Spacecraft structure micro stress detection device and use method thereof

Abstract

The invention relates to the technical field of spacecraft structure micro stress detection, in particular to a spacecraft structure micro stress detection device and a use method thereof, and the spacecraft structure micro stress detection device comprises a displacement amplification structure, a single mode fiber, a reflector, a broadband light source and a spectrum analyzer. An optical resonant cavity formed by an optical fiber end face and a lens is combined with a displacement amplification structure, when stress acts on a spacecraft structure, the cavity length of the Fabry-Perot resonant cavity changes, so that phase movement of light is caused, and accurate measurement of the stress can be achieved by detecting the phase movement. The antenna has extremely high anti-electromagnetic interference capability, and can work stably in a complex space environment. Even under the condition of violent temperature change, high measurement accuracy can be kept, and detection of tiny stress can be achieved through the amplification effect of the displacement amplification structure. And the stress characteristic signal and the background noise can be effectively separated, so that the detection reliability is remarkably improved.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of spacecraft structure micro stress detection, and particularly relates to a spacecraft structure micro stress detection device and a use method thereof. BACKGROUND

[0002] The spacecraft structure is affected by various stresses in the complex space environment, such as vibration stress in the launch process, thermal stress in the on-orbit operation, and space debris impact stress. Accurate detection of the stress suffered by the spacecraft structure is crucial to ensure the safe operation of the spacecraft. However, the traditional stress detection method has some limitations and cannot accurately detect the extremely small stress suffered by the spacecraft structure. SUMMARY

[0003] The present application provides a spacecraft structure micro stress detection device and a use method thereof to solve the problem that the traditional stress detection method has some limitations and cannot accurately detect the extremely small stress suffered by the spacecraft structure.

[0004] In one aspect, the present application provides a spacecraft structure micro stress detection device, comprising:

[0005] A displacement amplification structure is fixedly installed on one side of the spacecraft structure at opposite ends;

[0006] A single-mode optical fiber is installed on one side of the displacement amplification structure;

[0007] A reflecting mirror is installed on the other side of the displacement amplification structure and is arranged opposite to one end surface of the single-mode optical fiber. The reflecting mirror is close to the single-mode optical fiber, and a Fabry-Perot resonant cavity is formed between the side surface of the single-mode optical fiber and the end surface of the single-mode optical fiber;

[0008] A broadband light source is connected to the single-mode optical fiber through an optical fiber coupler;

[0009] A spectrum analyzer is connected to the single-mode optical fiber through the optical fiber coupler;

[0010] When the spacecraft structure is subjected to stress, the distance between the opposite ends of the displacement amplification structure changes, and the cavity distance of the Fabry-Perot resonant cavity changes.

[0011] In some embodiments, the displacement amplification structure is in the shape of a rhombus, comprising:

[0012] A first fixed seat is fixedly installed on the spacecraft structure;

[0013] A second fixed seat is arranged opposite to the first fixed seat and is fixedly installed on the spacecraft structure;

[0014] The fiber mounting seat is arranged between the first fixing seat and the second fixing seat, and a clearance hole is formed in the fiber mounting seat;

[0015] The sleeve is installed in the clearance hole and can allow the end of the single-mode optical fiber to pass through.

[0016] The lens adjusting seat is arranged between the first fixing seat and the second fixing seat, and a reflecting lens is installed on the side of the lens adjusting seat close to the fiber mounting seat.

[0017] The first connecting arm is fixedly connected at one end to the first fixing seat and at the other end to the fiber mounting seat.

[0018] The second connecting arm is fixedly connected at one end to the second fixing seat and at the other end to the fiber mounting seat.

[0019] The third connecting arm is fixedly connected at one end to the first fixing seat and at the other end to the lens adjusting seat.

[0020] The fourth connecting arm is fixedly connected at one end to the second fixing seat and at the other end to the lens adjusting seat.

[0021] In some embodiments, the lens adjusting seat comprises:

[0022] The first lens mounting seat is fixedly connected to the third connecting arm and the fourth connecting arm, respectively.

[0023] The second lens mounting seat is installed on the side of the first lens mounting seat close to the fiber mounting seat.

[0024] The elastic washer is installed between the second lens mounting seat and the first lens mounting seat.

[0025] The adjusting screw is used to connect the second lens mounting seat, the elastic washer and the first lens mounting seat.

[0026] In some embodiments, the first fixing seat comprises:

[0027] The fixing blocks are two and are connected by the connecting screw, one of the fixing blocks is fixedly connected to the first connecting arm, and the other fixing block is fixedly connected to the third connecting arm.

