Spacecraft primary mirror molded surface high-precision ground optical test system and method

Abstract

The invention discloses a high-precision ground optical testing system and method for a spacecraft primary mirror profile, and relates to the technical field of spacecraft optical testing, and the system comprises a truss assembly, a primary mirror, a sensing assembly, a vibration exciter, a demodulator and a computer. The vibration exciter is used for loading exciting force to the truss assembly; the sensing assembly is used for collecting strain signals of the truss assembly and the primary mirror. The demodulator is used for demodulating the truss assembly strain signal and the primary mirror strain signal respectively; and the computer is used for carrying out displacement field inversion on the truss assembly strain demodulation signal by adopting a Ko displacement theory algorithm, carrying out displacement field inversion on the primary mirror strain demodulation signal by adopting an inverse finite element algorithm, and superposing a truss assembly inversion result and a primary mirror inversion result to obtain a form reconstruction result, so that form reconstruction of the spacecraft primary mirror structure is realized. According to the invention, the accuracy of displacement field measurement can be improved, and the structural form reconstruction of the spacecraft primary mirror can be accurately realized.

Description

Technical Field

[0001] The present application relates to the technical field of spacecraft optical testing, and in particular to a high-precision ground-based optical testing system and method for a spacecraft primary mirror surface. Background Art

[0002] Under the influence of multiple complex environmental factors such as extreme temperature changes in space, microgravity, microvibration, and fatigue, in order to ensure the detection accuracy and image positioning and registration capabilities of ultra-large-scale complex flexible spacecraft, the deformation that the spacecraft needs to monitor spans from the millimeter level to the nanometer level in order to effectively implement feedback control. However, existing measurement systems and measurement methods generally have the problem of low accuracy in displacement field measurement, and cannot accurately achieve the reconstruction of the structure and morphology of the spacecraft primary mirror. Summary of the invention

[0003] The purpose of this application is to provide a high-precision ground-based optical testing system and method for the surface of a spacecraft primary mirror, which can improve the accuracy of displacement field measurement and accurately achieve structural morphology reconstruction of the spacecraft primary mirror.

[0004] To achieve the above objectives, this application provides the following solutions:

[0005] In a first aspect, the present application provides a high-precision ground optical test system for a spacecraft primary mirror surface, the high-precision ground optical test system for a spacecraft primary mirror surface comprising a truss assembly, a primary mirror, a sensor assembly, an exciter, a demodulator, and a computer;

[0006] The primary mirror is disposed at one end of the truss assembly, the exciter is disposed at the other end of the truss assembly, the sensor assembly is disposed at each measuring point of the truss assembly and each measuring point on the surface of the primary mirror, the output end of the sensor assembly is connected to the input end of the demodulator, and the output end of the demodulator is connected to the input end of the computer;

[0007] The truss assembly and the primary mirror are used as a force-bearing component and a core observation component of the spacecraft respectively;

[0008] The exciter is used to apply an excitation force to the truss assembly to cause the truss assembly and the primary mirror to vibrate;

[0009] The sensing component is used to collect strain signals of the truss component and the primary mirror respectively, to obtain a truss component strain signal and a primary mirror strain signal;

[0010] The demodulator is used to demodulate the truss assembly strain signal and the primary mirror strain signal respectively to obtain a truss assembly strain demodulation signal and a primary mirror strain demodulation signal;

[0011] The computer is used to perform displacement field inversion on the strain demodulation signal of the truss assembly using the Ko displacement theory algorithm to obtain a truss assembly inversion result, and simultaneously perform displacement field inversion on the strain demodulation signal of the primary mirror using the inverse finite element algorithm to obtain a primary mirror inversion result, and superimpose the truss assembly inversion result and the primary mirror inversion result to obtain a morphological reconstruction result, thereby realizing morphological reconstruction of the spacecraft primary mirror structure.

[0012] Optionally, the truss assembly includes three trusses, one end of the three trusses is fixed to the same point, and the other ends are evenly dispersed and fixed to the surface of the primary mirror.

[0013] Optionally, the sensing component includes a first sensing component, the first sensing component includes a plurality of first fiber grating strain sensors, the plurality of first fiber grating strain sensors are evenly arranged at various measuring points on the surface of the primary mirror, and the first fiber grating strain sensors are used to collect the strain signal of the primary mirror.

[0014] Optionally, the number of the first fiber grating strain sensors at each measuring point on the surface of the primary mirror is 3, one end of the three first fiber grating strain sensors is fixed at the same point, and the other ends are arranged along the horizontal directions of 0°, 45° and 90° to form a strain rosette structure.

[0015] Optionally, the sensing component also includes a second sensing component, which includes a plurality of second fiber grating strain sensors, and the plurality of second fiber grating strain sensors are evenly arranged at various measuring points on the surface of the truss component, and the second fiber grating strain sensors are used to collect strain signals of the truss component.

[0016] Optionally, the sensing assembly further comprises a third sensing assembly, wherein the third sensing assembly comprises a plurality of first fiber Bragg grating temperature compensation sensors and a plurality of second fiber Bragg grating temperature compensation sensors;

[0017] A first fiber Bragg grating temperature compensation sensor is arranged at each measuring point on the surface of the primary mirror, and the first fiber Bragg grating temperature compensation sensor is used to perform temperature compensation on the first fiber Bragg grating strain sensor;

[0018] A second fiber Bragg grating temperature compensation sensor is arranged at each measuring point on the surface of the truss assembly, and the second fiber Bragg grating temperature compensation sensor is used to perform temperature compensation on the second fiber Bragg grating strain sensor.

[0019] Optionally, the computer includes a user graphical interface, which includes operation buttons for initialization, data acquisition, data preprocessing, truss assembly displacement field inversion, primary mirror displacement field inversion and morphological reconstruction, and the user graphical interface is used to operate each of the operation buttons and display the truss assembly inversion results, the primary mirror inversion results and the morphological reconstruction results in real time.

