A manned spacecraft structure full-field stress inversion system based on optical fiber measurement

Abstract

The application provides a manned spacecraft structure full-field stress inversion system based on optical fiber measurement, the maximum principal stress and the minimum principal stress at key positions of a sealed cabin are obtained through sensor units arranged at the key positions of the outer wall of the sealed cabin, then the maximum principal stress and the minimum principal stress at each position of the outer wall of the sealed cabin are inverted from the maximum principal stress and the minimum principal stress at the key positions based on the finite element mode, the global stress of the sealed cabin is predicted, the problem that the stress measurement data is only local points and leads to low stress prediction accuracy is solved, meanwhile, a solution for spacecraft life extension is provided, and the application has high economic benefits; in addition, based on the discrete strain data of the fiber grating strain sensor, the application can also realize full-field inversion of the structural strain state of a single cabin, two-cabin and three-cabin combination of a space station.

Description

Technical Field

[0001] This invention belongs to the field of aerospace, and in particular relates to a full-field stress inversion system for manned spacecraft structures based on fiber optic measurement. Background Technology

[0002] The on-orbit service environment of manned spacecraft is complex and harsh. Under the effects of long-term pressurization, temperature changes and loading and repositioning, the key parts of the load-bearing structure are prone to strength problems. In addition, unexpected situations such as micrometeoroids and space debris impacts can cause major accidents such as cabin leakage or tearing. It is necessary to establish a strain inversion system for the sealed cabin structure based on fiber optic measurement to provide relevant support for the life extension and repair of the sealed cabin structure of manned spacecraft.

[0003] McDonnell Douglas developed the DC-X, a demonstrator of a single-engine, reusable launch vehicle. NASA validated numerous new technologies on the subsequent upgraded DC-XA launch vehicle, including a structural health monitoring system. This system comprises an onboard unit and a ground-based unit at the launch site, communicating via a remote computer at the flight control center. In the PC-based DC-XA structural health monitoring system, fiber optic strain and temperature sensors on a composite liquid hydrogen tank monitor the system via fiber optic links. The fiber optic sensor data is demodulated and converted into signals suitable for long-distance transmission. Data analysis, recording, and graphical display are performed on the remote PC. Historical strain and temperature data for both the structure and the tank can be displayed separately through a strain-temperature display window.

[0004] The alternating loads on the main structure of the International Space Station's (ISS) sealed module during its on-orbit operation primarily originate from temperature changes, internal pressure changes, and alternating loads caused by the movement of on-orbit moving parts. Loads from moving parts mainly act on localized areas, while internal pressure and temperature loads act on the overall main structure. Construction of the ISS began in October 1998. It is a multi-functional space research complex with the participation of 16 national and regional organizations. Based on the ISS's fiber optic measurement system's function of collecting strain parameters from various modules during the ascent phase and on-orbit operation, the ISS needs to collect multiple structural parameters in real time to predict structural loads, temperature, and dynamic characteristics. Therefore, with the participation of institutions such as the Johnson Space Center in the United States, various integrated acquisition systems have been developed, and the development is gradually moving towards wireless sensing systems.

[0005] The United States has embedded a fiber optic grating sensing system in the X-37B spaceplane to monitor information such as temperature and pressure online in real time. The types of data monitored include parameters such as temperature and strain. Parameters such as strain and temperature are mainly measured using embedded fiber optic sensors, while vibration is mainly measured using piezoelectric sensors.

[0006] However, existing measurement methods based on limited points can only measure local strain information and often cannot monitor the most severely loaded parts, resulting in poor measurement accuracy and posing a threat to the safe operation of spacecraft in orbit. Summary of the Invention

[0007] To address the aforementioned problems, this invention provides a full-field stress inversion system for manned spacecraft structures based on fiber optic measurement, which can predict the global stress of the spacecraft.

[0008] A full-field stress inversion system for manned spacecraft structure based on fiber optic measurement includes a fiber optic demodulation module, a strain inversion module, and a sensor unit array arranged on key parts of the outer wall of the sealed cabin. In the launch phase during flight, the key parts are high-strain regions of the structure, and in the on-orbit phase during flight, the key parts are both high-strain regions and fatigue-prone regions of the structure.