[0028] The second fixing seat has the same structure as the first fixing seat.

[0029] In some embodiments, the first fixing seat and the second fixing seat are respectively bonded to the spacecraft structure.

[0030] In some embodiments, the clearance hole is a stepped hole.

[0031] In some embodiments, the reflecting lens is installed on the second lens mounting seat by bonding.

[0032] In some embodiments, the surface of the reflecting lens is plated with a gold layer.

[0033] In some embodiments, the length of the first connecting arm, the second connecting arm, the third connecting arm and the fourth connecting arm is 145mm, the arm thickness is 4mm, and the angle between the extension direction and the horizontal direction is 2°.

[0034] In another aspect, the application also provides a method for using the spacecraft structure micro stress detection device, comprising:

[0035] The two ends of the displacement amplification structure of the spacecraft structure micro stress detection device are fixedly installed on one side surface of the spacecraft structure;

[0036] The first preset pressure is applied to the spacecraft structure by the driver;

[0037] The second preset pressure is applied to the spacecraft structure by the driver.

[0038] The beneficial effects of the application are as follows: the spacecraft structure micro stress detection device of the application is provided with a displacement amplification structure, a single-mode optical fiber, a reflecting lens, a broadband light source and a spectrum analyzer. The opposite two ends of the displacement amplification structure are fixedly installed on one side surface of the spacecraft structure. The single-mode optical fiber is installed on one side of the displacement amplification structure. The reflecting lens is installed on the other side of the displacement amplification structure and is arranged opposite to one end surface of the single-mode optical fiber. A Fabry-Perot resonant cavity is formed between the side surface of the reflecting lens close to the single-mode optical fiber and the end surface of the single-mode optical fiber. The broadband light source is connected with the single-mode optical fiber through an optical fiber coupler. The spectrum analyzer is connected with the single-mode optical fiber through the optical fiber coupler. When the spacecraft structure is subjected to stress, the distance between the opposite two ends of the displacement amplification structure changes, and then the cavity distance of the Fabry-Perot resonant cavity changes. The optical resonant cavity formed by the fiber end surface and the lens is combined with the displacement amplification structure. When stress acts on the spacecraft structure, the cavity length of the Fabry-Perot resonant cavity changes, thereby causing the phase shift of light. The precise measurement of stress can be realized by detecting the phase shift. It has extremely high anti-electromagnetic interference capability and can work stably in the complex space environment without being affected by electromagnetic fluctuations. Secondly, it has low temperature sensitivity. Even under severe temperature change conditions, it can maintain high measurement accuracy. Through the amplification effect of the displacement amplification structure, the detection of micro stress can be realized. At the same time, the sensor has small volume, high sensitivity, good repeatability and high linearity. It can effectively separate the stress characteristic signal and the background noise, thereby significantly improving the reliability of detection. BRIEF DESCRIPTION OF DRAWINGS

[0039] Figure 1 is a structural schematic diagram of some specific embodiments of the spacecraft structure micro stress detection device of the application;

[0040] Figure 2 is Figure 1 Figure 2 is a schematic diagram of the combination structure of the displacement amplification structure and the reflecting mirror in the spacecraft structure micro stress detection device shown in Figure 1;

[0041] Figure 3 Figure 3 is a schematic diagram of the combination structure of the mirror adjusting seat and the reflecting mirror;

[0042] Figure 4 Figure 4 is a schematic diagram of the triangular linear amplification principle;

[0043] Figure 5 Figure 5 is a schematic diagram of the double peak method principle.

[0044] In the drawings, 111 is a first fixed seat; 1111 is a fixed block; 112 is a second fixed seat; 113 is an optical fiber mounting seat; 114 is a sleeve; 115 is a mirror adjusting seat; 1151 is a first mirror mounting seat; 1152 is a second mirror mounting seat; 1153 is an elastic washer; 1154 is an adjusting bolt; 116 is a first connecting arm; 117 is a second connecting arm; 118 is a third connecting arm; 119 is a fourth connecting arm; 120 is a single mode optical fiber; 130 is a reflecting mirror; 140 is a broadband light source; 150 is an optical fiber coupler; 160 is a spectrum analyzer; and 200 is a spacecraft structure. DETAILED DESCRIPTION

[0045] The technical solutions of the present application will be described clearly and completely below in conjunction with the embodiments. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative labor fall within the scope of protection of the present application.

[0046] As described in the background, the traditional stress detection method has some limitations and cannot accurately detect the extremely small stress suffered by the spacecraft structure.

