Aircraft component thermal stress detection device and method

Abstract

The invention discloses a thermal stress detection device and method for an aircraft component, and the device is characterized in that a unidirectional strain measurement grating, a strain rosette 0-degree direction strain measurement grating, a strain rosette 45-degree direction strain measurement grating and a strain rosette 90-degree direction strain measurement grating of a fiber grating strain sensing network module are respectively pasted on preset detection parts; forming a test piece; the grating strain demodulator is connected with an optical fiber in the fiber grating strain sensing network module; the temperature compensation module comprises a temperature compensation grating, and the temperature compensation grating is located at a part which is not influenced by strain in the fiber grating strain sensing network module; the fiber bragg grating strain sensing network module, the aircraft component and the temperature compensation module are all located in the environment box. The embodiment of the invention can meet the thermal stress test requirements of the large structure of the airplane, is simple to install, and can avoid the influence of temperature and stress on thermal stress measurement accuracy factors as much as possible.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of thermal stress detection of aircraft components, and particularly relates to a thermal stress detection device and method for aircraft components. BACKGROUND

[0002] Composite and metal materials are inevitably connected with each other in the structure of an aircraft, and such connection is relatively common in important connection positions. The aircraft is assembled at room temperature, but is operated in different high and low temperature environments. Thermal stress is generated due to the significant difference in thermal performance between the composite and metal materials, and the influence of thermal stress on the static strength and fatigue life of the structure at some key positions of the aircraft cannot be ignored. Therefore, the thermal stress analysis method and test method for the aircraft structure containing a composite and metal hybrid structure are indispensable in the process of aircraft development, and are crucial to ensure the safety of the structure in use or service.

[0003] Since the thermal stress test needs to be implemented in an environmental chamber, the cost of the full-aircraft-level thermal stress test is high and there are many technical difficulties, and therefore, in engineering, the results of low-level tests are used to verify the thermal stress analysis method, and the verified analysis method is used to verify the airworthiness of the full-aircraft-level aircraft. The anisotropy of the composite material also brings inconvenience to the characterization of the thermal performance of the material. The thermal expansion coefficient of the composite material is closely related to the thermal expansion coefficient of the fiber, the arrangement mode, the volume content, and the thermal expansion coefficient of the matrix, the volume content of the matrix, and the interface properties. The anisotropy characteristics of the composite laminate make the thermal performance parameters significantly affected by the layup, and the large number of layup libraries in the structure design brings inconvenience to the determination of the material properties during thermal stress analysis. In engineering, the theoretical empirical formula based on the relationship between the thermal performance parameters and the layup obtained by a certain number of original tests is used to calculate the thermal parameters, and then the structure internal force under thermal load is analyzed by finite element analysis. The mismatch of the thermal expansion coefficients of the hybrid structure under high and low temperature environments brings not only local thermal stress, but also global thermal stress. Research shows that the thermal load caused by temperature change has a greater impact on the end bolt load, and the superposition of the thermal load and the mechanical load will exacerbate the uneven distribution of the end bolt load. In addition, the bolt diameter and the bolt hole gap also have a greater impact on the bolt load distribution. Since there are a large number of fasteners in the structure, the differences in the pre-tightening force of the fasteners and the structure gap will affect the bolt load distribution and the force transmission path, and these influences are difficult to reflect in the finite element model, so that the overall analysis results of the large structure are often inaccurate in complex structure regions, especially in the metal-composite hybrid connection region.

[0004] Therefore, thermal stress test of large components is inevitable, on the one hand to obtain accurate thermal stress distribution of structure, on the other hand test data can be used to correct and verify thermal stress analysis method of large components. Thermal stress test of large structure needs to be carried out in a large environmental chamber, a large number of strain and temperature sensors and data acquisition systems are needed, strain gauge is usually used to measure structure strain in traditional mechanics test, strain gauge is a very mature method, and the cost is also low, but its size is large, when involving large area and multi-position monitoring, the actual use process is more complex, the installation time is longer, and the detection channel is more. In addition, thermal stress test also requires that the sensing element and its connecting line have high temperature resistance and high temperature measurement accuracy, and the temperature is also changing in thermal stress test, so dynamic temperature compensation needs to be made for the strain gauge, which increases the difficulty of data measurement and engineering practical application. SUMMARY

[0005] The application provides a thermal stress detection device and method of an aircraft component to meet the thermal stress test requirements of large aircraft structures, which is simple to install and can avoid the influence of temperature and stress on the thermal stress measurement accuracy as much as possible.

[0006] In a first aspect, the application provides a thermal stress detection device for an aircraft component, the aircraft component comprising a plurality of hybrid structures formed by composite materials and metal materials connected to each other, the hybrid structure comprising a plurality of preset detection sites; the thermal stress detection device comprising: a fiber Bragg grating strain sensing network module, a temperature compensation module, an environmental chamber and a grating strain demodulator.

[0007] The fiber Bragg grating strain sensing network module comprises a plurality of single-direction strain measurement grating units and a plurality of three-direction strain flower grating units.

[0008] The single-direction strain measurement grating unit comprises a single-direction strain measurement grating; the three-direction strain flower grating unit comprises a strain flower 0° direction strain measurement grating, a strain flower 45° direction strain measurement grating and a strain flower 90° direction strain measurement grating arranged on the same optical fiber.

[0009] The single-direction strain measurement grating, the strain flower 0° direction strain measurement grating, the strain flower 45° direction strain measurement grating and the strain flower 90° direction strain measurement grating are respectively pasted on each of the preset detection sites to form a test piece.

[0010] The grating strain demodulator is connected with the optical fiber in the fiber grating strain sensing network module; the temperature compensation module comprises a temperature compensation grating, which is located at a part of the fiber grating strain sensing network module that is not affected by strain; the fiber grating strain sensing network module, the aircraft component and the temperature compensation module are all located in the environmental box.