[0009] Each sensor unit includes a temperature sensor and three fiber Bragg grating strain sensors installed at different angles. The fiber demodulation module provides a broadband light source for each fiber Bragg grating strain sensor and converts the optical signals fed back by each fiber Bragg grating strain sensor and the temperature signals measured by the temperature sensor into electrical signals. The strain inversion module obtains the maximum and minimum principal stresses at the installation positions of each sensor unit based on the electrical signals corresponding to the temperature sensor belonging to the same sensor unit and the fiber Bragg grating strain sensors at different angles. Then, based on the finite element method, it inverts the maximum and minimum principal stresses at each position on the outer wall of the entire sealed chamber based on the maximum and minimum principal stresses at each installation position.

[0010] Furthermore, the fiber optic strain sensor is elastically stretched and includes a spring sheet and an optical fiber disposed in a groove on the central axis of the spring sheet. The spring sheet has a centrally symmetrical structure, and the middle part of the spring sheet is a first rectangular piece. The two ends of the first rectangular piece are respectively connected to a hollow rectangular frame, and then connected to a second rectangular piece through the hollow rectangular frame. The height of the first rectangular piece is less than the height of the hollow rectangular frame and the second rectangular piece.

[0011] Furthermore, the overall size of the spring is designed to be 24mm×8mm×0.8mm, the length of the grating area on the optical fiber is 16mm, and solder joints are provided at both ends of the grating area. The middle part of the grating area is a 12mm grating, and the two ends are blank gratings of 2mm each used to isolate the solder joints.

[0012] Furthermore, the installation angles of the three fiber Bragg grating strain sensors in each sensor unit are 0°, 45°, and 90°, respectively. The calculation methods for the maximum and minimum principal stresses at the installation positions of each sensor unit are as follows:

[0013] Let εx =ε 0° , ε y =ε 90° γ xy =2ε 45° -(ε 0° +ε 90° ), where ε x With ε y ε represents the strain along the x and y axes at the current sensor unit installation location, respectively. 0° ε 45° ε 90° The strains γ are measured by fiber optic strain sensors installed at angles of 0°, 45°, and 90° in the current sensor unit, respectively. xy The shear strain at the current sensor unit installation location;

[0014] The maximum principal strain ε at the current sensor unit installation location max and minimum principal strain ε min The calculation method is as follows:

[0015]

[0016]

[0017] The maximum principal stress σ at the current sensor unit installation location max and minimum principal stress σ min The calculation method is as follows:

[0018]

[0019]

[0020] Where E is the elastic modulus and v is Poisson's ratio.

[0021] Furthermore, the method for calculating the strain measured by fiber optic strain sensors installed at angles of 0°, 45°, and 90° in the sensor unit is as follows:

[0022]

[0023] Where ε takes the value of ε 0° ε 45° ε 90° λ is the center wavelength measured by the fiber optic strain sensor after strain occurs, λ0 is the reference wavelength measured by the fiber optic strain sensor when no strain occurs, τ is the strain sensitivity coefficient of the fiber optic strain sensor, ΔT is the difference between the temperature measured by the temperature sensor and the zero temperature value, and P is the temperature compensation coefficient of the fiber optic strain sensor.

[0024] Furthermore, the method for deploying the sensor unit array at key locations on the outer wall of the sealed chamber is as follows:

[0025] Assuming the maximum number of sensor units that can be deployed in the current critical area is M, based on the structural dimensions of the current critical area, at least M / 2 sensor units are evenly deployed in the current critical area, and the position coordinates of the deployed sensor units are determined.

[0026] The coordinate information of the vulnerable high-strain region is extracted from the finite element model of the sealed chamber. New sensor units are deployed in the high-strain region until the number of sensor units in the current critical parts reaches the preset upper limit M, thus completing the deployment of sensor units. The newly deployed sensor units are located as far away from the existing sensor units as possible.

[0027] Furthermore, the method for retrieving the maximum and minimum principal stresses at various locations on the outer wall of the entire sealed chamber based on the maximum and minimum principal stresses at each installation location is as follows:

[0028] Obtain the position coordinates of each finite element and each sensor unit in the finite element model of the sealed chamber, and correct the mesh quality, load settings and boundary stiffness in the finite element model of the sealed chamber based on the position coordinates of each sensor unit.

