A device and method for thermal stress detection of an aircraft component

Abstract

The application discloses a kind of thermal stress detection device and method of aircraft component, one-way strain measurement grating of fiber grating strain sensing network module, strain rosette 0 ° direction strain measurement grating, strain rosette 45 ° direction strain measurement grating and strain rosette 90 ° direction strain measurement grating are respectively pasted on each preset detection site, constitute test piece;Strain demodulator and fiber in fiber grating strain sensing network module are connected;Temperature compensation module includes temperature compensation grating, temperature compensation grating is located in fiber grating strain sensing network module and is not affected by strain part;Fiber grating strain sensing network module, aircraft component and temperature compensation module are located in environmental box.The embodiment of the application can meet the demand of aircraft large structure thermal stress test, simple to install, and can avoid the influence of temperature, stress on thermal stress measurement accuracy factor as far as possible.

Description

Technical Field

[0001] This invention relates to the field of thermal stress testing technology for aircraft components, and more particularly to a thermal stress testing device and method for aircraft components. Background Technology

[0002] Aircraft structures inevitably involve the bonding of composite materials and metals, and this bonding is relatively common in critical connection areas. Aircraft are assembled at room temperature, but operate in varying high and low temperature environments. The significant differences in the thermal properties of composite materials and metals generate thermal stress, which has a non-negligible impact on the static strength and fatigue life of the structure in some critical locations. Therefore, thermal stress analysis and testing methods for aircraft structural components containing composite and metal hybrid structures are indispensable in the aircraft development process and are crucial for ensuring the safety of the structure during use and service.

[0003] Because thermal stress testing needs to be conducted in an environmental chamber, conducting full-aircraft-level thermal stress testing is costly and technically challenging. In engineering practice, lower-level test results are generally used to verify thermal stress analysis methods, and then the verified analysis methods are used to verify the airworthiness compliance of the entire aircraft. The anisotropy of composite materials also complicates the characterization of their thermal properties. The coefficient of thermal expansion of composite materials is closely related to the coefficient of thermal expansion of the fibers, their arrangement, volume content, as well as the coefficient of thermal expansion of the matrix, matrix volume content, and interface properties. This anisotropic characteristic of composite laminates makes their thermal performance parameters significantly affected by the ply configuration. The vast ply library in structural design makes it difficult to determine material properties during thermal stress analysis. In engineering, thermal stress analysis uses theoretical and empirical formulas derived from a certain number of prototype tests to calculate thermal parameters based on the relationship between thermal performance parameters and ply configurations. Then, finite element analysis is used to analyze the internal forces of the structure under thermal load. The mismatch in the coefficients of thermal expansion of hybrid structures under high and low temperature environments can lead to both local and global thermal stresses. Studies have shown that thermal loads caused by temperature changes have a significant impact on end-mounted loads. The superposition of thermal and mechanical loads exacerbates the uneven distribution of end-mounted loads. In addition, bolt diameter and bolt hole gap also have a significant impact on load distribution. Due to the large number of fasteners in the structure, differences in fastener preload and structural gaps will affect load distribution and force transmission paths. These effects are difficult to reflect in finite element models, so the overall analysis results of large structural components are often inaccurate in complex structural regions, especially in metal-composite hybrid connection areas.

[0004] Therefore, thermal stress testing of large components is unavoidable. This serves two purposes: firstly, to obtain accurate thermal stress distribution within the structure; and secondly, to allow for the correction and validation of thermal stress analysis methods for large components. Thermal stress testing of large structural components requires large environmental chambers and numerous strain and temperature sensors and data acquisition systems. Traditional mechanical testing typically uses strain gauges to measure structural strain. While strain gauge measurement is a mature and cost-effective method, its large size leads to complex setup and operation, longer installation times, and more detection channels when monitoring large areas and multiple locations. Furthermore, thermal stress testing requires sensing elements and their connecting circuits to have high-temperature resistance and high-temperature measurement accuracy. Since the temperature is constantly changing during thermal stress testing, dynamic temperature compensation for the strain gauges is necessary, further increasing the difficulty of data measurement and practical engineering applications. Summary of the Invention

[0005] This invention provides a thermal stress detection device and method for aircraft components to meet the requirements of thermal stress testing of large aircraft structures. It is simple to install and can minimize the influence of temperature and stress on the accuracy of thermal stress measurement.

[0006] In a first aspect, embodiments of the present invention provide a thermal stress detection device for an aircraft component. The aircraft component includes multiple hybrid structures formed by interconnecting composite materials and metal materials. The hybrid structures include multiple preset detection locations. The thermal stress detection device includes: a fiber optic strain sensing network module, a temperature compensation module, an environmental chamber, and a grating strain demodulator.

[0007] The fiber optic strain sensing network module includes multiple unidirectional strain measurement grating units and multiple tridirectional strain flower grating units.