[0011] Optionally, a sensitivity calibration module is further included, which comprises a shockproof platform, a constant-temperature heating table, a first displacement platform, a second displacement platform, a first displacement adjustment knob, a second displacement adjustment knob, a heat insulation ceramic cabin, a sensitivity calibration optical fiber, a sensitivity calibration grating and a sensitivity calibration thermocouple; the grating strain demodulator is connected with the output ends of the sensitivity calibration optical fiber and the sensitivity calibration thermocouple;

[0012] The constant-temperature heating table, the first displacement platform and the second displacement platform are all located on the shockproof platform; the first displacement adjustment knob is arranged on the first displacement platform, and the second displacement adjustment knob is arranged on the second displacement platform; the sensitivity calibration optical fiber at both ends of the sensitivity calibration grating is fixed at a first fixed point of the first displacement platform and a second fixed point of the second displacement platform respectively, and the first fixed point and the second fixed point are located at the same horizontal position; the sensitivity calibration grating is located in the heat insulation ceramic cabin and at the intermediate position between the first fixed point and the second fixed point; the heat insulation ceramic cabin is placed on the constant-temperature heating table, and the collection end of the sensitivity calibration thermocouple is located in the heat insulation ceramic cabin;

[0013] The first displacement adjustment knob is used to move the first displacement platform, so as to control the strain amount of the sensitivity calibration grating, and the second displacement adjustment knob is used to move the second displacement platform, so as to control the strain amount of the sensitivity calibration grating; the constant-temperature heating table is used to control the temperature of the environment where the sensitivity calibration grating is located;

[0014] When measuring the strain sensitivity coefficient of the sensitivity calibration grating at a preset temperature, at the preset temperature, the distance between the first fixed point and the second fixed point is first adjusted so that the sensitivity calibration optical fiber is arranged in a straight line, then the first displacement platform or the second displacement platform is controlled to move a preset distance for multiple times, the sensitivity calibration optical fiber is stretched, and the strain and wavelength of the sensitivity calibration grating after each movement are recorded, a first linear function of the wavelength of the sensitivity calibration grating with respect to the strain is fitted, and the slope of the first linear function is determined as the strain sensitivity coefficient;

[0015] When measuring the temperature sensitivity coefficient of the sensitivity calibration grating, wavelength data of the sensitivity calibration grating under different temperatures and without stretching of the sensitivity calibration fiber is selected, a second linear function of the wavelength of the sensitivity calibration grating with respect to the temperature is fitted, and a slope of the second linear function is determined as the temperature sensitivity coefficient;

[0016] The sensitivity calibration grating, the gratings in the fiber grating strain sensing network module, and the temperature compensation grating are of the same shape, size, and material.

[0017] Optionally, the thermal stress detection device for the aircraft component further comprises a sticking assembly.

[0018] The sticking assembly comprises an auxiliary sticking mold and a glue, the auxiliary sticking mold is used to fix the shape and volume of the glue, and the glue is used to stick the gratings in the fiber grating strain sensing network module to the preset detection sites.

[0019] Optionally, the thermal stress detection device further comprises a support fixture.

[0020] The support fixture comprises a wooden support, and the wooden support is used to support the aircraft component.

[0021] Optionally, the environmental box comprises a temperature monitoring module and an environmental box thermocouple.

[0022] The temperature monitoring module is used to adjust the temperature in the environmental box, and the environmental box thermocouple is used to detect the temperature in the environmental box.

[0023] Optionally, the wavelength resolution of the grating strain demodulator is 0.01 nm, and the demodulation rate of the grating strain demodulator is 1 Hz.

[0024] The grating strain demodulator comprises a real-time strain display unit, a grating number of a main interface of the real-time strain display unit corresponds to a grating number in the fiber grating strain sensing network module in a one-to-one manner, and the real-time strain display unit is used to display the original wavelength of the grating in the fiber grating strain sensing network module and a strain value monitored by the grating in real time.

[0025] In a second aspect, the embodiments of the present application further provide a thermal stress detection method for an aircraft component, which is implemented by using the thermal stress detection device for the aircraft component as described in the first aspect, and the thermal stress detection method for the aircraft component comprises the following steps.

[0026] Design and manufacture a fiber grating strain sensing network module according to the preset detection sites;

[0027] Stick the gratings in the fiber grating strain sensing network module to each of the preset detection sites to form a test piece.

[0028] placing the test piece in an environmental chamber;

[0029] determining the strain force on the gratings in the fiber grating strain sensing network module.

[0030] Optionally, after the fiber grating strain sensing network module is designed and manufactured according to the preset detection site, the method further comprises:

[0031] determining the strain sensitivity coefficient and the temperature sensitivity coefficient of the sensitivity calibration grating at different temperatures.

[0032] Optionally, the determining the strain sensitivity coefficient and the temperature sensitivity coefficient of the sensitivity calibration grating at different temperatures comprises:

[0033] when measuring the strain sensitivity coefficient of the sensitivity calibration grating at a preset temperature, at the preset temperature, the distance between the first fixed point and the second fixed point is first adjusted to be equal to the sensitivity calibration fiber, then the first displacement platform or the second displacement platform is controlled to move a preset distance for multiple times to stretch the sensitivity calibration fiber, and the strain and the wavelength of the sensitivity calibration grating after each movement are recorded, a first linear function of the wavelength of the sensitivity calibration grating with respect to the strain is fitted, and the slope of the first linear function is determined as the strain sensitivity coefficient;

[0034] when measuring the temperature sensitivity coefficient of the sensitivity calibration grating, the wavelength data of the sensitivity calibration grating under the condition that the sensitivity calibration fiber is not stretched at each temperature is selected, a second linear function of the wavelength of the sensitivity calibration grating with respect to the temperature is fitted, and the slope of the second linear function is determined as the temperature sensitivity coefficient.