[0020] In a second aspect, the present application proposes a high-precision ground optical test method for a spacecraft primary mirror surface, the high-precision ground optical test method for a spacecraft primary mirror surface comprising the following steps:

[0021] Acquire a truss assembly strain signal and a primary mirror strain signal; the truss assembly strain signal and the primary mirror strain signal are the truss assembly strain signal and the primary mirror strain signal respectively collected by the spacecraft primary mirror profile high-precision ground optical testing system described in the first aspect;

[0022] Demodulating the truss assembly strain signal and the primary mirror strain signal respectively to obtain a truss assembly strain demodulation signal and a primary mirror strain demodulation signal;

[0023] The Ko displacement theory algorithm is used to perform displacement field inversion on the strain demodulation signal of the truss component to obtain the inversion result of the truss component, and the inverse finite element algorithm is used to perform displacement field inversion on the strain demodulation signal of the primary mirror to obtain the inversion result of the primary mirror;

[0024] The truss assembly inversion result and the primary mirror inversion result are superimposed to obtain a morphological reconstruction result, thereby realizing the morphological reconstruction of the spacecraft primary mirror structure.

[0025] Optionally, before the step of respectively demodulating the truss assembly strain signal and the primary mirror strain signal to obtain a truss assembly strain demodulation signal and a primary mirror strain demodulation signal, the spacecraft primary mirror profile high-precision ground optical testing method further includes the following steps:

[0026] The truss assembly strain signal and the primary mirror strain signal are denoised respectively to obtain a preprocessed truss assembly strain signal and a preprocessed primary mirror strain signal; the preprocessed truss assembly strain signal and the preprocessed primary mirror strain signal are used for demodulation as the truss assembly strain signal and the primary mirror strain signal respectively to obtain the truss assembly strain demodulation signal and the primary mirror strain demodulation signal.

[0027] Optionally, after the step of respectively demodulating the truss assembly strain signal and the primary mirror strain signal to obtain a truss assembly strain demodulation signal and a primary mirror strain demodulation signal, the spacecraft primary mirror profile high-precision ground optical testing method further comprises the following steps:

[0028] An IP address is set, and the truss assembly strain demodulation signal and the primary mirror strain demodulation signal are imported into computer software using a UDP transmission protocol, and displacement field inversion and morphological reconstruction are performed in a user graphical interface of the computer software.

[0029] According to the specific embodiments provided in this application, this application discloses the following technical effects:

[0030] The present application provides a high-precision ground optical test system and method for the surface of a spacecraft primary mirror. An excitation force is applied to a truss assembly by an exciter, strain signals of the truss assembly and the primary mirror are respectively collected by a sensor assembly, the strain signals of the truss assembly and the primary mirror are respectively demodulated by a demodulator, and then a Ko displacement theory algorithm is used by a computer to perform displacement field inversion on the strain demodulated signal of the truss assembly, an inverse finite element algorithm is used to perform displacement field inversion on the strain demodulated signal of the primary mirror, and the inversion result of the truss assembly and the inversion result of the primary mirror are superimposed to obtain a morphological reconstruction result, thereby realizing morphological reconstruction of the spacecraft primary mirror structure. The present application collects and processes signals based on sensor components, exciters, demodulators and computers, and applies the Ko displacement theory algorithm and the inverse finite element algorithm to perform displacement field inversion and morphological reconstruction on the strain demodulation signal of the truss component and the strain demodulation signal of the primary mirror, respectively. The structural displacement field can be reversely deduced through strain without the need for various mechanical parameters of the structure, and there is no complicated calculation process and calculation error of various mechanical parameters. The accuracy of displacement field inversion and reconstruction can be effectively improved, the accuracy of displacement field measurement can be improved, and the morphological reconstruction of the spacecraft primary mirror structure can be accurately realized. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] In order to more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0032] Figure 1 A schematic structural diagram of a high-precision ground-based optical testing system for a spacecraft primary mirror surface provided in Example 1 of the present application.

[0033] Figure 2 A schematic diagram of the structure of a four-node reverse unit provided in Example 1 of the present application.

[0034] Figure 3 Schematic diagram of the local coordinate system on the neutral plane of the four-node inverse unit provided in Example 1 of the present application.

[0035] Figure 4 Schematic diagram of the strain rosette pasting provided in Example 1 of the present application.

[0036] Figure 5 Schematic diagram of strain measurement points provided in Example 1 of the present application.

[0037] Figure 6 A schematic flow chart of a method for high-precision ground-based optical testing of a spacecraft primary mirror surface provided in Example 2 of the present application.

[0038] Reference numerals:

[0039] 1-truss assembly; 2-primary mirror; 3-fiber Bragg grating strain sensor; 4-fiber Bragg grating temperature compensation sensor; 5-demodulator; 6-exciter; 7-computer; 8-network cable; 9-fiber jumper. DETAILED DESCRIPTION

[0040] The following will be combined with the drawings in the embodiments of the present application to clearly and completely describe the technical solutions in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of this application.

[0041] In order to make the above-mentioned objects, features and advantages of the present application more obvious and easy to understand, the present application is further described in detail below in conjunction with the accompanying drawings and specific implementation methods.

[0042] Example 1

[0043] like Figure 1 As shown, this embodiment provides a high-precision ground optical test system for the profile of a spacecraft primary mirror, which mainly includes a truss assembly 1, a primary mirror 2, a sensor assembly, an exciter 6, a demodulator 5, a computer 7, etc. Among them, one end of the truss assembly 1 is provided with the primary mirror 2, the other end of the truss assembly 1 is provided with the exciter 6, the sensor assembly is respectively arranged at each measuring point position of the truss assembly 1 and each measuring point position on the surface of the primary mirror 2, the output end of the sensor assembly is connected to the input end of the demodulator 5, and the output end of the demodulator 5 is connected to the input end of the computer 7.

[0044] The truss assembly 1 and the primary mirror 2 are used as a force-bearing component and a core observation component of the spacecraft, respectively.