[0008] The unidirectional strain measurement grating unit includes a unidirectional strain measurement grating; the three-directional strain flower grating unit includes 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 disposed on the same optical fiber.

[0009] 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 each of the preset detection locations to form a test piece;

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

[0011] Optionally, it also includes a sensitivity calibration module, which includes a vibration-proof platform, a constant-temperature heating stage, a first displacement platform, a second displacement platform, a first displacement adjustment knob, a second displacement adjustment knob, a heat-insulating ceramic chamber, a sensitivity calibration fiber, a sensitivity calibration grating, and a sensitivity calibration thermocouple; the grating strain demodulator is connected to the output end of the sensitivity calibration fiber and the sensitivity calibration thermocouple.

[0012] The constant temperature heating stage, the first displacement platform, and the second displacement platform are all located on the vibration-damping platform; the first displacement adjustment knob is located on the first displacement platform, and the second displacement adjustment knob is located on the second displacement platform; the sensitivity calibration optical fibers at both ends of the sensitivity calibration grating are respectively fixed at a first fixed point on the first displacement platform and a second fixed point on 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 inside the heat-insulating ceramic chamber, and is located at the midpoint between the first fixed point and the second fixed point; the heat-insulating ceramic chamber is placed on the constant temperature heating stage, and the acquisition end of the sensitivity calibration thermocouple is located inside the heat-insulating ceramic chamber;

[0013] 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.

[0014] When measuring the strain sensitivity coefficient of the sensitivity calibration grating at a preset temperature, at the preset temperature, first adjust the distance between the first fixed point and the second fixed point to make the sensitivity calibration fiber set in a straight line, then control the first displacement platform or the second displacement platform to move a preset distance multiple times to stretch the sensitivity calibration fiber, and record the strain and wavelength of the sensitivity calibration grating after each movement, fit the wavelength of the sensitivity calibration grating with a first linear function of strain, and determine the slope of the first linear function 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 without stretching are selected, and a second linear function of the wavelength of the sensitivity calibration grating with respect to temperature is fitted. The slope of the second linear function is then determined to be the temperature sensitivity coefficient.

[0016] The sensitivity calibration grating is identical in shape, size, and material to the grating in the fiber optic strain sensing network module and the temperature compensation grating.

[0017] Optionally, the thermal stress detection device for the aircraft component further includes an adhesive component;

[0018] The adhesive assembly includes an auxiliary adhesive mold and an adhesive. The auxiliary adhesive mold is used to fix the shape and volume of the adhesive, and the adhesive is used to attach the grating in the fiber optic strain sensing network module to the preset detection location.

[0019] Optionally, the thermal stress detection device further includes: support fasteners;

[0020] The support components include a wooden support for supporting the aircraft components.

[0021] Optionally, the environmental chamber includes a temperature monitoring module and an environmental chamber thermocouple;

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

[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 includes a real-time strain display unit. The grating number on the main interface of the real-time strain display unit corresponds one-to-one with the grating number in the fiber optic grating strain sensing network module. It is used to display the original wavelength of the grating in the fiber optic grating strain sensing network module and the strain value monitored by the grating in real time.

[0025] Secondly, embodiments of the present invention also provide a method for detecting thermal stress in aircraft components, implemented using the thermal stress detection device for aircraft components as described in the first aspect, the method comprising:

[0026] Fiber Bragg grating strain sensing network modules are designed and manufactured according to the preset detection locations;

[0027] The fiber grating strain sensing network module is attached to each of the preset detection locations to form a test specimen.

[0028] The test specimen was placed in an environmental chamber;

[0029] Determine the strain force on the grating in the fiber Bragg grating strain sensing network module.

[0030] Optionally, after designing and manufacturing the fiber Bragg grating strain sensing network module according to the preset detection location, the method further includes:

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

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

[0033] When measuring the strain sensitivity coefficient of the sensitivity calibration grating at a preset temperature, at the preset temperature, first adjust the distance between the first fixed point and the second fixed point to be equal to that of the sensitivity calibration fiber, then move the first displacement platform or the second displacement platform by a preset distance multiple times to stretch the sensitivity calibration fiber, and record the strain and wavelength of the sensitivity calibration grating after each movement, fit the wavelength of the sensitivity calibration grating as a first linear function of the strain, and determine the slope of the first linear function as the strain sensitivity coefficient;

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

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

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

[0037] The thermal stress on the grating in the fiber optic strain sensing network module is obtained based on the center wavelength offset of the grating and the temperature-compensated grating in the fiber optic strain sensing network module, the strain sensitivity coefficient, the temperature sensitivity coefficient, and the temperature change.