[0035] Optionally, the determining the strain force on the gratings in the fiber grating strain sensing network module comprises:

[0036] determining the center wavelength offset of the gratings in the fiber grating strain sensing network module and the temperature compensation grating;

[0037] obtaining the thermal stress on the gratings in the fiber grating strain sensing network module according to the center wavelength offset of the gratings in the fiber grating strain sensing network module and the temperature compensation grating, the strain sensitivity coefficient, the temperature sensitivity coefficient, and temperature change.

[0038] The embodiment of the present application provides a thermal stress detection device and method of an aircraft component, the aircraft component comprising a plurality of hybrid structures formed by composite materials and metal materials being connected with each other, the hybrid structure comprising a plurality of preset detection parts; the thermal stress detection device comprises: a fiber grating strain sensing network module, a temperature compensation module, an environmental box and a grating strain demodulator; the fiber grating strain sensing network module comprises a plurality of single-direction strain measurement grating units and a plurality of three-direction strain flower grating units; the single-direction strain measurement grating unit comprises a single-direction strain measurement grating; the three-direction strain flower grating unit comprises a strain flower 0° direction strain measurement grating, a strain flower 45° direction strain measurement grating and a strain flower 90° direction strain measurement grating arranged on the same optical fiber; the single-direction strain measurement grating, the strain flower 0° direction strain measurement grating, the strain flower 45° direction strain measurement grating and the strain flower 90° direction strain measurement grating are respectively pasted on each preset detection part to form a test piece; the grating strain demodulator is further connected with the optical fiber in the fiber grating strain sensing network module; the temperature compensation module comprises a temperature compensation grating, and the temperature compensation grating is located at a part in the fiber grating strain sensing network module which is not affected by strain; the fiber grating strain sensing network module, the aircraft component and the temperature compensation module are all located in the environmental box. The embodiment of the present application can meet the thermal stress test demand of a large aircraft structure, the fiber grating strain sensing network module is adopted, installation is simple, the temperature compensation grating is located at a part in the fiber grating strain sensing network module which is not affected by strain, and the influence of temperature and stress on the thermal stress measurement accuracy factor can be avoided as much as possible.

[0039] It should be understood that the description in this section is not intended to identify key or critical features of the embodiments of the present application or to limit the scope of the present application. Other features of the present application will become apparent from the following description. BRIEF DESCRIPTION OF DRAWINGS

[0040] In order to more clearly illustrate the technical solutions in the embodiments of the present application, the following will briefly introduce the drawings needed to be used in the embodiment description. Obviously, the drawings in the following description are only some embodiments of the present application, and other drawings can also be obtained by those skilled in the art without any creative effort on the basis of these drawings.

[0041] Figure 1 A structure schematic diagram of a hybrid structure provided by the embodiment of the present application;

[0042] Figure 2 A structure schematic diagram of another hybrid structure provided by the embodiment of the present application;

[0043] Figure 3 A structure schematic diagram of another hybrid structure provided by the embodiment of the present application;

[0044] Figure 4 A structural diagram of another hybrid structure provided for an embodiment of the present application;

[0045] Figure 5 A structural diagram of a thermal stress detection device for an aircraft component provided for an embodiment of the present application;

[0046] Figure 6 A structural diagram of a unidirectional strain measurement grating unit provided for an embodiment of the present application;

[0047] Figure 7 A structural diagram of a three-directional strain measurement grating unit provided for an embodiment of the present application;

[0048] Figure 8 A structural diagram of a sensitivity calibration module provided for an embodiment of the present application;

[0049] Figure 9 A test diagram of the wavelength and strain relationship of a sensitivity calibration grating at various temperatures provided for an embodiment of the present application;

[0050] Figure 10 A test diagram of the wavelength and temperature relationship of a sensitivity calibration grating provided for an embodiment of the present application;

[0051] Figure 11 A flowchart of a thermal stress detection method for an aircraft component provided for an embodiment of the present application;

[0052] Figure 12 A flowchart of another thermal stress detection method for an aircraft component provided for an embodiment of the present application. DETAILED DESCRIPTION

[0053] In order to make the personnel in the art better understand the present application, the technical solutions in the embodiments of the present application will be described clearly and completely in conjunction with the drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, but not all the embodiments. Based on the embodiments in the present application, all the other embodiments obtained by the personnel in the art without creative labor should belong to the protection scope of the present application.

[0054] It should be noted that the terms "first", "second", and the like in the description and claims of the application and the above drawings are used to distinguish similar objects, and do not necessarily indicate a specific order or sequence. It should be understood that the data used in this way can be interchanged under appropriate circumstances, so that the embodiments of the application described herein can be implemented in an order other than those illustrated or described herein. In addition, the terms "include" and "have" and any variations thereof are intended to cover non-exclusive inclusion, for example, a process, method, system, product or device including a series of steps or units does not necessarily limit to those steps or units clearly listed, but can include other steps or units not clearly listed or inherent to these processes, methods, products or devices.

[0055] Figure 1 A structural diagram of a hybrid structure is provided for an embodiment of the application, specifically a structural diagram of the hybrid structure under normal temperature conditions. Figure 1 The hybrid structure includes a composite material 1, a metal material 2, and a metal fastener 3, and the composite material 1 and the metal material 2 are fixedly connected through the metal fastener 3. Figure 2 Another structural diagram of a hybrid structure is provided for an embodiment of the application, specifically a structural diagram of the hybrid structure under the condition of excessively low temperature, Figure 3 Another structural diagram of a hybrid structure is provided for an embodiment of the application, specifically a structural diagram of the hybrid structure under the condition of excessively high temperature, referring to Figure 2 And Figure 3 Excessively low temperature or excessively high temperature will cause the hybrid structure to deform. Figure 4 Another structural diagram of a hybrid structure is provided for an embodiment of the application, referring to Figure 4 The hybrid structure composed of the mixed connection of the composite material 1 and the metal material 2 causes the local stress change of the metal material 2 due to the overall thermal effect of the aircraft structure, that is, the overall thermal stress is generated.