[0047] To solve the above problems, refer to Figure 1 , Figure 2 , Figure 3 , Figure 4 and Figure 5On the one hand, the present invention provides a device for detecting micro-stress in a spacecraft structure, comprising a displacement amplification structure, a single-mode optical fiber 120, a reflective lens 130, a broadband light source 140, and a spectrum analyzer 160. The two opposite ends of the displacement amplification structure are fixedly mounted on the same side of the spacecraft structure 200. The single-mode optical fiber 120 is mounted on one side of the displacement amplification structure. The reflective lens 130 is mounted on the other side of the displacement amplification structure, facing one end face of the single-mode optical fiber 120. A Fabry-Perot resonant cavity is formed between the side face of the reflective lens 130 proximal to the single-mode optical fiber 120 and the end face of the single-mode optical fiber 120. The broadband light source 140 is connected to the single-mode optical fiber 120 via a fiber coupler 150. The spectrum analyzer 160 is connected to the single-mode optical fiber 120 via a fiber coupler 150. When the spacecraft structure 200 is subjected to stress, it should be noted that the direction of the stress is parallel to the side face of the spacecraft structure 200 and points from one end of the displacement amplification structure to the other end. The distance between the two opposite ends of the displacement amplification structure changes, thereby changing the cavity distance of the Fabry-Perot resonant cavity, converting the stress change into a change in the optical signal. The wavelength and light intensity information in the optical signal are then processed using equipment such as an optical spectrum analyzer 160 and a computer. The measured cavity length change is substituted into the theoretical model to obtain the measured stress magnitude.

[0048] The working process and principle of the spacecraft structure micro-stress detection device are as follows:

[0049] When performing stress detection, first, the opposite ends of the displacement amplification structure are fixedly installed on the same side of the spacecraft structure 200. When the detection area is subjected to stress, the distance between the opposite ends of the displacement amplification structure will change linearly with the stress change. According to the principle of triangular linear amplification, the opposite sides of the displacement amplification structure will linearly amplify the displacement. In this process, the broadband light source 140 outputs continuous broadband light, which enters the single-mode optical fiber 120 through the optical fiber coupler 150, and after being reflected by the reflective lens 130, it is coupled into the single-mode optical fiber 120 again. The output light is transmitted to the spectrum analyzer 160 through the optical fiber coupler 150, and the reflection spectrum is monitored by the spectrum analyzer 160. The change in the distance between the opposite sides of the displacement amplification structure will cause a corresponding change in the reflection spectrum. By analyzing the change in the reflection spectrum, the cavity length change of the Fabry-Perot resonant cavity can be demodulated in real time. Combined with analysis, calculation and test calibration, the stress signal and the cavity length change signal can be matched one by one, thereby achieving the purpose of high-precision measurement of the stress in the detection area. As Figure 1 As shown, it is assumed that the gauge length of the two opposite ends of the displacement amplification structure is L g , the initial cavity length of the Fabry-Perot resonant cavity is s, the change in the cavity length of the Fabry-Perot resonant cavity is Δs, and the strain in the detection area is ε, then:

[0050]

[0051] In order to optimize the quality of the interference spectrum, the lens adjusting seat 115 is arranged to solve the problem of poor quality of the interference spectrum caused by the angle deviation of the single-mode optical fiber 120 and the reflecting lens 130.

[0052] And using the principle of triangular linear amplification as shown in Figure 4 , assuming that a displacement of Δx is formed on the x-axis and a displacement of Δy is formed on the y-axis, and the amplification factor is A, then:

[0053] A = Δy / Δx

[0054] Assuming that the length of the first connecting arm 116 is l, the initial angle between the extension direction of the first connecting arm 116 and the horizontal direction is α, when a displacement of Δx is formed on the x-axis, the angle between the extension direction of the first connecting arm 116 and the horizontal direction is β, and the difference between β and α is θ, then:

[0055] Δx = lcosα-lcosβ

[0056] Δy = lsinβ-lsinθ

[0057] Therefore, we have:

[0058]

[0059] Different cavity lengths correspond to different spectral distributions, and different spectral distributions have completely different phases in the same section, that is, the phase has a strict correspondence with the cavity length. Since the spectral distribution is a combination of multiple single-wavelength information, its information quantity is greater than that of the light intensity output under the single-wavelength condition, and therefore, the demodulation accuracy can be improved by using phase demodulation. Using the double-peak method to obtain the cavity length is a simple and direct demodulation method. The principle is shown in Figure 5 .