[0038] This invention provides a thermal stress detection device and method for aircraft components. The aircraft components include multiple hybrid structures formed by interconnecting composite materials and metal materials. Each hybrid structure includes multiple preset detection locations. The thermal stress detection device includes: a fiber optic strain sensing network module, a temperature compensation module, an environmental chamber, and a grating strain demodulator. The fiber optic strain sensing network module includes multiple unidirectional strain measurement grating units and multiple tridirectional strain flower grating units. Each unidirectional strain measurement grating unit includes a unidirectional strain measurement grating. Each tridirectional strain flower grating unit includes a strain flower 0° direction strain measurement grating disposed on the same optical fiber. The test specimen consists of a strain gauge grating with a 45° and a 90° direction; a unidirectional strain gauge grating, a 0° direction strain gauge grating, a 45° direction strain gauge grating, and a 90° direction strain gauge grating, each attached to a pre-set detection location. The grating strain demodulator is also connected to the optical fiber in the fiber optic strain sensing network module. The temperature compensation module includes a temperature compensation grating located in a strain-independent part of the fiber optic strain sensing network module. The fiber optic strain sensing network module, the aircraft component, and the temperature compensation module are all located within an environmental chamber. This embodiment of the invention can meet the requirements of thermal stress testing for large aircraft structures. The use of a fiber optic strain sensing network module simplifies installation, and the temperature compensation grating, located in a strain-independent part of the module, minimizes the impact of temperature and stress on the accuracy of thermal stress measurements.

[0039] It should be understood that the description in this section is not intended to identify key or essential features of the embodiments of the present invention, nor is it intended to limit the scope of the invention. Other features of the invention will become readily apparent from the following description. Attached Figure Description

[0040] To more clearly illustrate the technical solutions in the embodiments of the present invention, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0041] Figure 1 This is a schematic diagram of a hybrid structure provided in an embodiment of the present invention;

[0042] Figure 2 This is a schematic diagram of another hybrid structure provided in an embodiment of the present invention;

[0043] Figure 3 This is a schematic diagram of another hybrid structure provided in an embodiment of the present invention;

[0044] Figure 4 This is a schematic diagram of another hybrid structure provided in an embodiment of the present invention;

[0045] Figure 5 This is a schematic diagram of the structure of a thermal stress detection device for an aircraft component provided in an embodiment of the present invention;

[0046] Figure 6 This is a schematic diagram of the structure of a unidirectional strain measurement grating unit provided in an embodiment of the present invention;

[0047] Figure 7 This is a schematic diagram of the structure of a three-directional strain measurement grating unit provided in an embodiment of the present invention;

[0048] Figure 8 This is a schematic diagram of the structure of a sensitivity calibration module provided in an embodiment of the present invention;

[0049] Figure 9 A schematic diagram illustrating the relationship between the wavelength and strain of a sensitivity calibration grating at various temperatures, provided as an embodiment of the present invention;

[0050] Figure 10 A schematic diagram illustrating the relationship between wavelength and temperature of a sensitivity calibration grating, provided as an embodiment of the present invention;

[0051] Figure 11 A flowchart of a method for detecting thermal stress in aircraft components provided by an embodiment of the present invention;

[0052] Figure 12 A flowchart of another method for detecting thermal stress in aircraft components provided in an embodiment of the present invention. Detailed Implementation

[0053] To enable those skilled in the art to better understand the present invention, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings of the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort should fall within the scope of protection of the present invention.

[0054] It should be noted that the terms "first," "second," etc., in the specification, claims, and accompanying drawings of this invention are used to distinguish similar objects and are not necessarily used to describe a specific order or sequence. It should be understood that such data can be interchanged where appropriate so that the embodiments of the invention described herein can be implemented in orders other than those illustrated or described herein. Furthermore, the terms "comprising" and "having," and any variations thereof, are intended to cover non-exclusive inclusion; for example, a process, method, system, product, or apparatus that comprises a series of steps or units is not necessarily limited to those steps or units explicitly listed, but may include other steps or units not explicitly listed or inherent to such processes, methods, products, or apparatus.

[0055] Figure 1 This is a schematic diagram of a hybrid structure provided in an embodiment of the present invention, specifically a schematic diagram of the hybrid structure at room temperature. (Reference) Figure 1 The hybrid structure includes a composite material 1, a metal material 2, and a metal fastener 3, with the composite material 1 and the metal material 2 being fixedly connected by the metal fastener 3. Figure 2 This is a schematic diagram of another hybrid structure provided in an embodiment of the present invention, specifically a schematic diagram of the hybrid structure under excessively low temperature conditions. Figure 3 This is a schematic diagram of another hybrid structure provided in an embodiment of the present invention, specifically a schematic diagram of the hybrid structure under excessively high temperature conditions, see reference. Figure 2 and Figure 3 Both excessively low and excessively high temperatures can cause deformation of the hybrid structure. Figure 4 This is a schematic diagram of another hybrid structure provided in an embodiment of the present invention, with reference to... Figure 4 The hybrid structure, composed of composite material 1 and metal material 2, is subjected to localized stress changes in metal material 2 due to the overall thermal effect of the aircraft structure, resulting in overall thermal stress.