[0056] It can be understood that the aircraft is assembled under room temperature conditions, but operates in different high and low temperature environments. Due to the significant difference in thermal performance between the composite material and the metal material, thermal stress is generated, and the thermal stress at some key positions has a non-negligible influence on the static strength and fatigue life of the structure. The local thermal stress of the hybrid structure can be defined as the internal deformation and internal stress of the mixed connection area caused by the mismatch of the thermal expansion coefficient and the temperature effect. In order to meet the thermal stress test requirements of the large structure of the aircraft, an embodiment of the application provides a thermal stress detection device and method for an aircraft component, specifically as follows:

[0057] The aircraft component of the embodiment of the application includes a plurality of hybrid structures connected by the composite material and the metal material, and the hybrid structure includes a plurality of preset detection sites. Figure 5A structural schematic diagram of a thermal stress detection device for an aircraft component is provided in an embodiment of the present application, referring to Figure 5 The thermal stress detection device 100 comprises a fiber Bragg grating strain sensing network module 110, a temperature compensation module 120, an environmental box 130, and a grating strain demodulator 140.

[0058] Figure 6 A structural schematic diagram of a single-direction strain measurement grating unit is provided in an embodiment of the present application, Figure 7 A structural schematic diagram of a three-direction strain measurement grating unit is provided in an embodiment of the present application. Referring to Figures 6-7 In an embodiment of the present application, the fiber Bragg grating strain sensing network module 110 comprises a plurality of single-direction strain measurement grating units 111 and a plurality of three-direction strain flower grating units 112. The single-direction strain measurement grating unit 111 comprises a single-direction strain measurement grating 1111. The three-direction strain flower grating unit 112 comprises a strain flower 0° direction strain measurement grating 1121, a strain flower 45° direction strain measurement grating 1122, and a strain flower 90° direction strain measurement grating 1123 arranged on the same optical fiber 113.

[0059] In an embodiment of the present application, the single-direction strain measurement grating 1111, the strain flower 0° direction strain measurement grating 1121, the strain flower 45° direction strain measurement grating 1122, and the strain flower 90° direction strain measurement grating 1123 are respectively pasted on each preset detection site to form a test piece.

[0060] It should be noted that by analyzing and evaluating the regions with relatively large stress in the aircraft component under the target thermal load, the key positions to be detected, i.e., the preset detection sites, are determined. Generally, the thermal stress of the hybrid structure formed by the mutual connection of the composite material and the metal material is relatively complex, and a large number of monitoring points are arranged. A small number of monitoring points are arranged in the region without hybrid structure connection or simple structure. The fiber routing arrangement is designed according to the preset detection sites, the grating length and interval are designed, the number of gratings connected in series by a single optical fiber is determined according to the number of monitoring points and the demodulation strategy, and thus the number of channels is determined. The fiber grating array is prepared by the dynamic online grating technology of the drawing tower at one time, and the fiber grating has high-temperature strain and temperature test stability and accuracy through the carbon coating process.

[0061] For example, through finite element analysis, 72 single pieces and 56 flower pieces, i.e., a total of 240 monitoring point positions, are required to be arranged for strain monitoring of a certain large aircraft structure. According to the selection of the monitoring point density and quantity, the minimum spatial resolution of the grating measurement point is 50 mm, the grating length is 50 mm, and 32 channels are used to design the arrangement of the fiber grating sensing network. The corresponding grating array is prepared by the dynamic online grating technology of the drawing tower at one time, and the fiber grating has high-temperature test function through the carbon coating process.

[0062] In the embodiment of the present application, the grating strain demodulator 140 is connected with the optical fiber in the fiber grating strain sensing network module 110; the temperature compensation module 120 comprises a temperature compensation grating, and the temperature compensation grating is located at a position in the fiber grating strain sensing network module 110 which is not affected by strain; the fiber grating strain sensing network module 110, the aircraft component and the temperature compensation module 120 are all located in the environmental box 130.

[0063] It can be understood that the temperature compensation grating is suspended to ensure that there is no interaction force between the temperature compensation grating and the test piece. By increasing a plurality of force-free temperature compensation gratings, when the fiber grating sensor is used as a strain measurement sensor, the effect of the change of the ambient temperature on the center wavelength shift can be offset, so that the purpose of strain measurement is achieved. When the test piece is subjected to a thermal load, the strain measurement grating pasted on the surface of the aircraft component is subjected to the effects of temperature and thermal strain. When the temperature changes by ΔT, the center wavelength shift can be expressed as: wherein Δλ1 is the center wavelength shift of the grating pasted on the aircraft component, λ1 is the initial center wavelength of the grating pasted on the aircraft component, K1 is the strain sensitivity coefficient, K2 is the temperature sensitivity coefficient, ε is the thermal strain of the grating pasted on the aircraft component, and ΔT is the temperature change.

[0064] For the temperature compensation grating, when the temperature changes by ΔT, the center wavelength shift can be expressed as: wherein Δλ2 is the center wavelength shift of the temperature compensation grating, λ2 is the initial center wavelength of the temperature compensation grating, K2 is the temperature sensitivity coefficient, and ΔT is the temperature change.

[0065] Therefore, the thermal strain of the grating pasted on the aircraft component is:

[0066] According to the center wavelength shift, the initial center wavelength and the strain sensitivity coefficient of the grating pasted on the aircraft component and the temperature compensation grating in the force-free temperature state, the measured strain can be determined.