[0060] From equation (1), the change of the cavity length Δs with the stress is obtained, and then the strain change can be obtained. Two wavelengths λ m and λ m+q at the maximum light intensity in the spectrum before and after the stress change can be selected to process the information (m and m+q are the interference orders corresponding to them, respectively), and then λ m , λ m+q and the cavity length s1 satisfy the following relationship:

[0061]

[0062] When the cavity length changes, the wavelengths corresponding to the peak values in the spectrum will also change, and we can obtain another pair of wavelengths λ m ' and λ m+q '. In this way, the changed cavity length s2 can be obtained.

[0063]

[0064] Therefore, the displacement change amount is:

[0065] Δs = |s2-s1| (4)

[0066] Substituting the cavity length change amount into formula (4), the strain amount to be measured can be obtained. This measurement method realizes strain measurement by wavelength instead of light intensity, so that the influence of light source light intensity fluctuation on measurement can be overcome. Therefore, higher measurement accuracy and better stability are achieved.

[0067] Since the displacement amplification structure is combined with the Fabry-Perot resonant cavity, the actual deformation amount of the aluminum alloy sample is amplified N times by the rhombic displacement amplification structure to the F-P cavity. Therefore, the actual strain value detected in formula (1) needs to be divided by the amplification multiple N.

[0068] Therefore, the strain value of the detection area is

[0069]

[0070] Where ΔL is the sample detection area change amount, N is the displacement amplification structure amplification multiple, and L is the sample detection area length.

[0071] The amplification ratio formula of the rhombic displacement amplification structure can be simplified as:

[0072]

[0073] Where H is the change amount of the Fabry-Perot cavity length, ΔL f is the sample detection area change amount, and θ is the included angle between the first connecting arm 116 and the horizontal line at the start of movement.

[0074] Therefore, it can be known that:

[0075]

[0076] Where σ is the test stress value under the applied load, Δs is the cavity length change value under the load, E is the elastic modulus of the structure material to be measured, L is the distance between the two fixed points of the displacement amplification structure; and N is the amplification multiple.

[0077] By using the stress-sensitive characteristics of the fiber Fabry-Perot sensor, the stress change is converted into the change of the optical signal, and then the wavelength and light intensity information in the optical signal are processed by a demodulator and a computer to obtain the cavity length information of the Fabry-Perot cavity. Finally, the measured cavity length change is substituted into the theoretical model to obtain the measured stress value. The stress detection model formula is:

[0078]

[0079] wherein, a is a stress correction coefficient, calibrated by test. σ' is the test stress value under the action of the corrected load. Δs is the change value of the cavity length of the Fabry-Perot resonant cavity. E is the elastic modulus of the material to be tested, about 70 GPa. N is the magnification, about 16.92 times. L is the distance between the two ends of the displacement amplification structure, the value is 115 mm.

[0080] Overall, the displacement amplification structure is combined with the Fabry-Perot resonant cavity, the displacement amplification structure is used to convert the micro / nano stress change into significant displacement change, and the Fabry-Perot resonant cavity is used to realize high-precision optical measurement, which significantly improves the detection sensitivity and precision. It has the advantages of strong anti-electromagnetic interference ability, little influence of stress detection on temperature, high measurement precision and convenient installation, and can effectively adapt to the complex scene of the spacecraft structure 200. In addition, through calculation analysis and test calibration, the stress information and the cavity length change information are one-to-one corresponding, realizing high-precision, real-time online measurement of the stress of the detection area. Compared with the traditional strain gauge detection method, it is not affected by electromagnetic interference and has low temperature sensitivity, and can maintain high measurement precision in complex space environment. Compared with the traditional ultrasonic guided wave detection method, the signal interference problem caused by the scattering of acoustic waves in the complex structure is avoided, the clarity and reliability of the detection signal are improved, the lower limit of stress detection is reduced, and the strain of 1.41×10 -9 mm can be effectively identified, and the cavity length change is 3nm. At the same time, the spacecraft structure micro stress detection device has small overall volume, high sensitivity, good repeatability and high linearity, which can effectively separate the stress characteristic signal and the background noise, thereby significantly improving the detection reliability.

