurement platform. The accelerometer has a measurement center, and the center of mass position of the spacecraft simulator is offset from the position of the measurement center. A torque excitation is applied to the spacecraft simulator, causing the spacecraft simulator to oscillate periodically on the center of mass measurement platform; The attitude measurement instrument is used to measure the attitude of the spacecraft simulator in a swinging motion, and the accelerometer is used to measure the linear acceleration of the spacecraft simulator located at the center of the accelerometer. The position of the spacecraft simulator's center of mass is calculated by measuring the attitude and acceleration of the spacecraft simulator. Based on the calculated center of mass position, the center of mass position of the spacecraft simulator is adjusted to control the distance between the center of mass of the spacecraft simulator and the center of the accelerometer. Repeat the experiment, and calculate the error between the spacecraft simulator's center of mass measurement and the ideal value based on the results of the repeated experiment to ensure the accuracy of the ground verification experiment; Methods for calculating errors include: Obtain the angular velocity and angular acceleration of the spacecraft simulator. Set the Asum matrix, Matrix and The matrix, the Asum matrix, contains the three-dimensional coordinates (x, y, z) of the spacecraft. The matrix represents the classification of the angular velocity of the spacecraft simulator in the x, y, and z directions. ); the The matrix represents the classification of angular acceleration in the x, y, and z directions of the spacecraft simulator. ); exist , Add noise to x, y, and set the data after adding noise as follows: , Set one Matrix A and a Matrix a, where: The following formulas are used to calculate the deviations in the x and y directions of the spacecraft simulator: error=l-[0.005;0.005] in, The first column represents the deviation of the spacecraft simulator in the x-direction. The second column represents the deviation of the spacecraft simulator in the y-direction; The error matrix represents the error between the current estimated parameter and the expected value [0.005; 0.005]. The first column of the error matrix represents the error in the x-direction, denoted as […]. The second column of the error matrix represents the error in the y-direction, denoted as . ; Repeat the experiment and record the error each time; let the error in the x-direction and the error in the y-direction of the i-th experiment be respectively... (i) and (i); Calculate the sum of squares of the errors in the X and Y directions, and obtain the root mean square (RMS) error. and ;in Combining the root mean square errors in the X and Y directions, the overall error is calculated using the following formula: This is the comprehensive error we are looking for.

2. The ground verification method for on-orbit identification technology of spacecraft center of mass according to claim 1, characterized in that, The center of mass measurement platform includes a torsion pendulum base and a torsion pendulum platform. The formula for measuring the center of mass position of the spacecraft simulator is as follows: In the formula denoted as the offset of the spacecraft simulator's center of mass; k is the torsional stiffness coefficient of the torsional platform. The mass of the spacecraft simulator; The mass of the torsional platform; and These are the vibration frequencies of the spacecraft simulator when it rotates 0° and 180° around the axis of the torsion platform, respectively. R is the angle between the horizontal projection of the spacecraft simulator and the perpendicular line of the spacecraft simulator; R is the distance between the geometric center of the spacecraft simulator and the axis of rotation of the torsion platform.

3. A ground verification device for on-orbit identification technology of spacecraft center of mass, characterized in that, The ground verification method for implementing the spacecraft center of mass identification technology in orbit as described in claim 1 includes: A center of mass measurement platform, comprising a torsion pendulum platform, a torsion pendulum base, and a suspension wire, wherein the torsion pendulum platform is used to place a spacecraft simulator, the torsion pendulum platform is positioned above the torsion pendulum base, the torsion pendulum platform and the torsion pendulum base are connected by the suspension wire, and the torsion pendulum platform and the torsion pendulum base retain only the degree of freedom of rotation about the direction of gravity. A drive mechanism, mounted on the center of mass measurement platform, is used to apply torque excitation to the spacecraft simulator; An accelerometer, which is mounted above the center of mass measurement platform, is used to measure the acceleration of the spacecraft simulator; A center of mass adjustment mechanism is installed on the center of mass measurement platform and is used to adjust the position of the center of mass of the spacecraft simulator.

4. The ground verification device for spacecraft center of mass identification technology in orbit according to claim 3, characterized in that, The suspension wires include a first suspension wire, a second suspension wire, a third suspension wire, a fourth suspension wire, and a fifth suspension wire. The first suspension wire is arranged in a direction parallel to the direction of gravity. The second suspension wire intersects the third suspension wire at a first intersection point, and the fourth suspension wire intersects the fifth suspension wire at a second intersection point. Both the first intersection point and the second intersection point are located on the torsion platform.

5. The ground verification device for spacecraft center of mass identification technology in orbit according to claim 3, characterized in that, The drive mechanism is a reaction flywheel.

6. The ground verification device for spacecraft center of mass identification technology in orbit according to claim 3, characterized in that, The center of mass adjustment mechanism includes a moving rail, a mass block, and a drive unit. The moving rail is mounted on the spacecraft simulator, the mass block is movably mounted on the moving rail, and the drive unit is used to drive the mass block to move.

7. The ground verification device for spacecraft center of mass identification technology in orbit according to claim 3, characterized in that, The torsion platform includes a connecting part, which has a bending part. The bending part is located at the bottom of the torsion base, and a support line is provided on the bending part. The support line is connected to the torsion base, and the direction of the support line is parallel to the direction of gravity.

8. The ground verification device for spacecraft center of mass identification technology in orbit according to claim 3, characterized in that, The accelerometer is connected to a moving mechanism, which is used to move the position of the accelerometer.

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