[0029] The modified finite element model of the sealed chamber was sampled based on the Latin hypersolution method, and the sampled finite element calculated stress values ​​were used as the LFM dataset, while the measured maximum principal stress and minimum principal stress of each sensor unit were used as the HFM dataset.

[0030] By integrating LFM and HFM datasets using weighted interpolation or polynomial response surface methods, an in-orbit digital twin of the sealed cabin is obtained. Then, based on the inversion of the in-orbit digital twin of the sealed cabin, the maximum and minimum principal stresses at each sampled finite element location are obtained.

[0031] Beneficial effects:

[0032] 1. This invention provides a full-field stress inversion system for manned spacecraft structures based on fiber optic measurement. It acquires the maximum and minimum principal stresses at key locations on the outer wall of the sealed cabin using sensor units deployed at these key locations. Then, based on the finite element method, it inverts the maximum and minimum principal stresses at various locations on the entire outer wall of the sealed cabin from these maximum and minimum principal stresses, achieving global stress prediction for the sealed cabin. This solves the problem of low stress prediction accuracy caused by current stress measurement data only being available at local points. It also provides a solution for extending the lifespan of spacecraft and has high economic benefits. Furthermore, based on discrete strain data from fiber optic strain sensors, this invention can also achieve full-field inversion of the structural strain state in single-module, two-module, and three-module combined configurations of a space station.

[0033] 2. This invention provides a full-field stress inversion system for manned spacecraft structures based on fiber optic measurement. The hollow rectangular frame set on the spring sheet enables the fiber optic strain sensor to have spring-like stretchable properties. In other words, this invention can realize the main stress measurement of sealed cabins with different curvatures without affecting the long life of the fiber optic strain sensor by setting the elastic stretching body.

[0034] 3. This invention provides a full-field stress inversion system for manned spacecraft structures based on fiber optic measurement. The sensor units are optimized and arranged in a hierarchical and phased manner, which can enhance the accuracy of structural measurement and inversion.

[0035] 4. This invention provides a full-field stress inversion system for manned spacecraft structures based on fiber optic measurement. The overall size of the spring sheet is designed to be 24mm×8mm×0.8mm, which has the advantages of convenient machining and operation, and also ensures the strength of the sensor itself. At the same time, this invention reserves 2mm blank optical fibers at both ends of the grating area, which can avoid the solder joints from affecting the grating. Attached Figure Description

[0036] Figure 1 This is a schematic diagram of the composition of a full-field stress inversion system for manned spacecraft structures based on fiber optic measurement according to the present invention.

[0037] Figure 2 This is a schematic diagram of the sensor unit layout of the present invention;

[0038] Figure 3 This is a two-dimensional schematic diagram of the fiber optic strain sensor of the present invention;

[0039] Figure 4 This is a three-dimensional schematic diagram of the fiber optic strain sensor of the present invention;

[0040] Figure 5 This is a schematic diagram showing the dimensions of the fiber optic strain sensor of the present invention from its main view.

[0041] Figure 6 This is a schematic diagram of the dimensions of the fiber optic strain sensor of the present invention from a side view.

[0042] Figure 7 This is a schematic diagram of the installation method of the fiber optic strain sensor of the present invention;

[0043] Figure 8 This is a schematic diagram of the adhesive area of ​​the fiber optic strain sensor of the present invention;

[0044] Figure 9 This is a schematic diagram of the sensor unit layout of the present invention;

[0045] Figure 10 The master strain cloud map of the digital twin effect obtained by using the inversion system of this invention for the stable operation of a spacecraft in orbit;

[0046] Figure 11 The master strain cloud map is a digital twin effect obtained using the inversion system of this invention when a spacecraft is subjected to an impact in orbit. Detailed Implementation

[0047] To enable those skilled in the art to better understand the present application, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the accompanying drawings.