[0045] The exciter 6 is used to apply an excitation force to the truss assembly 1 to cause the truss assembly 1 and the primary mirror 2 to vibrate.

[0046] The sensor assembly is used to collect strain signals of the truss assembly 1 and the primary mirror 2 respectively, to obtain the strain signal of the truss assembly 1 and the strain signal of the primary mirror 2.

[0047] The demodulator 5 is used to demodulate the strain signal of the truss assembly 1 and the strain signal of the primary mirror 2 respectively to obtain a strain demodulation signal of the truss assembly 1 and a strain demodulation signal of the primary mirror 2.

[0048] The computer 7 is used to perform displacement field inversion on the strain demodulation signal of the truss assembly 1 using the Ko displacement theory algorithm to obtain the inversion result of the truss assembly 1, and at the same time, perform displacement field inversion on the strain demodulation signal of the primary mirror 2 using the inverse finite element algorithm to obtain the inversion result of the primary mirror 2, and superimpose the inversion result of the truss assembly 1 and the inversion result of the primary mirror 2 to obtain a morphological reconstruction result, thereby realizing the morphological reconstruction of the spacecraft primary mirror 2 structure.

[0049] In this embodiment, the truss assembly 1 includes three trusses, one end of the three trusses is fixed at the same point, and the other ends are evenly distributed and fixed to the surface of the primary mirror 2.

[0050] In this embodiment, the sensing component includes a first sensing component, which includes a plurality of first fiber grating strain sensors, which are evenly arranged at various measuring points on the surface of the primary mirror 2, and the first fiber grating strain sensors are used to collect the strain signal of the primary mirror 2.

[0051] In this embodiment, the number of the first fiber Bragg grating strain sensors at each measuring point on the surface of the primary mirror 2 is 3, one end of the three first fiber Bragg grating strain sensors is fixed at the same point, and the other ends are arranged along the horizontal directions of 0°, 45° and 90° to form a strain rosette structure.

[0052] In this embodiment, the sensing component also includes a second sensing component, which includes a plurality of second fiber grating strain sensors, and the plurality of second fiber grating strain sensors are evenly arranged at various measuring points on the surface of the truss component 1, and the second fiber grating strain sensor is used to collect the strain signal of the truss component 1.

[0053] In this embodiment, the sensing assembly further includes a third sensing assembly, and the third sensing assembly includes a plurality of first fiber Bragg grating temperature compensation sensors and a plurality of second fiber Bragg grating temperature compensation sensors.

[0054] A first fiber Bragg grating temperature compensation sensor is arranged at each measuring point on the surface of the primary mirror 2 , and the first fiber Bragg grating temperature compensation sensor is used to perform temperature compensation on the first fiber Bragg grating strain sensor.

[0055] A second fiber Bragg grating temperature compensation sensor is arranged at each measuring point on the surface of the truss assembly 1 , and the second fiber Bragg grating temperature compensation sensor is used to perform temperature compensation on the second fiber Bragg grating strain sensor.

[0056] In this embodiment, the structure of the spacecraft primary mirror 2 of the experimental model is a structure of a truss assembly 1, the primary mirror 2 is fixed by three trusses, and the overall structure is that the ends of the three trusses suspend and fix the primary mirror 2. A plurality of first fiber Bragg grating strain sensors are evenly pasted on the upper left, lower left, upper right and lower right positions of the primary mirror 2, and a plurality of second fiber Bragg grating strain sensors are arranged at each measuring point position on the three trusses. The fiber Bragg grating strain sensor 3 network formed by the first fiber Bragg grating strain sensor and the second fiber Bragg grating strain sensor constitutes a basic sensing system for the morphological perception of the spacecraft primary mirror 2.

[0057] In this embodiment, the first fiber grating strain sensor forms a relatively independent sensing unit at each measuring point position on the surface of the primary mirror 2. Each sensing unit at each measuring point position includes 4 sensors, of which 3 sensors are arranged at angles of 0°, 45°, and 90° respectively in the horizontal direction, and the 3 sensors are fixed at the same point at one end, thereby forming a strain flower structure. The strain flower is fixed at the sensor installation position on the surface of the primary mirror 2 using AB epoxy glue to monitor the strain changes at various angles of the surface of the spacecraft primary mirror 2. The other sensor is the first fiber grating temperature compensation sensor, which is used to eliminate the influence of wavelength deviation caused by temperature changes on the first fiber grating strain sensor and improve the accuracy of strain monitoring. The first fiber grating temperature compensation sensor is fixed at the position of the strain flower using ordinary fixing methods such as pasting, but it does not contact the surface of the primary mirror 2. The strain flower is pasted as Figure 4 The strain measurement points are shown in Figure 5 shown.

[0058] In this embodiment, the first fiber Bragg grating temperature compensation sensor is fixed at both ends near each first fiber Bragg grating strain sensor and is not attached to the surface of the structure, so as to eliminate the influence of the first fiber Bragg grating strain sensor itself on strain due to its temperature sensitivity.

[0059] In this embodiment, the second fiber Bragg grating strain sensor forms a relatively independent sensing unit at each measuring point of the truss assembly 1. Each sensing unit at each measuring point includes two sensors, one of which is fixed at the sensor installation position of the truss assembly 1 using AB epoxy glue to monitor strain at various angles. The other sensor is a second fiber Bragg grating temperature compensation sensor, which is fixed at the second fiber Bragg grating strain sensor position using common fixing methods such as pasting, but does not contact the surface of the truss assembly 1.

[0060] In this embodiment, the computer 7 includes a user graphical interface, which includes operation buttons for initialization, data acquisition, data preprocessing, truss component 1 displacement field inversion, primary mirror 2 displacement field inversion and morphological reconstruction. The user graphical interface is used to operate each of the operation buttons and display the inversion results of the truss component 1, the inversion results of the primary mirror 2 and the morphological reconstruction results in real time.