[0067] The embodiment of the present application can meet the demand of the thermal stress test of the large aircraft structure by reasonably designing the optical fiber sensing network, pasting the single direction strain measurement grating unit 111 or the three direction strain flower grating unit 112 at the preset detection position of the aircraft part, controlling the temperature in the environmental box 130 to apply the required temperature load, processing the temperature influence of the grating strain monitoring through the temperature compensation module 120, analyzing the strain of the preset detection position through the grating strain demodulator 140, and monitoring the change of the strain with the temperature in real time, so as to complete the thermal stress test of the test piece according to the required temperature change condition. The optical fiber grating strain sensing network module 110 is simple to install, the temperature compensation grating is located at the position not affected by the strain in the optical fiber grating strain sensing network module 110, and the influence of the temperature and the stress on the thermal stress measurement precision can be avoided as much as possible.

[0068] Figure 8 A structural schematic diagram of a sensitivity calibration module provided by the embodiment of the present application is shown in FIG. 13. Optionally, on the basis of the above embodiment, referring to FIG. 13, Figure 8 The thermal stress detection device of the aircraft part further includes a sensitivity calibration module 150, and the sensitivity calibration module 150 includes a shockproof platform 151, a constant temperature heating table 152, a first displacement platform 153, a second displacement platform 154, a first displacement adjusting knob 155, a second displacement adjusting knob 156, a heat insulation ceramic cabin 157, a sensitivity calibration optical fiber 158, a sensitivity calibration grating 159, and a sensitivity calibration thermocouple 1510. In the embodiment of the present application, the grating strain demodulator 140 is connected with the output ends of the sensitivity calibration optical fiber 158 and the sensitivity calibration thermocouple 1510.

[0069] Continuing to refer to Figure 8The constant temperature heating platform 152, the first displacement platform 153 and the second displacement platform 154 are all located on the anti-vibration platform 151; the first displacement adjustment knob 155 is set on the first displacement platform 153, and the second displacement adjustment knob 156 is set on the second displacement platform 154; the sensitivity calibration optical fiber 158 at both ends of the sensitivity calibration grating 159 is respectively fixed to the first fixed point A of the first displacement platform 153 and the second fixed point B of the second displacement platform 154, and the first fixed point A and the second fixed point B are located at the same horizontal position X; the sensitivity calibration grating 159 is located in the thermal insulation ceramic cabin 157, and is located between the first fixed point A and the second fixed point B; the thermal insulation ceramic cabin 157 is placed on the constant temperature heating platform 152. In this embodiment of the present invention, the collection end of the sensitivity calibration thermocouple 1510 is located in the thermal insulation ceramic cabin 157. The first displacement adjustment knob 155 is used to move the first displacement platform 153, thereby controlling the strain of the sensitivity calibration grating 159, and the second displacement adjustment knob 156 is used to move the second displacement platform 154, thereby controlling the strain of the sensitivity calibration grating 159; the constant temperature heating platform 152 is used to control the temperature of the environment in which the sensitivity calibration grating 159 is located.

[0070] In an embodiment of the present invention, when measuring the strain sensitivity coefficient of the sensitivity calibration grating 159 at a preset temperature, at the preset temperature, the distance D between the first fixed point A and the second fixed point B is first adjusted so that the sensitivity calibration optical fiber 158 is arranged in a straight line. Then, the sensitivity calibration optical fiber 158 is stretched by repeatedly controlling the first displacement platform 153 or the second displacement platform 154 to move the preset distance. The strain and wavelength of the sensitivity calibration grating 159 after each movement are recorded, and a first linear linear function of the wavelength of the sensitivity calibration grating 159 with respect to the strain is fitted. The slope of the first linear linear function is determined as the strain sensitivity coefficient. When measuring the temperature sensitivity coefficient of the sensitivity calibration grating 159, wavelength data of the sensitivity calibration grating 159 at different temperatures when the sensitivity calibration optical fiber 158 is not stretched is selected to fit a second linear linear function of the wavelength of the sensitivity calibration grating 159 with respect to the temperature. The slope of the second linear linear function is determined as the temperature sensitivity coefficient.

[0071] Specifically, Figure 9 A schematic diagram of a test of the relationship between the wavelength and strain of a sensitivity calibration grating at various temperatures provided by an embodiment of the present invention. Figure 10 A schematic diagram of a test of the relationship between the wavelength and temperature of a sensitivity calibration grating provided in an embodiment of the present invention. It can be understood that by controlling the movement of the first displacement platform 153 or the second displacement platform 154, with a preset distance Δd as a step, the strain each time is equal to n·Δd / D, where n is the number of movements. The strain and the wavelength of the sensitivity calibration grating 159 are recorded each time, and a first linear function of the wavelength of the sensitivity calibration grating 159 with respect to the strain is fitted, with reference toFigure 9 The slope of the first linear function of the wavelength of the sensitivity calibration grating 159 with respect to the strain is the strain sensitivity coefficient K1 of the sensitivity calibration grating 159 at the temperature, i.e., λ=K1·ε, where λ is the wavelength of the sensitivity calibration grating 159, and ε is the strain. The strain sensitivity coefficient of the sensitivity calibration grating 159 at different temperatures is measured in the embodiment of the present application. The wavelength data of the sensitivity calibration grating 159 when the sensitivity calibration fiber 158 is not stretched at each temperature is selected in the embodiment of the present application, and a second linear function of the wavelength of the sensitivity calibration grating 159 with respect to the temperature is fitted, i.e., λ=K2·T, where λ is the wavelength of the sensitivity calibration grating 159, K2 is the temperature sensitivity coefficient of the sensitivity calibration grating 159, and T is the temperature. Figure 10 The slope of the linear function is the temperature sensitivity coefficient of the sensitivity calibration grating 159.