[0048] like Figure 1 As shown, a full-field stress inversion system for manned spacecraft structures based on fiber optic measurement includes a fiber optic demodulation module, a strain inversion module, and a sensor unit array arranged at key locations on the outer wall of the sealed cabin. During the launch phase of flight, these key locations are high-strain regions of the structure; during the on-orbit phase of flight, these key locations are both high-strain regions and areas of structural fatigue risk. Furthermore, based on manned spacecraft test and simulation results, typical key locations during the launch phase are high-strain regions, while typical key locations during the on-orbit phase include the connection between the wall panel and the frame, the robotic arm, and the solar array mounting wall panel, all of which are high-strain regions and areas of structural fatigue risk.

[0049] like Figure 2As shown, each sensor unit includes a temperature sensor and three fiber Bragg grating strain sensors with different installation angles. The fiber demodulation module provides a broadband light source for each fiber Bragg grating strain sensor and converts the optical signals fed back by each fiber Bragg grating strain sensor and the temperature signals measured by the temperature sensor into electrical signals. The strain inversion module obtains the maximum and minimum principal stresses at the installation positions of each sensor unit based on the electrical signals corresponding to the temperature sensor belonging to the same sensor unit and the fiber Bragg grating strain sensors at different angles. Then, based on the finite element method, it inverts the maximum and minimum principal stresses at each position on the outer wall of the entire sealed chamber based on the maximum and minimum principal stresses at each installation position.

[0050] Therefore, the inversion system of this invention includes a fiber optic sensor unit, a fiber optic cable assembly, a fiber optic demodulation unit, a control computer, and strain inversion software. The fiber optic sensor unit and fiber optic demodulation unit collect and store strain information from typical key parts of the cabin during the pre-launch and ascent phases, orbit changes, and rendezvous / docking and separation phases. The control computer then performs a full-field inversion of structural strain. The structural strain inversion system operates at regular intervals during different flight phases, including the pre-launch phase, launch phase, orbit changes, and rendezvous / docking.

[0051] The working principle of this invention is as follows: Based on the analysis results, fiber optic sensor units are arranged on the outer wall of the sealed cabin before the manned spacecraft launch. Before the sealed cabin structure is pressurized and depressurized and before orbit changes, the fiber optic sensor's fiber optic demodulation unit is powered on, and data acquisition begins. Data acquisition ends upon completion of the operation, and the acquired data is transmitted to the ground via the telemetry and control communication network. The control computer, through the strain inversion module, retrieves the cabin strain field results measured by the fiber optic sensors in the orbital segment.

[0052] In other words, the fiber optic demodulation module of this invention uses an internal light source to provide a broadband light source for the strain sensor, and the demodulator demodulates the feedback signal of the sensor unit, converting the wavelength change into an electrical signal that enters the data acquisition and processing system. The data can then be processed by the FPGA acquisition circuit, and the acquired sensor information is output to the central control module via the internal communication bus for packaging and storage of structural force parameter measurement fiber optic data.

[0053] It should be noted that ordinary fiber optic sensors typically use simple encapsulation of bare fibers. However, in-orbit health monitoring requires a long lifespan, and considering machining capabilities and ensuring the sensor's own strength, the thickness and width of the sensor should be minimized, and the fiber height can be reduced. Through simulation verification, the fiber optic grating strain sensor of this invention is shown below. Figures 3-8As shown, the fiber optic strain sensor is capable of elastic stretching and includes a spring sheet and an optical fiber disposed in a groove on the central axis of the spring sheet. The spring sheet has a centrally symmetrical structure, and the middle part of the spring sheet is a first rectangular piece. The two ends of the first rectangular piece are respectively connected to a hollow rectangular frame, and then connected to a second rectangular piece through the hollow rectangular frame. The height of the first rectangular piece is less than the height of the hollow rectangular frame and the second rectangular piece.

[0054] Furthermore, the overall dimensions of the spring are designed to be 24mm × 8mm × 0.8mm, ensuring that the sensor itself will not bend due to pre-strain. During fiber optic grating writing, the fiber coating needs to be stripped and recoated. A 1mm positioning tolerance is allowed at each end of the grating, therefore the grating length in the middle of the grating area is at least 12mm. To avoid the solder joints affecting the grating, 2mm of blank fiber is reserved at both ends of the grating in the middle, therefore the length of the grating area is at least 16mm. The solder joint size is approximately 0.8mm, therefore the distance between the outer edges of the two solder joints is at least 17.6mm. For all these reasons, the distance between the two solder joints is designed to be 18mm. In addition, a 3mm length is reserved on the outer side of each of the two solder joints of the sensor for fixing the sleeve.