[0061] In this embodiment, the sampling frequency of the demodulator 5 is 1000 Hz.

[0062] The system of this embodiment mainly consists of a truss assembly 1 (three trusses), a primary mirror 2, a plurality of fiber Bragg grating strain sensors 3 (including a first fiber Bragg grating strain sensor and a second fiber Bragg grating strain sensor), a plurality of fiber Bragg grating temperature compensation sensors 4 (including a first fiber Bragg grating temperature compensation sensor and a second fiber Bragg grating temperature compensation sensor), a demodulator 5, an exciter 6, a computer 7, a network cable 8 and a plurality of fiber jumpers 9. Among them, a plurality of fiber Bragg grating strain sensors 3 are fixed on the three trusses, the strain flower is fixed on the surface of the primary mirror 2, and a fiber Bragg grating temperature compensation sensor 4 is fixed near each measuring point where the strain flower is arranged. The fiber Bragg grating strain sensor 3 is connected to the demodulator 5 through the fiber jumper, and the demodulator 5 converts the Bragg wavelength signal into an electrical signal and sends it to the computer 7 through the network cable 8. The computer 7 decodes and communicates according to the transmission protocol, transmits the strain signal to the user graphical interface, performs data processing and displacement reconstruction on the strain signal through the interactive user graphical interface, completes the high-precision optical measurement of the primary mirror 2 of the spacecraft, and performs visual display on the user graphical interface.

[0063] In this embodiment, an excitation force is applied to the truss assembly 1 by means of a vibrator 6, strain signals of the truss assembly 1 and the primary mirror 2 are respectively collected by means of a sensor assembly, the strain signals of the truss assembly 1 and the primary mirror 2 are respectively demodulated by means of a demodulator 5, and then a computer 7 is used to perform displacement field inversion on the strain demodulated signal of the truss assembly 1 by means of a Ko displacement theory algorithm, and to perform displacement field inversion on the strain demodulated signal of the primary mirror 2 by means of an inverse finite element algorithm, and the inversion results of the truss assembly 1 and the primary mirror 2 are superimposed to obtain a morphological reconstruction result, thereby realizing morphological reconstruction of the structure of the primary mirror 2 of the spacecraft. The present application collects and processes signals based on a sensor assembly, an exciter 6, a demodulator 5 and a computer 7, and applies the Ko displacement theory algorithm and the inverse finite element algorithm to perform displacement field inversion and morphological reconstruction on the strain demodulation signal of the truss assembly 1 and the strain demodulation signal of the primary mirror 2, respectively. The structural displacement field can be reversely deduced through strain without the need for various mechanical parameters of the structure, and there is no complicated calculation process and calculation error of various mechanical parameters. The accuracy of displacement field inversion and reconstruction can be effectively improved, the accuracy of displacement field measurement can be improved, and the morphological reconstruction of the spacecraft primary mirror 2 structure can be accurately realized.

[0064] In this embodiment, a compilation and decoding method based on UDP (User Datagram Protocol) is adopted. The method is determined by the internal hardware output signal of the demodulator 5, and the decoding steps are as follows:

[0065] (1) Set the target IP address and port of computer 7.

[0066] (2) Open the target port.

[0067] (3) Read in an 8-bit unsigned integer.

[0068] (4) Convert the above 8-bit unsigned integer into a hexadecimal number.

[0069] (5) Every three hexadecimal numbers are integrated into one number, and the data are merged by connecting the head and the tail.

[0070] (6) Convert the combined hexadecimal number into a decimal number and multiply it by 0.1 to obtain the frequency.

[0071] (7) The wavelength data of the signal is calculated based on the speed of light calculation formula.

[0072] In this embodiment, the user graphical interface is a high-precision optical measurement platform for the spacecraft primary mirror 2 under micro-vibration interference, which is a software system for strain signal processing and displacement reconstruction of the spacecraft primary mirror 2 installed in the computer 7. This embodiment uses the platform to perform morphological perception of the spacecraft primary mirror 2, and the operation method and steps are as follows:

[0073] (1) A strain rosette consisting of three fiber Bragg grating strain sensors 3 is fixed at each measuring point on the surface of the spacecraft primary mirror 2 with AB epoxy glue, and a fiber Bragg grating temperature compensation sensor 4 is fixed at each strain rosette position to measure temperature; a fiber Bragg grating strain sensor 3 is fixed at each measuring point on the spacecraft truss with AB epoxy glue, and a fiber Bragg grating temperature compensation sensor 4 is fixed at each strain sensor position to measure temperature and temperature compensation.

[0074] (2) Use the exciter 6 to load the truss assembly 1 and the primary mirror 2 of the spacecraft to cause them to deform.

[0075] (3) The optical fiber jumpers connected to the optical fibers are connected to the demodulator 5 according to different channels. The demodulator 5 is connected to the computer 7 via the network cable 8 to transmit the data to the computer.

[0076] (4) Set the IP address of computer 7 and input the digital signal into the computer 7 software according to the UDP transmission protocol.

[0077] (5) In the user graphical interface, select the operation button on the user graphical interface to implement various functions and display the operation results on the user graphical interface.

[0078] In this embodiment, the user graphical interface includes the following functions:

[0079] (1) Data initialization, to enable the computer 7 to receive the strain signal of the spacecraft primary mirror 2. It is necessary to convert the wavelength data into strain data, and determine whether the absolute value of the strain signal exceeds the preset amplitude. If it exceeds the preset amplitude, the abnormal value at this point is removed and the value is supplemented by the interpolation method; if it does not exceed the preset amplitude, the subsequent steps are performed. The temperature compensation is subtracted from the processed strain data.

[0080] (2) Data preprocessing, in which the computer 7 performs denoising on the strain signal and determines whether the absolute value of the amplitude of the strain signal exceeds a preset amplitude.