[0056] It is understandable that aircraft are assembled at room temperature, but operate under varying high and low temperature conditions. The significant difference in thermal properties between composite materials and metallic materials generates thermal stress, and this thermal stress at some critical locations has a non-negligible impact on the static strength and fatigue life of the structure. The local thermal stress of this hybrid structure can be defined as the internal deformation and internal stress in the hybrid connection area caused by the mismatch in thermal expansion coefficients and temperature effects. To meet the requirements of thermal stress testing for large aircraft structures, this invention provides a thermal stress detection device and method for aircraft components, as detailed below:

[0057] The aircraft component of this invention includes multiple hybrid structures made of composite materials and metal materials connected together, and the hybrid structures include multiple preset detection areas. Figure 5This is a schematic diagram of the structure of a thermal stress detection device for an aircraft component provided in an embodiment of the present invention, with reference to... Figure 5 The thermal stress detection device 100 includes: a fiber optic strain sensing network module 110, a temperature compensation module 120, an environmental chamber 130, and a grating strain demodulator 140.

[0058] Figure 6 This is a schematic diagram of the structure of a unidirectional strain measurement grating unit provided in an embodiment of the present invention. Figure 7 This is a schematic diagram of a three-directional strain measurement grating unit provided in an embodiment of the present invention. (Reference) Figure 6-7 In this embodiment of the invention, the fiber optic strain sensing network module 110 includes multiple unidirectional strain measurement grating units 111 and multiple tridirectional strain flower grating units 112. The unidirectional strain measurement grating unit 111 includes a unidirectional strain measurement grating 1111. The tridirectional strain flower grating unit 112 includes 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 disposed on the same optical fiber 113.

[0059] In this embodiment of the invention, a unidirectional strain measurement grating 1111, a strain rosette strain measurement grating 1121 in the 0° direction, a strain rosette strain measurement grating 1122 in the 45° direction, and a strain rosette strain measurement grating 1123 in the 90° direction are respectively attached to each preset detection location to form a test piece.

[0060] It should be noted that by analyzing and evaluating areas of high stress in aircraft components under target thermal loads, key locations requiring inspection are identified, i.e., pre-set inspection areas. Generally, the thermal stress is more complex in hybrid structures formed by the interconnection of composite and metal materials, requiring a higher density of monitoring points. In areas without hybrid structures or with simple structures, fewer monitoring points are set. The fiber optic cabling layout, grating length, and spacing are designed based on the pre-set inspection areas. The number of gratings connected in series in a single fiber is determined based on the number of monitoring points, demodulation strategy, and thus the number of channels. The fiber optic grating array is fabricated in one step using dynamic online grating technology, and a carbon coating process ensures the fiber optic gratings possess high-temperature strain, temperature testing stability, and accuracy.

[0061] For example, based on finite element analysis, a large aircraft structural component requires 240 monitoring points (72 single-piece and 56 multi-piece) for strain monitoring. Based on the density and number of monitoring points, a fiber optic grating sensor network with a minimum spatial resolution of 50 mm, a grating length of 50 mm, and 32 channels is selected. The corresponding grating array is fabricated in one step using a dynamic online grating fabrication technique, and a carbon coating process enables the fiber optic grating to perform high-temperature testing.

[0062] In this embodiment of the invention, the grating strain demodulator 140 is connected to the optical fiber in the fiber optic strain sensing network module 110; the temperature compensation module 120 includes a temperature compensation grating, which is located in a part of the fiber optic strain sensing network module 110 that is not affected by strain; the fiber optic strain sensing network module 110, the aircraft component, and the temperature compensation module 120 are all located inside the environmental chamber 130.

[0063] It is understandable that the temperature-compensated grating is suspended to ensure no interaction force between it and the test piece. By adding several unloaded temperature-compensated gratings, the fiber optic grating sensor, when used as a strain measurement sensor, can counteract the effect of ambient temperature changes on its center wavelength shift, thereby achieving the purpose of strain measurement. For the grating attached to the preset detection location, when the test piece is subjected to thermal load, the strain measurement grating attached to the surface of the aircraft component will be simultaneously subjected to temperature and thermal strain. When the temperature changes by ΔT, its center wavelength shift can be expressed as: Where Δλ1 is the center wavelength offset of the grating attached to the aircraft component, λ1 is the initial center wavelength of the grating attached to the aircraft component, K1 is the strain sensitivity coefficient, K2 is the temperature sensitivity coefficient, ε is the thermal strain of the grating attached to the aircraft component, and ΔT is the temperature change.