[0055] Furthermore, the fiber Bragg grating strain sensor has a measurement range covering ±1700 με, and the temperature sensor measures the temperature change at the strain point, providing temperature compensation for the strain measurement results. The fiber Bragg grating strain sensor has an installation area of ​​55×55 mm and a fiber coil area of ​​280×180 mm. The temperature sensor has an accuracy of ±0.5℃ and a measurement range of -30 to 60℃; specifically, the fiber Bragg grating strain sensor has a measurement range of -2660 με to 2420 με and a sampling rate greater than 800 Hz. For the fiber optic cable assembly, the insertion loss of a single cable assembly is no greater than 1.5 dB, and the return loss is no less than 40 dB.

[0056] The following describes the method for deploying the sensor unit array in key areas of the sealed chamber's outer wall, as detailed below:

[0057] Assuming the maximum number of sensor units that can be deployed in the current critical area is M, based on the structural dimensions of the current critical area, at least M / 2 sensor units are uniformly deployed on the current critical area, and the position coordinates of the deployed sensor units are determined. The coordinates of the vulnerable high-strain regions are extracted from the finite element model of the sealed chamber, and new sensor units are deployed in the high-strain regions until the number of sensor units in the current critical area reaches the preset maximum limit M, thus completing the deployment of sensor units. The positions of the newly deployed sensor units are as far away from the existing sensor units as possible.

[0058] In other words, based on the finite element simulation results of the spacecraft structure, this invention identifies the high-strain zone of the launch section structure as a typical critical component and deploys sensors at this critical component. The proposed sensor deployment optimization method mainly includes the following steps: ① Based on the structural dimensions, a small number of sensors are evenly deployed on the structure to determine their position coordinates; ② In the structural finite element model, the coordinates of vulnerable high-strain regions are extracted, and new sensors are deployed around these locations, ensuring that the new sensors are as far away from existing sensors as possible to guarantee spatial coverage; ③ The number of sensors is continuously increased until the total number of sensors on the structure meets a pre-set maximum value, thus completing the sensor deployment. The specific implementation effect is as follows: Figure 9 As shown.

[0059] The following describes the installation method of the sensor unit array on key parts of the outer wall of the sealed chamber, as detailed below:

[0060] ① Preparation of fiber optic sensors and mounting fixtures. Remove adhesive, inspect and clean the fiber optic sensor connector end face, wipe and clean the fixtures. Grind the structural mounting area. Grind the structural mounting area with appropriate sandpaper, ensuring no deep scratches are created. Clean the ground area with a lint-free cloth after grinding. ② Sensor installation on the fixture. Install the sensor on the fixture according to the fiber optic coiling method and detailed requirements for sensor installation on the fixture. ③ Pre-installation and feeler gauge check. Pre-install without adhesive, and check the gap between the fiber optic sensor and the structural mounting surface with a feeler gauge of appropriate size. If it cannot be inserted, it is acceptable. Apply adhesive to the adhesive area. Apply adhesive to the sensor adhesive area using an adhesive applicator, scraper, etc. The amount of adhesive on the sensor should be appropriate, without overflow or shortage. ④ Sensor installation. Remove the adhesive applicator and quickly check the adhesive on the back of the applicator. If it meets the requirements, quickly press the sensor onto the structural mounting surface using the fixing device within the specified time. Do not rotate or lift the installed fiber optic sensor before the structure has cured. After the adhesive has cured, remove the fixture and use a feeler gauge to check the gap between the fiber optic sensor and the structural mounting surface.