[0081] First, the strain signal wavelet packet decomposition uses the db4 wavelet function (a Dobesi wavelet function), and the strain signal denoising wavelet decomposition is defined as follows:

[0082] For functions If its Fourier transform Satisfies the admissibility condition, that is, the constant of the wavelet transform It is expressed as:

[0083]

[0084] Where ω is the angular frequency of the strain signal, is a basic wavelet or wavelet mother function, and the wavelet mother function Scaling and translation are performed, and the scale factor is set to a, the translation factor is set to b, and the function after translation and scaling is set to Then we have:

[0085]

[0086] Any L 2 The function f(t) in (R) space is expanded under the wavelet basis. This expansion is called the continuous wavelet transform of the function f(t), and the expression is:

[0087]

[0088] From the above formula, we can see that the wavelet transform Wf(a,b) is a function of the scale factor a and the translation factor b. The signal is analyzed by scaling and translating in the spatial domain (time domain).

[0089] In this embodiment, the corresponding inverse transformation is expressed as:

[0090]

[0091] From a mathematical perspective, wavelet transform is to transform a function (or signal) into In practical applications, the scale of wavelet transform does not need to be continuous, but the continuous wavelet and its transform are discretized in some way. Usually, the scale factor a is discretized according to the power series, and the translation factor b is uniformly discrete. Let (m,n∈Z,a0≠0), where m and n are discretization parameters and Z is an integer set. Then formula (3) can be transformed into:

[0092]

[0093] In this embodiment, the wavelet decomposition decomposes the low-frequency part layer by layer, and the resolution increases with the increase of the decomposition level, while the high-frequency part is not considered.

[0094] (3) The Ko displacement theory algorithm is used to invert the displacement field of the truss structure strain signal.

[0095] The principle of the Ko displacement theory algorithm in this embodiment is as follows:

[0096] Ko displacement theory is based on the classical beam theory of material mechanics, using piecewise linear assumptions and differential means to transform the deformation problem of the structure into a summation problem. In material mechanics, according to the classical beam theory of Euler-Bernoulli, the structure is subjected to bending loads. On the cross section of a beam with a uniform cross-section, the calculation formula for the bending normal stress is given by the bending theory of the beam. For a simply supported beam, at the top and bottom of the beam, the normal stress in the cross section can be calculated by the following formula:

[0097]

[0098] Among them, σ(x) represents stress, M(x) is the bending moment of the beam at that position, c is the distance from a point on the upper surface of the beam to the neutral plane, also known as the cross-sectional eccentricity, and I is the moment of inertia of the cross section. Combined with the stress-strain relationship Where ε(x) is strain and E is elastic modulus. The relationship between surface strain and bending moment of a beam structure with uniform cross-section can be deduced:

[0099]

[0100] In this embodiment, the approximate differential equation of the deflection curve of the cantilever beam structure with arbitrary cross-section is Substituting into the above formula, we can get the relationship between surface strain and bending deflection y:

[0101]

[0102] Integrating the above equation, we can get the slope equation of the structure at position x, tanθ(x), where tanθ i-1 The slope value at the point i-1 before the current point i is:

[0103]

[0104] The displacement equation y(x) at this position can be integrated again by the formula:

[0105]

[0106] The truss of this embodiment can be made of uniform steel long rods, and its c(x) value can be considered unchanged. The Ko displacement theory algorithm equation of the uniform cross-section beam is:

[0107]

[0108] Among them, Δl is the distance from point i-1 to point i, ε i , ε i-1 are the strain values ​​at point i and point i-1 respectively.

[0109] (1) The inverse finite element algorithm is used to invert the displacement field of the strain signal of the primary mirror 2, and the morphological reconstruction and visualization of the structure of the spacecraft primary mirror 2 are performed.

[0110] In this embodiment, the principle of the inverse finite element algorithm is as follows:

[0111] Figure 2 is a four-node inverse unit (referred to as unit) consisting of four nodes. Figure 3The local coordinates xyz are established on the neutral plane of a four-node inverse unit, and the origin of the local coordinate system is at the centroid position in the plane of the four-node inverse unit. There are two degrees of freedom in each coordinate axis direction. u, v and w represent the translation along the three coordinate axes x, y, and z, respectively, and θ xi ,θ yi ,θ zi They represent the rotation angles around the x-axis, y-axis, and z-axis respectively. The thickness of the four-node inverse element is 2h, and the four nodes P1, P2, P3, and P4 are located on the same plane.

[0112] The deformation of any point on the structure includes displacements in three directions. According to the first-order shear deformation theory, the following formula can be used:

[0113]

[0114] In this embodiment, each node of the spatial beam unit has six degrees of freedom, which are displacement u0, v0, w0 and rotation component around the coordinate axis. For the deformation of any point in the structure in three directions, the displacement in the z direction does not change along the thickness direction. According to the finite element theory, the displacement of any point in the unit can be expressed by the linear interpolation of the shape function of the four nodes in the unit. The formula is as follows:

[0115]

[0116] Among them, L i 、M i 、N i is the shape function of the unit, which is a function that characterizes the shape characteristics within the unit. i 、M i 、N i Related to the unit size and node position coordinates; i represents the number of the unit node, i = 1, 2, 3, 4.

[0117] In this embodiment, each unit is composed of four nodes, and the degree of freedom of a single node is 6, so the total degree of freedom of a unit is 24. Each node can be divided into translation and rotation along the three coordinate axes and expressed as:

[0118]

[0119] The superscript e represents the unit number, and the displacement matrix u of the four nodes in the unit is e It can be expressed as the following formula:

[0120]

[0121] In this embodiment, the strain ε on the surface of the structure e Tensile strain and bending strain and transverse shear strain The relationship is as follows:

[0122]

[0123] Among them, ε xx is the strain in the x direction, ε yy is the strain in the y direction, γ xy is the shear strain in the xy direction, γ xz is the shear strain in the xz direction, γ yz is the shear strain in the yz direction, and T represents the matrix transpose.