[0064] For a temperature-compensated grating, since it is not subjected to any force, when the temperature change is ΔT, its center wavelength shift can be expressed as: Where Δλ2 is the center wavelength offset 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 experienced by the grating attached to the aircraft component is:

[0066] The strain to be measured can be determined by the center wavelength offset, initial center wavelength, and strain sensitivity coefficient of the grating attached to the aircraft component and the temperature compensation grating in a state of only sensing temperature and not stress.

[0067] This invention, through a rational fiber optic sensor network design, attaches a unidirectional strain measurement grating unit 111 or a three-directional strain rosette grating unit 112 to a preset detection location on the aircraft component. The temperature within the environmental chamber 130 is controlled to apply the required temperature load for the test. A temperature compensation module 120 processes the temperature influence of the grating strain monitoring, and a grating strain demodulator 140 analyzes the strain at the preset detection location, monitoring the strain change with temperature in real time. The thermal stress test of the test piece is completed according to the required temperature change conditions. This method can meet the thermal stress testing requirements of large aircraft structures. The use of the fiber optic grating strain sensor network module 110 is simple to install, and the temperature compensation grating is located in a strain-independent part of the fiber optic grating strain sensor network module 110, minimizing the influence of temperature and stress on the accuracy of thermal stress measurement.

[0068] Figure 8 This is a schematic diagram of a sensitivity calibration module provided in an embodiment of the present invention. Optionally, based on the above embodiment, refer to... Figure 8 The thermal stress detection device for the aircraft component also includes a sensitivity calibration module 150. The sensitivity calibration module 150 includes a vibration-damping platform 151, a constant-temperature heating stage 152, a first displacement platform 153, a second displacement platform 154, a first displacement adjustment knob 155, a second displacement adjustment knob 156, a heat-insulating ceramic chamber 157, a sensitivity calibration fiber optic cable 158, a sensitivity calibration grating 159, and a sensitivity calibration thermocouple 1510. In this embodiment of the invention, the grating strain demodulator 140 is connected to the output terminals of the sensitivity calibration fiber optic cable 158 and the sensitivity calibration thermocouple 1510.

[0069] Continue to refer to Figure 8The constant temperature heating stage 152, the first displacement platform 153, and the second displacement platform 154 are all located on the vibration-proof 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 fibers 158 at both ends of the sensitivity calibration grating 159 are respectively fixed at 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 inside the heat-insulating ceramic chamber 157, and is located at the middle position between the first fixed point A and the second fixed point B; the heat-insulating ceramic chamber 157 is placed on the constant temperature heating stage 152. In this embodiment of the invention, the acquisition end of the sensitivity calibration thermocouple 1510 is located inside the heat-insulating ceramic chamber 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; 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 stage 152 is used to control the temperature of the environment in which the sensitivity calibration grating 159 is located.

[0070] In this embodiment of the 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 to make the sensitivity calibration fiber 158 linearly set. Then, by repeatedly controlling the first displacement platform 153 or the second displacement platform 154 to move a preset distance, the sensitivity calibration fiber 158 is stretched, and the strain and wavelength of the sensitivity calibration grating 159 after each movement are recorded. A first linear function of the wavelength of the sensitivity calibration grating 159 with respect to strain is fitted, and the slope of the first linear function is determined to be the strain sensitivity coefficient. When measuring the temperature sensitivity coefficient of the sensitivity calibration grating 159, wavelength data of the sensitivity calibration grating 159 under different temperatures without stretching are selected, and a second linear function of the wavelength of the sensitivity calibration grating 159 with respect to temperature is fitted. The slope of the second linear function is determined to be the temperature sensitivity coefficient.

[0071] Specifically, Figure 9 This is a schematic diagram illustrating the relationship between the wavelength of a sensitivity calibration grating and strain at various temperatures, provided as an embodiment of the present invention. Figure 10 This is a schematic diagram illustrating the relationship between the wavelength of a sensitivity calibration grating and temperature, provided as 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 in steps of a preset distance Δd, the strain at each step is equal to n·Δd / D, where n is the number of movements. The strain at each step is recorded in relation to the wavelength of the sensitivity calibration grating 159, and a first-order linear function of the wavelength of the sensitivity calibration grating 159 with respect to strain is fitted. (Refer to...) Figure 9 The slope of the first linear function of the wavelength of the sensitivity calibration grating 159 with respect to strain is the strain sensitivity coefficient K1 of the sensitivity calibration grating 159 at that temperature, i.e., λ = K1·ε, where λ is the wavelength of the sensitivity calibration grating 159 and ε is the strain. This embodiment of the invention measured the strain sensitivity coefficient of the sensitivity calibration grating 159 at different temperatures. This embodiment of the invention, by selecting the wavelength data of the sensitivity calibration grating 159 under the condition that the sensitivity calibration fiber 158 is not stretched at each temperature, fits a second linear function of the wavelength of the sensitivity calibration grating 159 with respect to temperature, 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. (Reference) Figure 10 The slope of the linear function is the temperature sensitivity coefficient of the sensitivity calibration grating 159.