[0072] It can be understood that the strain sensitivity coefficient of the grating usually does not change under the influence of different temperatures, and the average value of the strain sensitivity coefficient of the grating under various working conditions is taken as the strain sensitivity coefficient of the grating.

[0073] The sensitivity calibration grating 159, the grating in the fiber grating strain sensing network module 110, and the temperature compensation grating are of the same shape, size, and material.

[0074] Optionally, on the basis of the above embodiment, reference is made to Figure 6 and Figure 7 The thermal stress detection device 100 of the aircraft component further includes a sticking assembly 160. The sticking assembly 160 includes an auxiliary sticking mold 161 and a sticking glue 162, the auxiliary sticking mold 161 is used to fix the shape and volume of the sticking glue 162, and the sticking glue 162 is used to stick the grating in the fiber grating strain sensing network module 110 to a preset detection position.

[0075] Specifically, in order to make all the gratings have uniform sensitivity after being stuck, the auxiliary sticking mold 161 is used to stick and fix the shape and volume of the sticking glue 162. The sticking glue 162 of the embodiment of the present application has high temperature resistance, which can ensure the stability of the grating position and strain transmission during the entire high temperature test. The gratings are stuck in three different directions based on the sticking method to form a three-direction strain flower grating unit 112.

[0076] For example, the fiber direction is adjusted along the measuring direction, the center of the grating is aligned with the preset detection part, a certain pre-stress is applied to the fiber grating, the two ends of the grating are fixed by using the adhesive 162, the auxiliary bonding mold 161 is pressed around the grating to make the grating located in the center of the long rectangular hole, the auxiliary bonding mold 161 is fixed around the frame by using the adhesive 162, the fixed adhesive 162 needs to be a high-temperature-resistant adhesive, the bonding performance is maintained during the whole test, and no deformation and displacement occurs, the adhesive should be uniformly distributed and thin during the gluing, so that the fiber grating is maximally prevented from being affected by the non-uniformity of the adhesive in the high-temperature and low-temperature environments. After the fixed adhesive 162 is completely solidified, the fixed auxiliary bonding mold 161 is removed, and the grating bonding of one strain monitoring point is completed. For the monitoring points that need to be monitored in three directions, the grating in three directions is sequentially bonded in the same way, and the reference Figure 7 direction is monitored by using the fiber grating in the other two directions.

[0077] Optionally, on the basis of the above-mentioned embodiment, the thermal stress detection device 100 further comprises a support fixture. The support fixture comprises a wooden support, and the wooden support is used for supporting the aircraft component.

[0078] It can be understood that the support fixture, such as the wooden support, which is less affected by temperature, is used to support the test piece, so that the test piece can freely expand under temperature stress, and the detailed support position and the profile of the wooden support can be designed according to actual conditions.

[0079] Optionally, on the basis of the above-mentioned embodiment, the environmental box 130 comprises a temperature monitoring module and environmental box thermocouples. The temperature monitoring module is used for adjusting the temperature in the environmental box 130, and the environmental box thermocouples are used for detecting the temperature in the environmental box 130.

[0080] In the embodiment of the application, the grating strain demodulator 140 is first zeroed, and initial data is recorded. The temperature monitoring module of the environmental box 130 is adjusted, and the temperature in the environmental box 130 is increased to the target temperature. After each target temperature is stabilized, the temperature is kept for 15 minutes, strain data is measured and collected, the temperature of the environmental box 130 is then decreased, and the temperature in the environmental box 130 is decreased to the target temperature of the cooling stage. After each target temperature is stabilized, the temperature is kept for 15 minutes, strain data is measured and collected, and the temperature setting maximum overshoot temperature is less than or equal to 5℃.

[0081] Optionally, on the basis of the above embodiment, the wavelength resolution of the grating strain demodulator 140 is 0.01 nm, and the demodulation rate of the grating strain demodulator 140 is 1 Hz.

[0082] Optionally, on the basis of the above embodiment, the grating strain demodulator 140 includes a real-time strain display unit, the grating number of the main interface of the real-time strain display unit corresponds to the number of the gratings in the fiber grating strain sensing network module 110 one by one, and the real-time strain display unit is used to display the original wavelength of the grating in the fiber grating strain sensing network module 110 and the strain value monitored by the grating.

[0083] It can be understood that the grating strain demodulator 140 has a real-time strain display function, and the upper computer sub-interface can display the grating duration strain curve and update in real time.

[0084] It should be noted that the monitoring and recording of the grating spectrum can be realized by a spectrum analyzer or a wavelength and time division sensing demodulator and other spectrum monitoring instruments. According to the required number of measurement points, the wavelength resolution of the demodulator and the demodulation rate, the number of channels that meet the requirements is designed.

[0085] Compared with the single-point measurement of thermocouples and strain gauges, the thermal stress detection device for aircraft parts provided by the embodiment of the present application has the characteristics of more wiring. The fiber grating can realize dense grating sensors on a single optical fiber through various multiplexing methods, simplify the data acquisition channel, and the fiber grating sensor has the advantages of small size, light weight, corrosion resistance, high temperature resistance, and electromagnetic interference resistance. The grating sensitivity test device of the present application can simultaneously measure the high temperature strain sensitivity coefficient and the temperature sensitivity coefficient of the grating. The grating pasting method of the provided pasting assembly can ensure the stability of the grating position and strain transmission during the entire high temperature test. The present application can be used to monitor the mechanical state of the structure under environmental thermal load, realize the thermal stress test of the composite and metal large hybrid structure, and at the same time, the method can be used under cyclic thermal load, and can be used to evaluate the residual strength and fatigue performance of the structure under cyclic thermal load. Thus, the present application provides a method and condition for verifying the implementation of the thermal stress test of the composite and metal hybrid structure under environmental load, and can also be used to verify and correct the structure thermal stress analysis method.