[0124] In this embodiment, if the general situation is taken into consideration, the strain value generated on the surface of the component is can be divided into tensile strain e(u e ) and bending strain k(u e ), the three are connected by a coefficient equation. Taking the partial derivative of the shape function, we get the matrix Processing the equation mathematically yields the following result:

[0125]

[0126] in, is the strain value generated on the component surface, e(u e ) represents tensile strain, k(u e ) represents the bending strain, z is the bending strain coefficient, u e represents the displacement matrix, and Taking partial derivatives of the shape function yields a matrix.

[0127] In this embodiment, the cross-sectional shear strain We can use the derivative of the shape function and the node displacement matrix u e express:

[0128]

[0129] in, represents the transverse shear strain, g(u e ) represents the displacement matrix u e The equation function of .

[0130] It can be further expressed as the following formula:

[0131]

[0132] in, They are The matrix after partial derivative of the shape function at nodes P1, P2, P3, and P4 can be expressed as the following formula:

[0133]

[0134]

[0135] In this embodiment, the inverse finite element algorithm achieves the ultimate goal of solving the inverse displacement field by comparing the strain measurement value and the calculated value, constructing the least squares error function between the two, and taking the strain value at the minimum value when the error function takes the minimum value, and then further inferring the reconstructed displacement field. The least squares error function Δ e (u e ) is constructed as follows:

[0136]

[0137] Among them, e(u e ),κ(u e ),γ(u e ) are the theoretical tensile and compressive strains, bending strains and shear strains, respectively, e ε , κ ε , γ ε are the corresponding measured strain values, w e 、w κ 、w γ is the weighting coefficient of the strain least squares method.

[0138] In this embodiment, an error function between theoretical strain and measured strain is established. When the error function takes a minimum value, an equation for solving displacement can be obtained. The equations of multiple units are converted to the overall coordinate system in the form of a matrix for assembly. Boundaries are imposed and combined with deformation continuity conditions to solve the displacement of each unit node.

[0139] The specific experimental method and process of the inverse finite element algorithm in this embodiment are as follows:

[0140] (1) The component to be measured is divided into units, and the unit plane is composed of several nodes. The calculation formula between unit strain and unit node displacement is derived.

[0141] (2) The fiber Bragg grating strain sensor 3 is used to obtain the actual measured strain data.

[0142] (3) The error function is expressed by the least square method. The minimum value of the error function is obtained by the derivative function, and the minimum value represents the unit stiffness matrix.

[0143] (4) Assembly of the unit stiffness matrix. This part of the assembly process is similar to the method in finite element analysis.

[0144] (5) By setting the unit boundary conditions, the displacement value is inversely calculated based on the relationship equation between displacement and force.

[0145] The truss assembly 1, fiber Bragg grating strain sensor 3, demodulator 5, exciter 6, etc. used in this embodiment respectively realize the functions of strain data acquisition and transmission, fiber optic sensing signal demodulation, vibration excitation, loading, etc. The acquisition system is composed of fiber Bragg grating strain sensor 3, demodulator 5, fiber jumper and signal processing software.

[0146] This embodiment is aimed at the spacecraft primary mirror 2 model, and the primary mirror 2 profile displacement field inversion is performed. The specific model structure installation and experimental steps are as follows:

[0147] The spacecraft primary mirror 2 model is suspended and fixed at the support rod, and an air flotation device is installed to simulate the spacecraft microgravity environment.

[0148] In this embodiment, symmetrical points are taken on the upper left, lower left, upper right and lower right of the mirror surface of the primary mirror 2 to fix the sensor, and each point is 20 cm away from the center of the profile. The fiber Bragg grating strain sensor 3 strain rosette is glued to the surface of the structure by AB epoxy glue and connected to the demodulator 5 through an optical fiber jumper. A fiber Bragg grating temperature compensation sensor 4 is fixed to each measuring point. The fiber Bragg grating strain sensor 3 is arranged at the same distance interval position on the truss, and is glued to the surface of the structure by AB epoxy glue and connected to the demodulator 5 through an optical fiber jumper. A fiber Bragg grating temperature compensation sensor 4 is fixed to each measuring point.

[0149] In this embodiment, reflective marks are pasted all over the test piece, and stickers with reflective material dots printed on the back are pasted as evenly and firmly as possible on the mirror surface of the truss and the primary mirror 2. An image displacement detector is used to shoot the unloaded test piece at multiple angles to determine the initial position of each reference point of the test piece.

[0150] In this embodiment, a loading test of the exciter 6 is carried out at the position of the primary mirror 2 of the spacecraft, and the test piece is photographed after loading. Three-dimensional reconstruction is completed based on the position change of the calibration point in the photographing result, and the displacement field after loading is determined.

[0151] In this embodiment, the strain data is connected to the computer through the demodulator 5 and the network cable 8. The user graphical interface is opened, and the operation buttons such as initialization, data acquisition, data preprocessing, primary mirror 2 truss displacement field reconstruction and primary mirror 2 displacement field reconstruction are clicked to realize the corresponding functions. After the button is selected and the function is completed, the button turns yellow.

[0152] In this experiment, the average error between the displacement change measured by the system proposed in this embodiment and the displacement measured by the imaging displacement sensor is less than 5%, which can meet the deformation monitoring requirements of the spacecraft primary mirror 2 under micro-disturbance.

[0153] This embodiment can realize the surface strain data collection of the main mirror 2 of the spacecraft under small disturbance conditions, providing a good data collection basis for the subsequent displacement reconstruction algorithm. The inverse finite element algorithm adopted does not need to know the mechanical parameters of the structure itself to realize the displacement field reconstruction and perform high-precision morphological perception of the main mirror 2 of the spacecraft.

[0154] This embodiment uses a strain rosette composed of three fiber Bragg grating strain sensors 3 and a fiber Bragg grating temperature compensation sensor 4 to collect the strain signal of the primary mirror 2 of the spacecraft for inverse finite element algorithm, improves the accuracy of strain sensing, and can fully monitor strains at different angles. On the other hand, the fiber Bragg grating strain sensor 3 has strong anti-interference ability and good durability in various environments. In addition, the fiber Bragg grating strain sensor 3 of this embodiment has a clear network structure, which is convenient for later maintenance.