[0072] It is understandable that the strain sensitivity coefficient of a grating usually does not change much under different temperature conditions, so the average value of each working condition is taken as the strain sensitivity coefficient of the grating.

[0073] The sensitivity calibration grating 159 is identical in shape, size, and material to the grating and temperature compensation grating in the fiber optic strain sensing network module 110.

[0074] Optionally, based on the above embodiments, continue to refer to... Figure 6 and Figure 7 The thermal stress detection device 100 for aircraft components also includes an adhesive assembly 160. The adhesive assembly 160 includes an auxiliary adhesive mold 161 and an adhesive 162. The auxiliary adhesive mold 161 is used to fix the shape and volume of the adhesive 162, and the adhesive 162 is used to attach the grating in the fiber optic strain sensing network module 110 to the preset detection location.

[0075] Specifically, to ensure uniform sensitivity across all gratings after attachment, an auxiliary attachment mold 161 is used to fix the shape and volume of the adhesive 162. The adhesive 162 in this embodiment of the invention has high-temperature resistance, ensuring the stability of the grating position and strain transfer throughout the high-temperature test. Based on this attachment method, gratings are attached along three different directions to form a three-directional strain rosette grating unit 112.

[0076] For example, adjust the fiber optic direction along the measurement direction so that the center of the grating is aligned with the preset detection location. Apply a certain prestress to the fiber optic grating. Fix both ends of the grating with adhesive 162. Place the auxiliary pasting mold 161 around the grating, positioning the grating in the center of the rectangular hole. Fix the edges of the auxiliary pasting mold 161 with adhesive 162. Apply adhesive 162 to the rectangular hole in the center of the auxiliary pasting mold 161. The adhesive 162 must be a high-temperature resistant adhesive to maintain adhesion throughout the test without deformation or displacement. Ensure the adhesive is evenly distributed and applied thinly to minimize the impact of adhesive non-uniformity in high and low temperature environments. After the adhesive 162 has completely cured, remove the auxiliary pasting mold 161 to complete the grating pasting for one strain monitoring point. For monitoring points requiring monitoring in three directions, use the same method, referring to... Figure 7 The fiber Bragg gratings are pasted sequentially in three directions. When multiple fiber Bragg gratings are connected in series, care should be taken to ensure that the fiber optic patch cords are fixed in appropriate positions to prevent fiber breakage during specimen movement. If fiber optic loops exist, the inner diameter of the loops should be increased as much as possible to reduce optical loss and avoid interruption of the grating signal.

[0077] Optionally, based on the above embodiments, the thermal stress detection device 100 further includes: a support component. The support component includes a wooden support for supporting aircraft components.

[0078] It is understandable that using fasteners less affected by temperature, such as wooden supports, to support the test specimens allows the specimens to expand freely under temperature stress. The specific support locations and wooden support profiles can be designed according to actual conditions.

[0079] Optionally, based on the above embodiments, the environmental chamber 130 includes a temperature monitoring module and an environmental chamber thermocouple. The temperature monitoring module is used to regulate the temperature inside the environmental chamber 130, and the environmental chamber thermocouple is used to detect the temperature inside the environmental chamber 130.

[0080] In this embodiment of the invention, the grating strain demodulator 140 is first zeroed, and initial data is recorded. The temperature monitoring module of the environmental chamber 130 is adjusted to raise the temperature until the temperature of all environmental chamber thermocouples in the environmental chamber 130 is displayed one by one to the target temperature. After each target temperature stabilizes, it is held for 15 minutes, and strain data is measured and collected. Then, the temperature of the environmental chamber 130 is adjusted to start cooling down until the temperature of all environmental chamber thermocouples in the environmental chamber 130 is displayed one by one to the target temperature of the cooling stage, until room temperature. After each target temperature stabilizes, it is held for 15 minutes, and strain data is measured and collected. It is recommended that the heating rate of the test equipment be 1-5℃ / min, the cooling rate be 1-3℃ / min, and the maximum overshoot temperature be set to ≤5℃.

[0081] Optionally, based on the above embodiments, 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, based on the above embodiments, the grating strain demodulator 140 includes a real-time strain display unit. The grating number on the main interface of the real-time strain display unit corresponds one-to-one with the grating number in the fiber optic grating strain sensing network module 110. It is used to display the original wavelength of the grating in the fiber optic grating strain sensing network module 110 and the strain value monitored by the grating in real time.

[0083] Understandably, the grating strain demodulator 140 has a real-time strain display function, and the host computer sub-interface can display the grating strain curve over time and update it in real time.