[0061] The calculation methods for the maximum and minimum principal stresses at the installation locations of each sensor unit are described below:

[0062] Let ε x =ε 0° , ε y =ε 90° γ xy =2ε 45° -(ε 0° +ε 90° ), where ε x With ε y ε represents the strain along the x and y axes at the current sensor unit installation location, respectively. 0° ε45° ε 90° The strains γ are measured by fiber optic strain sensors installed at angles of 0°, 45°, and 90° in the current sensor unit, respectively. xy The shear strain at the current sensor unit installation location;

[0063] The maximum principal strain ε at the current sensor unit installation location max and minimum principal strain ε min The calculation method is as follows:

[0064]

[0065]

[0066] The maximum principal stress σ at the current sensor unit installation location max and minimum principal stress σ min The calculation method is as follows:

[0067]

[0068]

[0069] Where E is the elastic modulus and v is Poisson's ratio.

[0070] The calculation method for strain measured by fiber optic strain sensors installed at angles of 0°, 45°, and 90° in the sensor unit is as follows:

[0071]

[0072] Where ε is the strain value (με), and takes values ​​of ε. 0° ε 45° ε 90° λ is the center wavelength (pm) measured by the fiber Bragg grating strain sensor after strain occurs, λ0 is the reference wavelength (pm) measured by the fiber Bragg grating strain sensor before strain occurs, τ is the strain sensitivity coefficient (pm / με) of the fiber Bragg grating strain sensor, ΔT is the difference (°C) between the temperature measured by the temperature sensor and the zero temperature value, P is the temperature compensation coefficient (pm / °C) of the fiber Bragg grating strain sensor, and 22.8 is the expansion coefficient of 5A06 aluminum alloy at 20°C.

[0073] The following describes how to deduce the maximum and minimum principal stresses at various locations on the outer wall of the entire sealed chamber based on the maximum and minimum principal stresses at each installation location. The details are as follows:

[0074] Step 1: Perform coordinate system management on the finite element measurement point positions and sensor measurement point positions, and extract the finite element and measured data for the corresponding measurement point positions.

[0075] Step 2: Analyze potential errors in the finite element simulation calculation based on sensor data (mesh quality, load settings, boundary stiffness, etc.), and adjust the finite element model to improve its accuracy and ensure a good correlation with the sensor data.

[0076] Step 2: The finite element results have a large amount of data, and directly carrying out deep learning or building a proxy model is computationally expensive. Therefore, sampling is performed based on the Latin hypercube sampling (LHS) method. The stress values ​​calculated by the finite element simulation after sampling are used as the LFM (latent factor model) dataset, and the sensor measured data are used as the HFM dataset.

[0077] Step 3: Based on techniques such as weighted interpolation and polynomial response surface methodology, high-precision integration of various data types, including sensor measured data and finite element simulation data, is achieved to construct a digital twin. Based on the proposed method, on-orbit digital twins of typical modules are constructed. During stable on-orbit operation of the spacecraft, sensor data and finite element simulation data are integrated to construct the digital twin, and the principal strain contour map is obtained by inverting the mechanical field. Figure 10 Simulating impact loads, demonstrating the principal strain cloud map of a digital twin when a spacecraft is subjected to an impact in orbit, such as... Figure 11 .

[0078] Therefore, this invention achieves global prediction of principal stresses through structural inversion algorithm.

[0079] Of course, the present invention may have other various embodiments. Without departing from the spirit and essence of the present invention, those skilled in the art can make various corresponding changes and modifications according to the present invention, but these corresponding changes and modifications should all fall within the protection scope of the appended claims.