[0086] Figure 11 The flowchart of the thermal stress detection method for aircraft parts provided by the embodiment of the present application is realized by using the thermal stress detection device 100 for aircraft parts provided by the above embodiment, and reference is made to Figure 11 , the thermal stress detection method for aircraft parts includes:

[0087] S210, designing and manufacturing the fiber grating strain sensing network module according to the preset detection position.

[0088] S220, paste the grating in the fiber grating strain sensing network module with each preset detection site to form a test piece.

[0089] S230, place the test piece in the environmental box.

[0090] S240, determine the strain force borne by the grating in the fiber grating strain sensing network module.

[0091] Figure 12 Another flowchart of the method for detecting thermal stress of an aircraft component is provided in the embodiments of the present application, and optionally, based on the above-mentioned embodiments, with reference to Figure 12 , the method comprises:

[0092] S310, design and manufacture the fiber grating strain sensing network module according to the preset detection site.

[0093] S320, determine the strain sensitivity coefficient and the temperature sensitivity coefficient of the sensitivity calibration grating under different temperatures.

[0094] S330, paste the grating in the fiber grating strain sensing network module with each preset detection site to form a test piece.

[0095] S340, place the test piece in the environmental box.

[0096] S350, determine the strain force borne by the grating in the fiber grating strain sensing network module.

[0097] Optionally, based on the above-mentioned embodiments, the step S320 comprises: when measuring the strain sensitivity coefficient of the sensitivity calibration grating at the preset temperature, at the preset temperature, first adjust the distance between the first fixed point and the second fixed point to be equal to the sensitivity calibration fiber, then control the first displacement platform or the second displacement platform to move a preset distance for multiple times to stretch the sensitivity calibration fiber, and record the strain and wavelength of the sensitivity calibration grating after each movement, fit the first linear function of the wavelength of the sensitivity calibration grating with respect to the strain, and determine the slope of the first linear function as the strain sensitivity coefficient. When measuring the temperature sensitivity coefficient of the sensitivity calibration grating, select the wavelength data of the sensitivity calibration grating under the condition that the sensitivity calibration fiber is not stretched at each temperature, fit the second linear function of the wavelength of the sensitivity calibration grating with respect to the temperature, and determine the slope of the second linear function as the temperature sensitivity coefficient.

[0098] Optionally, based on the above-mentioned embodiment, the step S350 comprises: determining the center wavelength offset of the grating and the temperature compensation grating in the fiber grating strain sensing network module. The thermal stress suffered by the grating in the fiber grating strain sensing network module is obtained according to the center wavelength offset of the grating and the temperature compensation grating in the fiber grating strain sensing network module, the strain sensitivity coefficient, the temperature sensitivity coefficient and the temperature change.

[0099] The method for detecting thermal stress of an aircraft component provided by the embodiments of the present application is implemented by using the aircraft component thermal stress detection device provided by the above-mentioned embodiments, and thus has the same beneficial effects. The content not described in detail in the embodiments of the present application can be referred to the aircraft component thermal stress detection device provided by the above-mentioned embodiments.

[0100] It should be understood that the various forms of flow shown above can be used to reorder, add or delete steps. For example, each step described in the present application can be executed in parallel, sequentially or in a different order, as long as the desired results of the technical solutions of the present application can be achieved, which is not limited herein.

[0101] The above detailed description does not constitute a limitation on the protection scope of the present application. Those skilled in the art should understand that various modifications, combinations, sub-combinations and substitutions can be made according to design requirements and other factors. Any modification, equivalent replacement and improvement within the spirit and principles of the present application should be included in the protection scope of the present application.

Claims

1. A thermal stress detection device for aircraft components, characterized in that: The aircraft component comprises a plurality of hybrid structures formed by interconnecting composite materials and metal materials, wherein the hybrid structure includes a plurality of preset detection locations; the thermal stress detection device comprises: a fiber Bragg grating strain sensing network module, a temperature compensation module, an environmental chamber, and a Bragg grating strain demodulator; The fiber Bragg grating strain sensing network module includes a plurality of unidirectional strain measurement grating units and a plurality of three-directional strain rosette grating units; The unidirectional strain measurement grating unit includes a unidirectional strain measurement grating; the three-directional strain rosette grating unit includes a strain rosette 0° direction strain measurement grating, a strain rosette 45° direction strain measurement grating, and a strain rosette 90° direction strain measurement grating arranged on the same optical fiber; The unidirectional strain measurement grating, the strain rosette 0° direction strain measurement grating, the strain rosette 45° direction strain measurement grating, and the strain rosette 90° direction strain measurement grating are respectively attached to the preset detection locations to form a test piece; The grating strain demodulator is connected to the optical fiber in the fiber Bragg grating strain sensing network module; the temperature compensation module includes a temperature compensation grating, which is located in a part of the fiber Bragg grating strain sensing network module that is not affected by strain; the fiber Bragg grating strain sensing network module, the aircraft component and the temperature compensation module are all located in the environmental chamber.