[0084] It should be noted that the monitoring and recording of grating spectra can be achieved through a spectrometer, wavelength division and time division sensor demodulators, or other spectrometer monitoring instruments. The number of channels required should be designed based on the number of measurement points, the wavelength resolution of the demodulator, and the demodulation rate.

[0085] Compared to single-point measurements using thermocouples and strain gauges, which involve numerous wiring connections, the thermal stress detection device for aircraft components provided in this invention utilizes fiber optic gratings (FGFs). FGFs can achieve dense grating sensors on a single fiber through various multiplexing methods, simplifying data acquisition channels. Furthermore, FGF sensors offer advantages such as small size, light weight, corrosion resistance, high temperature resistance, and electromagnetic interference resistance. The grating sensitivity testing device of this invention can simultaneously measure the high-temperature strain sensitivity coefficient and temperature sensitivity coefficient of the grating. The grating bonding method provided ensures the stability of the grating position and strain transfer throughout the high-temperature test. This invention can be used to monitor the mechanical state of structures under environmental thermal loads, enabling thermal stress testing of large hybrid composite and metal structures. Simultaneously, this method can be used under cyclic thermal loads to evaluate the residual strength and fatigue performance of structures under cyclic thermal loads. Therefore, it provides methods and conditions for verifying the implementation of thermal stress tests on composite and metal hybrid structures under environmental loads, and can also be used to verify and revise structural thermal stress analysis methods.

[0086] Figure 11 This is a flowchart of a method for detecting thermal stress in aircraft components according to an embodiment of the present invention. It is implemented using the thermal stress detection device 100 for aircraft components provided in the above embodiment. (Refer to...) Figure 11 Methods for detecting thermal stress in aircraft components include:

[0087] S210. Design and manufacture fiber optic strain sensing network modules according to preset detection locations.

[0088] S220. The grating in the fiber optic strain sensor network module is attached to each preset detection part to form a test piece.

[0089] S230. Place the test specimen in the environmental chamber.

[0090] S240. Determine the strain force on the grating in the fiber Bragg grating strain sensing network module.

[0091] Figure 12 The flowchart illustrates another method for detecting thermal stress in aircraft components provided by an embodiment of the present invention. Optionally, based on the above embodiments, refer to... Figure 12 The method includes:

[0092] S310. Design and manufacture fiber optic strain sensing network modules according to preset detection locations.

[0093] S320. Determine the strain sensitivity coefficient and temperature sensitivity coefficient of the sensitivity calibration grating at different temperatures.

[0094] S330. Attach the grating in the fiber optic strain sensor network module to each preset detection location to form a test specimen.

[0095] S340. Place the test specimen in the environmental chamber.

[0096] S350. Determine the strain force on the grating in the fiber Bragg grating strain sensing network module.

[0097] Optionally, based on the above embodiments, step S320 includes: when measuring the strain sensitivity coefficient of the sensitivity calibration grating at a 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 repeatedly control the first displacement platform or the second displacement platform to move the preset distance to stretch the sensitivity calibration fiber, and record the strain and wavelength of the sensitivity calibration grating after each movement, fit a first linear function of the wavelength of the sensitivity calibration grating with respect to 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 a second linear function of the wavelength of the sensitivity calibration grating with respect to temperature, and determine the slope of the second linear function as the temperature sensitivity coefficient.

[0098] Optionally, based on the above embodiments, step S350 includes: determining the center wavelength offset of the grating and the temperature-compensated grating in the fiber Bragg grating strain sensing network module. The thermal stress on the grating in the fiber Bragg grating strain sensing network module is obtained based on the center wavelength offset, strain sensitivity coefficient, temperature sensitivity coefficient, and temperature change of the grating and the temperature-compensated grating in the fiber Bragg grating strain sensing network module.

[0099] The thermal stress detection method for aircraft components provided in this embodiment of the invention is implemented using the thermal stress detection device for aircraft components provided in the above embodiment, and therefore has the same beneficial effects. For contents not described in detail in this embodiment of the invention, please refer to the thermal stress detection device for aircraft components provided in the above embodiment.

[0100] It should be understood that the various forms of processes shown above can be used, with steps reordered, added, or deleted. For example, the steps described in this invention can be executed in parallel, sequentially, or in different orders, as long as the desired result of the technical solution of this invention can be achieved, and this is not limited herein.

[0101] The specific embodiments described above do not constitute a limitation on the scope of protection of this invention. 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 modifications, equivalent substitutions, and improvements made within the spirit and principles of this invention should be included within the scope of protection of this invention.