[0155] In this embodiment, the signal obtained by the demodulator 5 is directly integrated into the user graphical interface through the UDP connection, and the UDP communication data decoding is completed in the user graphical interface general program. There is no need to use supporting software to receive wavelength data and save the file for subsequent data processing, and the data import and subsequent processing tasks are directly completed in the user graphical interface.

[0156] This embodiment uses the Ko displacement theory algorithm and the inverse finite element algorithm to reconstruct the displacement field of the spacecraft primary mirror 2. This method can reversely derive the structural displacement field through strain without requiring the mechanical parameters of the structure, and can improve the reconstruction accuracy of the displacement field. The inverse finite element algorithm based on the fiber Bragg grating strain sensor 3 has lower cost than the traditional displacement field perception and can achieve passive monitoring.

[0157] This embodiment uses the optical fiber sensing system to monitor the structural strain and realizes the surface monitoring of the primary mirror 2. It provides a graphical user interface to realize data visualization, making the software easier to operate, integrating multiple algorithms, making it convenient for users to call various algorithms and functions on the interface, and reducing the learning cost. It can be used for ground test verification of advanced spacecraft testing.

[0158] This embodiment combines the micro-strain monitoring research on the on-orbit structure of spacecraft at home and abroad and a complete set of monitoring systems and data processing methods to design a structural deformation perception method for the spacecraft main mirror 2 structure. By arranging fiber Bragg grating strain sensors 3 at various positions of the spacecraft main mirror 2 structure, and applying the Ko displacement theory algorithm and the inverse finite element algorithm to process the strain, the displacement field is reconstructed. When the spacecraft structure is deformed, the fiber Bragg grating strain sensor 3 is affected by the strain and the Bragg wavelength shift occurs. The wavelength signal is transmitted to the demodulator 5, and the demodulator 5 converts the optical signal into an electrical signal. The UDP communication signal transmission is completed by connecting to the computer through the network cable 8, and the signal decoding is displayed on the operation interface. After the signal is processed and the displacement is reconstructed in turn in the operation interface, the morphological perception of the spacecraft's main mirror 2 is visualized.

[0159] Example 2

[0160] In the second aspect, the present application proposes a high-precision ground optical test method for the primary mirror surface of a spacecraft, such as Figure 6 As shown, the high-precision ground optical test method for the spacecraft primary mirror surface includes the following steps:

[0161] Step S1, obtaining a strain signal of the truss assembly 1 and a strain signal of the primary mirror 2; the strain signal of the truss assembly 1 and the strain signal of the primary mirror 2 are strain signals of the truss assembly 1 and strain signals of the primary mirror 2 respectively collected by the high-precision ground optical test system for the spacecraft primary mirror profile described in Example 1.

[0162] Step S2, respectively demodulating the strain signal of the truss assembly 1 and the strain signal of the primary mirror 2 to obtain a truss assembly 1 strain demodulation signal and a primary mirror 2 strain demodulation signal.

[0163] Step S3, using the Ko displacement theory algorithm to perform displacement field inversion on the strain demodulation signal of the truss assembly 1 to obtain the inversion result of the truss assembly 1, and at the same time using the inverse finite element algorithm to perform displacement field inversion on the strain demodulation signal of the primary mirror 2 to obtain the inversion result of the primary mirror 2.

[0164] Step S4, superimposing the inversion result of the truss assembly 1 and the inversion result of the primary mirror 2 to obtain a morphological reconstruction result, thereby realizing the morphological reconstruction of the spacecraft primary mirror 2 structure.

[0165] In this embodiment, before the step of respectively demodulating the strain signal of the truss assembly 1 and the strain signal of the primary mirror 2 to obtain the strain demodulation signal of the truss assembly 1 and the strain demodulation signal of the primary mirror 2, the method for high-precision ground optical testing of the surface of the spacecraft primary mirror further includes the following steps:

[0166] The truss assembly 1 strain signal and the primary mirror 2 strain signal are denoised respectively to obtain a preprocessed truss assembly 1 strain signal and a preprocessed primary mirror 2 strain signal; the preprocessed truss assembly 1 strain signal and the preprocessed primary mirror 2 strain signal are used for demodulation as the truss assembly 1 strain signal and the primary mirror 2 strain signal respectively to obtain the truss assembly 1 strain demodulation signal and the primary mirror 2 strain demodulation signal.

[0167] In this embodiment, after respectively demodulating the strain signal of the truss assembly 1 and the strain signal of the primary mirror 2 to obtain the strain demodulation signal of the truss assembly 1 and the strain demodulation signal of the primary mirror 2, the high-precision ground optical testing method for the primary mirror profile of a spacecraft further includes the following steps:

[0168] An IP address is set, and the strain demodulation signal of the truss assembly 1 and the strain demodulation signal of the primary mirror 2 are imported into the computer 7 software using the UDP protocol, and displacement field inversion and morphological reconstruction are performed in the user graphical interface of the computer 7 software.

[0169] The technical features of the above embodiments may be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0170] This article uses specific examples to illustrate the principles and implementation methods of this application. The description of the above embodiments is only used to help understand the method and core ideas of this application. At the same time, for those skilled in the art, according to the ideas of this application, there will be changes in the specific implementation methods and application scope. In summary, the content of this specification should not be understood as limiting this application.