2. The thermal stress detection device for aircraft components according to claim 1, characterized in that: The apparatus further comprises a sensitivity calibration module, the sensitivity calibration module comprising an anti-vibration platform, a constant temperature heating platform, a first displacement platform, a second displacement platform, a first displacement adjustment knob, a second displacement adjustment knob, a thermal insulation ceramic cabin, a sensitivity calibration optical fiber, a sensitivity calibration grating, and a sensitivity calibration thermocouple; the grating strain demodulator is connected to the output ends of the sensitivity calibration optical fiber and the sensitivity calibration thermocouple; The constant temperature heating platform, the first displacement platform, and the second displacement platform are all located on the anti-vibration platform; the first displacement adjustment knob is provided on the first displacement platform, and the second displacement adjustment knob is provided on the second displacement platform; The sensitivity calibration optical fibers at both ends of the sensitivity calibration grating are respectively fixed to a first fixed point of the first displacement platform and a second fixed point of the second displacement platform, and the first fixed point and the second fixed point are located at the same horizontal position; the sensitivity calibration grating is located in the thermal insulation ceramic cabin and is located between the first fixed point and the second fixed point; the thermal insulation ceramic cabin is placed on the constant temperature heating table, and the collection end of the sensitivity calibration thermocouple is located in the thermal insulation ceramic cabin; The first displacement adjustment knob is used to move the first displacement platform, thereby controlling the strain of the sensitivity calibration grating; the second displacement adjustment knob is used to move the second displacement platform, thereby controlling the strain of the sensitivity calibration grating; the constant temperature heating stage is used to control the temperature of the environment in which the sensitivity calibration grating is located; When measuring the strain sensitivity coefficient of the sensitivity calibration grating at a preset temperature, at the preset temperature, first adjusting the distance between the first fixed point and the second fixed point so that the sensitivity calibration optical fiber is arranged in a straight line, then repeatedly controlling the first displacement platform or the second displacement platform to move the preset distance, stretching the sensitivity calibration optical fiber, and recording the strain and wavelength of the sensitivity calibration grating after each movement, fitting a first linear function of the wavelength of the sensitivity calibration grating with respect to the strain, and determining the slope of the first linear linear function as the strain sensitivity coefficient; When measuring the temperature sensitivity coefficient of the sensitivity calibration grating, wavelength data of the sensitivity calibration grating when the sensitivity calibration optical fiber is not stretched at different temperatures are selected, a second linear function of the wavelength of the sensitivity calibration grating with respect to temperature is fitted, and the slope of the second linear function is determined to be the temperature sensitivity coefficient; The sensitivity calibration grating, the grating in the fiber Bragg grating strain sensing network module, and the temperature compensation grating are all the same in shape, size and material.

3. The thermal stress detection device for aircraft components according to claim 1, characterized in that: The aircraft component thermal stress detection device further includes an adhesive assembly; The pasting component includes an auxiliary pasting mold and adhesive. The auxiliary pasting mold is used to fix the shape and volume of the adhesive. The adhesive is used to paste the grating in the fiber grating strain sensing network module on the preset detection position.

4. The thermal stress detection device for aircraft components according to claim 1, characterized in that: The thermal stress detection device further includes: a supporting fixture; The supporting fixture includes a wooden support, and the wooden support is used to support the aircraft component.

5. The thermal stress detection device for aircraft components according to claim 1, characterized in that: The environmental chamber includes a temperature monitoring module and an environmental chamber thermocouple; The temperature monitoring module is used to adjust the temperature in the environmental box, and the environmental box thermocouple is used to detect the temperature in the environmental box.

6. The thermal stress detection device for aircraft components according to claim 1, characterized in that: The wavelength resolution of the grating strain demodulator is 0.01 nm, and the demodulation rate of the grating strain demodulator is 1 Hz; The grating strain demodulator includes a real-time strain display unit. The grating numbers on the main interface of the real-time strain display unit correspond one-to-one with the grating numbers in the fiber Bragg grating strain sensing network module, and are used to display the original wavelength of the grating in the fiber Bragg grating strain sensing network module and the strain value monitored by the grating in real time.

7. A method for detecting thermal stress of aircraft components, characterized in that: The method for detecting thermal stress of an aircraft component is implemented by using the thermal stress detection device of any one of claims 1 to 6, and the method for detecting thermal stress of an aircraft component comprises: Design and manufacture fiber Bragg grating strain sensing network modules according to preset detection locations; Pasting the grating in the fiber Bragg grating strain sensing network module to each of the preset detection locations to form a test piece; placing the test piece in an environmental chamber; The strain force exerted on the grating in the fiber optic Bragg grating strain sensing network module is determined.

8. The method for detecting thermal stress of aircraft components according to claim 7, characterized in that: After designing and manufacturing the fiber Bragg grating strain sensing network module according to the preset detection position, the method further includes: Determine the strain sensitivity coefficient and temperature sensitivity coefficient of the sensitivity calibration grating at different temperatures.

9. The method for detecting thermal stress of aircraft components according to claim 8, characterized in that: Determining the strain sensitivity coefficient and temperature sensitivity coefficient of the sensitivity calibration grating at different temperatures includes: When measuring the strain sensitivity coefficient of the sensitivity calibration grating at a preset temperature, at the preset temperature, first adjusting the distance between the first fixed point and the second fixed point to be equal to the sensitivity calibration optical fiber, then repeatedly controlling the first displacement platform or the second displacement platform to move the preset distance, stretching the sensitivity calibration optical fiber, and recording the strain and wavelength of the sensitivity calibration grating after each movement, fitting a first linear function of the wavelength of the sensitivity calibration grating with respect to the strain, and determining the slope of the first linear function as the strain sensitivity coefficient; When measuring the temperature sensitivity coefficient of the sensitivity calibration grating, the wavelength data of the sensitivity calibration grating when the sensitivity calibration optical fiber is not stretched at each temperature is selected, and a second linear linear function of the wavelength of the sensitivity calibration grating with respect to temperature is fitted, and the slope of the second linear linear function is determined to be the temperature sensitivity coefficient.

10. The method for detecting thermal stress of aircraft components according to claim 8, characterized in that: Determining the strain force exerted on the grating in the fiber Bragg grating strain sensing network module includes: Determining the center wavelength offset of the grating in the fiber Bragg grating strain sensing network module and the temperature compensation grating; The thermal stress on the grating in the fiber Bragg grating strain sensing network module is obtained according to the central wavelength offset of the grating in the fiber Bragg grating strain sensing network module and the temperature compensation grating, the strain sensitivity coefficient, the temperature sensitivity coefficient and the temperature change.

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