Claims

1. A full-field stress inversion system for manned spacecraft structures based on fiber optic measurement, characterized in that, It includes an optical fiber demodulation module, a strain inversion module, and a sensor unit array arranged on key parts of the outer wall of the sealed cabin. During the flight period, the key parts are the high strain zone of the structure during the launch phase and the high strain zone and the structural fatigue danger zone during the on-orbit phase of the flight period. Each sensor unit includes a temperature sensor and three fiber Bragg grating strain sensors installed at different angles. The fiber demodulation module provides a broadband light source for each fiber Bragg grating strain sensor and converts the optical signals fed back by each fiber Bragg grating strain sensor and the temperature signals measured by the temperature sensor into electrical signals. The strain inversion module obtains the maximum and minimum principal stresses at the installation positions of each sensor unit based on the electrical signals corresponding to the temperature sensor belonging to the same sensor unit and the fiber Bragg grating strain sensors at different angles. Then, based on the finite element method, it inverts the maximum and minimum principal stresses at each position on the outer wall of the entire sealed chamber based on the maximum and minimum principal stresses at each installation position. The installation angles of the three fiber Bragg grating strain sensors in each sensor unit are 0°, 45°, and 90°, respectively. The calculation methods for the maximum and minimum principal stresses at the installation positions of each sensor unit are as follows: make , , ,in, and These represent the strain along the x and y axes at the current sensor unit installation location, respectively. , , The strains measured are those obtained by fiber optic strain sensors installed at angles of 0°, 45°, and 90° in the current sensor unit. The shear strain at the current sensor unit installation location; Maximum principal strain at the current sensor unit installation location and minimum principal strain The calculation method is as follows: Maximum principal stress at the current sensor unit installation location and minimum principal stress The calculation method is as follows: Where E is the elastic modulus. v Poisson's ratio; The calculation method for strain measured by fiber optic strain sensors installed at angles of 0°, 45°, and 90° in the sensor unit is as follows: in, The values ​​are respectively , , , The center wavelength is obtained by the fiber optic strain sensor after strain has occurred. The reference wavelength is obtained by the fiber optic strain sensor when no strain has occurred. This represents the strain sensitivity coefficient of the fiber Bragg grating strain sensor. P is the temperature difference between the temperature measured by the temperature sensor and the zero-point temperature value, and P is the temperature compensation coefficient of the fiber optic strain sensor.

2. The full-field stress inversion system for manned spacecraft structures based on fiber optic measurement as described in claim 1, characterized in that, The fiber optic strain sensor is capable of elastic stretching and includes a spring sheet and an optical fiber disposed in a groove on the central axis of the spring sheet. The spring sheet has a centrally symmetrical structure, and the middle part of the spring sheet is a first rectangular piece. The two ends of the first rectangular piece are respectively connected to a hollow rectangular frame, and then connected to a second rectangular piece through the hollow rectangular frame. The height of the first rectangular piece is less than the height of the hollow rectangular frame and the second rectangular piece.

3. The full-field stress inversion system for manned spacecraft structures based on fiber optic measurement as described in claim 2, characterized in that, The overall dimensions of the spring are designed to be 24mm×8mm×0.8mm. The length of the grating area on the optical fiber is 16mm, and solder joints are provided at both ends of the grating area. The middle part of the grating area is a 12mm grating, and the two ends are blank gratings of 2mm each used to isolate the solder joints.

4. A full-field stress inversion system for manned spacecraft structures based on fiber optic measurement as described in any one of claims 1 to 3, characterized in that, The method for deploying the sensor unit array in key areas of the sealed chamber's outer wall is as follows: Assuming the maximum number of sensor units that can be deployed in the current critical area is M, based on the structural dimensions of the current critical area, at least M / 2 sensor units are evenly deployed in the current critical area, and the position coordinates of the deployed sensor units are determined. The coordinate information of the vulnerable high-strain region is extracted from the finite element model of the sealed chamber. New sensor units are deployed in the high-strain region until the number of sensor units in the current critical parts reaches the preset upper limit M, thus completing the deployment of sensor units. The newly deployed sensor units are located as far away from the existing sensor units as possible.

5. A full-field stress inversion system for manned spacecraft structures based on fiber optic measurement as described in any one of claims 1 to 3, characterized in that, The method for retrieving the maximum and minimum principal stresses at various locations on the outer wall of the entire sealed chamber from the maximum and minimum principal stresses at each installation location is as follows: Obtain the position coordinates of each finite element and each sensor unit in the finite element model of the sealed chamber, and correct the mesh quality, load settings and boundary stiffness in the finite element model of the sealed chamber based on the position coordinates of each sensor unit. The modified finite element model of the sealed chamber was sampled based on the Latin hypersolution method, and the sampled finite element calculated stress values ​​were used as the LFM dataset, while the measured maximum principal stress and minimum principal stress of each sensor unit were used as the HFM dataset. By integrating LFM and HFM datasets using weighted interpolation or polynomial response surface methods, an in-orbit digital twin of the sealed cabin is obtained. Then, based on the inversion of the in-orbit digital twin of the sealed cabin, the maximum and minimum principal stresses at each sampled finite element location are obtained.

Images