Claims

1. A high-precision ground optical test system for the primary mirror surface of a spacecraft, characterized in that: The high-precision ground optical test system for the spacecraft primary mirror profile comprises a truss assembly, a primary mirror, a sensor assembly, an exciter, a demodulator and a computer; The primary mirror is disposed at one end of the truss assembly, the exciter is disposed at the other end of the truss assembly, the sensor assembly is disposed at each measuring point of the truss assembly and each measuring point on the surface of the primary mirror, the output end of the sensor assembly is connected to the input end of the demodulator, and the output end of the demodulator is connected to the input end of the computer; The truss assembly and the primary mirror are used as a force-bearing component and a core observation component of the spacecraft respectively; The exciter is used to apply an excitation force to the truss assembly to cause the truss assembly and the primary mirror to vibrate; The sensing component is used to collect strain signals of the truss component and the primary mirror respectively, to obtain a truss component strain signal and a primary mirror strain signal; The demodulator is used to demodulate the truss assembly strain signal and the primary mirror strain signal respectively to obtain a truss assembly strain demodulation signal and a primary mirror strain demodulation signal; The computer is used to perform displacement field inversion on the strain demodulation signal of the truss assembly using the Ko displacement theory algorithm to obtain a truss assembly inversion result, and simultaneously perform displacement field inversion on the strain demodulation signal of the primary mirror using the inverse finite element algorithm to obtain a primary mirror inversion result, and superimpose the truss assembly inversion result and the primary mirror inversion result to obtain a morphological reconstruction result, thereby realizing morphological reconstruction of the spacecraft primary mirror structure.

2. The high-precision ground optical test system for the primary mirror surface of a spacecraft according to claim 1, characterized in that: The truss assembly includes three trusses, one end of the three trusses is fixed at the same point, and the other ends are evenly dispersed and fixed to the surface of the primary mirror.

3. The high-precision ground optical test system for the primary mirror of a spacecraft according to claim 1, characterized in that: The sensing component includes a first sensing component, which includes a plurality of first fiber grating strain sensors, which are evenly arranged at various measuring points on the surface of the primary mirror, and the first fiber grating strain sensors are used to collect the strain signal of the primary mirror.

4. The high-precision ground optical test system for the primary mirror surface of a spacecraft according to claim 3, characterized in that: The number of the first fiber Bragg grating strain sensors at each measuring point on the surface of the primary mirror is 3, one end of the three first fiber Bragg grating strain sensors is fixed at the same point, and the other ends are arranged at 0°, 45° and 90° in the horizontal direction to form a strain rosette structure.

5. The high-precision ground optical test system for the primary mirror surface of a spacecraft according to claim 3, characterized in that: The sensing component also includes a second sensing component, which includes a plurality of second fiber grating strain sensors, which are evenly arranged at various measuring points on the surface of the truss component, and the second fiber grating strain sensors are used to collect strain signals of the truss component.

6. The high-precision ground optical test system for the primary mirror surface of a spacecraft according to claim 5, characterized in that: The sensing assembly further comprises a third sensing assembly, wherein the third sensing assembly comprises a plurality of first fiber Bragg grating temperature compensation sensors and a plurality of second fiber Bragg grating temperature compensation sensors; A first fiber Bragg grating temperature compensation sensor is arranged at each measuring point on the surface of the primary mirror, and the first fiber Bragg grating temperature compensation sensor is used to perform temperature compensation on the first fiber Bragg grating strain sensor; A second fiber Bragg grating temperature compensation sensor is arranged at each measuring point on the surface of the truss assembly, and the second fiber Bragg grating temperature compensation sensor is used to perform temperature compensation on the second fiber Bragg grating strain sensor.

7. The high-precision ground optical test system for the primary mirror surface of a spacecraft according to claim 1, characterized in that: The computer includes a user graphical interface, which includes operation buttons for initialization, data acquisition, data preprocessing, truss component displacement field inversion, primary mirror displacement field inversion and morphological reconstruction. The user graphical interface is used to operate each of the operation buttons and display the truss component inversion result, the primary mirror inversion result and the morphological reconstruction result in real time.

8. A high-precision ground optical test method for the primary mirror surface of a spacecraft, characterized in that: The high-precision ground optical testing method for the spacecraft primary mirror surface comprises: Acquire a truss assembly strain signal and a primary mirror strain signal; the truss assembly strain signal and the primary mirror strain signal are respectively collected by a high-precision ground optical test system for the primary mirror profile of a spacecraft according to any one of claims 1 to 7; Demodulating the truss assembly strain signal and the primary mirror strain signal respectively to obtain a truss assembly strain demodulation signal and a primary mirror strain demodulation signal; The Ko displacement theory algorithm is used to perform displacement field inversion on the strain demodulation signal of the truss component to obtain the inversion result of the truss component, and the inverse finite element algorithm is used to perform displacement field inversion on the strain demodulation signal of the primary mirror to obtain the inversion result of the primary mirror; The truss assembly inversion result and the primary mirror inversion result are superimposed to obtain a morphological reconstruction result, thereby realizing the morphological reconstruction of the spacecraft primary mirror structure.

9. The high-precision ground optical testing method for the primary mirror surface of a spacecraft according to claim 8, characterized in that: Before the step of respectively demodulating the truss assembly strain signal and the primary mirror strain signal to obtain a truss assembly strain demodulation signal and a primary mirror strain demodulation signal, the spacecraft primary mirror profile high-precision ground optical testing method further includes: The truss assembly strain signal and the primary mirror strain signal are denoised respectively to obtain a preprocessed truss assembly strain signal and a preprocessed primary mirror strain signal; the preprocessed truss assembly strain signal and the preprocessed primary mirror strain signal are used for demodulation as the truss assembly strain signal and the primary mirror strain signal respectively to obtain the truss assembly strain demodulation signal and the primary mirror strain demodulation signal.

10. The high-precision ground optical testing method for the primary mirror surface of a spacecraft according to claim 8, characterized in that: After the steps of respectively demodulating the truss assembly strain signal and the primary mirror strain signal to obtain a truss assembly strain demodulation signal and a primary mirror strain demodulation signal, the spacecraft primary mirror profile high-precision ground optical testing method further includes: An IP address is set, and the truss assembly strain demodulation signal and the primary mirror strain demodulation signal are imported into computer software using a UDP transmission protocol, and displacement field inversion and morphological reconstruction are performed in a user graphical interface of the computer software.

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