Claims

1. A thermal stress detection device for aircraft components, characterized in that, The aircraft component includes multiple hybrid structures made of composite materials and metal materials, and the hybrid structure includes multiple preset detection points; the thermal stress detection device includes: a fiber optic strain sensing network module, a temperature compensation module, an environmental chamber, and a grating strain demodulator. The fiber optic strain sensing network module includes multiple unidirectional strain measurement grating units and multiple tridirectional strain flower grating units. The unidirectional strain measurement grating unit includes a unidirectional strain measurement grating; the three-directional strain flower grating unit includes 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 disposed 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 each of the preset detection locations to form a test piece; The grating strain demodulator is connected to the optical fiber in the fiber optic strain sensing network module; the temperature compensation module includes a temperature compensation grating, which is located in a part of the fiber optic strain sensing network module that is not affected by strain; the fiber optic strain sensing network module, the aircraft component, and the temperature compensation module are all located inside the environmental chamber.

2. The thermal stress detection device for aircraft components according to claim 1, characterized in that, It also includes a sensitivity calibration module, which comprises a vibration-proof platform, a constant-temperature heating stage, a first displacement platform, a second displacement platform, a first displacement adjustment knob, a second displacement adjustment knob, a heat-insulating ceramic chamber, a sensitivity calibration fiber, a sensitivity calibration grating, and a sensitivity calibration thermocouple; the grating strain demodulator is connected to the output end of the sensitivity calibration fiber and the sensitivity calibration thermocouple. The constant temperature heating stage, the first displacement platform, and the second displacement platform are all located on the vibration-damping platform; the first displacement adjustment knob is located on the first displacement platform, and the second displacement adjustment knob is located on the second displacement platform; The sensitivity calibration optical fibers at both ends of the sensitivity calibration grating are fixed at a first fixed point on the first displacement platform and a second fixed point on the second displacement platform, respectively, with the first and second fixed points at the same horizontal position. The sensitivity calibration grating is located inside the heat-insulating ceramic chamber and is positioned between the first and second fixed points. The heat-insulating ceramic chamber is placed on the constant-temperature heating platform, and the acquisition end of the sensitivity calibration thermocouple is located inside the heat-insulating ceramic chamber. 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 adjust the distance between the first fixed point and the second fixed point to make the sensitivity calibration fiber set in a straight line, then control the first displacement platform or the second displacement platform to move a preset distance multiple times to stretch the sensitivity calibration fiber, and record the strain and wavelength of the sensitivity calibration grating after each movement, fit the wavelength of the sensitivity calibration grating with a first linear function of 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, wavelength data of the sensitivity calibration grating under different temperatures without stretching are selected, and a second linear function of the wavelength of the sensitivity calibration grating with respect to temperature is fitted. The slope of the second linear function is then determined to be the temperature sensitivity coefficient. The sensitivity calibration grating is identical in shape, size, and material to the grating in the fiber optic strain sensing network module and the temperature compensation grating.

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

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

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 inside the environmental chamber, and the environmental chamber thermocouple is used to detect the temperature inside the environmental chamber.

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 number on the main interface of the real-time strain display unit corresponds one-to-one with the grating number in the fiber optic grating strain sensing network module. It is used to display the original wavelength of the grating in the fiber optic grating strain sensing network module and the strain value monitored by the grating in real time.

7. A method for detecting thermal stress in aircraft components, characterized in that, The thermal stress detection device for aircraft components as described in any one of claims 1-6 is used, and the thermal stress detection method for the aircraft components includes: Fiber Bragg grating strain sensing network modules are designed and manufactured according to the preset detection locations; The fiber grating strain sensing network module is attached to each of the preset detection locations to form a test specimen. The test specimen was placed in an environmental chamber; Determine the strain force on the grating in the fiber Bragg grating strain sensing network module.

8. The method for detecting thermal stress in 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 location, 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 in aircraft components according to claim 8, characterized in that, The determination of 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 adjust the distance between the first fixed point and the second fixed point to be equal to that of the sensitivity calibration fiber, then move the first displacement platform or the second displacement platform by a preset distance multiple times to stretch the sensitivity calibration fiber, and record the strain and wavelength of the sensitivity calibration grating after each movement, fit the wavelength of the sensitivity calibration grating as a first linear function of 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, wavelength data of the sensitivity calibration grating under the condition that the sensitivity calibration fiber is not stretched are selected at each temperature, and a second linear function of the wavelength of the sensitivity calibration grating with respect to temperature is fitted. The slope of the second linear function is then determined to be the temperature sensitivity coefficient.

10. The method for detecting thermal stress in aircraft components according to claim 8, characterized in that, The determination of the strain force on the grating in the fiber Bragg grating strain sensing network module includes: Determine 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 optic strain sensing network module is obtained based on the center wavelength offset of the grating and the temperature-compensated grating in the fiber optic strain sensing network module, the strain sensitivity coefficient, the temperature sensitivity coefficient, and the temperature change.

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