[0081] Preferably, as Figure 1 , Figure 2 and Figure 3As shown, the displacement amplification structure includes a first fixing seat 111, a second fixing seat 112, a fiber mounting seat 113, a sleeve 114, a lens adjustment seat 115, a first connecting arm 116, a second connecting arm 117, a third connecting arm 118, and a fourth connecting arm 119. The first fixing seat 111 is fixedly mounted on the spacecraft structure 200. The second fixing seat 112 is positioned opposite the first fixing seat 111 and is fixedly mounted on the spacecraft structure 200. The fiber mounting seat 113 is positioned between the first fixing seat 111 and the second fixing seat 112. A clearance hole is formed in the fiber mounting seat 113. The sleeve 114 is installed in the clearance hole, allowing the end of the single-mode optical fiber 120 to pass through. The lens adjustment seat 115 is positioned between the first fixing seat 111 and the second fixing seat 112 and is used to fine-tune the relative angle and distance between the reflective lens 130 and the single-mode optical fiber 120. The reflective lens 130 is mounted on a side of the lens adjustment seat 115 near the fiber mounting seat 113. One end of the first connecting arm 116 is fixedly connected to the first fixing seat 111, and the other end is fixedly connected to the optical fiber mounting seat 113. One end of the second connecting arm 117 is fixedly connected to the second fixing seat 112, and the other end is fixedly connected to the optical fiber mounting seat 113. One end of the third connecting arm 118 is fixedly connected to the first fixing seat 111, and the other end is fixedly connected to the lens adjustment seat 115. One end of the fourth connecting arm 119 is fixedly connected to the second fixing seat 112, and the other end is fixedly connected to the lens adjustment seat 115.

[0082] Preferably, if Figure 3 As shown, the lens adjustment seat 115 includes a first lens mounting seat 1151, a second lens mounting seat 1152, three elastic washers 1153, and three adjustment bolts 1154. The first lens mounting seat 1151 is fixedly connected to the third connecting arm 118 and the fourth connecting arm 119, respectively. The second lens mounting seat 1152 is mounted on a side surface of the first lens mounting seat 1151 near the optical fiber mounting seat 113. The three elastic washers 1153 are respectively mounted between the second lens mounting seat 1152 and the first lens mounting seat 1151. The three adjustment bolts 1154 are respectively used to connect the second lens mounting seat 1152, the corresponding elastic washers 1153, and the first lens mounting seat 1151. The three adjustment bolts 1154 and the three elastic washers 1153 are used to fix the reflective lens 130 and fine-tune its angle. The elastic washers 1153 serve to prevent loosening.

[0083] Preferably, the first fixing base 111 includes two fixing blocks 1111. The two fixing blocks 1111 are detachably connected by connecting bolts, facilitating assembly and disassembly for maintenance. One fixing block 1111 is fixedly connected to the first connecting arm 116, and the other fixing block 1111 is fixedly connected to the third connecting arm 118. The structure of the second fixing base 112 is identical to that of the first fixing base 111.

[0084] Preferably, the first fixing seat 111 and the second fixing seat 112 are respectively adhered to the spacecraft structure 200 by epoxy resin.

[0085] Preferably, the yielding hole is a stepped hole. The sleeve 114 is adhered in the large-diameter hole section of the stepped hole.

[0086] Preferably, the reflecting mirror 130 is installed on the second mirror mounting seat 1152 by adhesion.

[0087] Preferably, the surface of the reflecting mirror 130 is plated with a gold layer, which has a higher reflectivity and effectively improves the detection performance.

[0088] Preferably, the lengths of the first connecting arm 116, the second connecting arm 117, the third connecting arm 118 and the fourth connecting arm 119 are 145 mm, the arm thickness is 4 mm, the included angle between the extension direction and the horizontal direction is 2°, and the theoretical amplification ratio is cotα=28.6. The actual amplification ratio needs to be calibrated and tested.

[0089] On the other hand, the application also provides a use method of the spacecraft structure micro stress detection device, which comprises the following steps: fixing and installing the two ends of the displacement amplification structure of the spacecraft structure micro stress detection device on one side surface of the spacecraft structure 200, respectively; then, applying a first preset pressure to the spacecraft structure 200 by using the driver; and then, applying a second preset pressure to the spacecraft structure 200 by using the driver. It needs to be noted that the second preset pressure is greater than the first preset pressure.

[0090] Specifically, first, a sample to be measured is prepared to ensure that the surface is smooth and has no obvious defects. Then, the two ends of the displacement amplification structure of the spacecraft structure micro stress detection device are installed at predetermined positions of the sample to be measured by epoxy resin glue, respectively, to ensure that the sample to be measured is closely attached to the spacecraft structure micro stress detection device. After that, a micro stress test is performed. Finally, a large stress test is performed.

[0091] When the micro stress test is performed, first, the first pressure sensor is calibrated by using a weight, and the driving force of the piezoelectric ceramic driver is calibrated by using the calibrated first pressure sensor. Then, the first pressure sensor is placed between the driving end of the piezoelectric ceramic driver and the sample to be measured. The piezoelectric ceramic driver is used to apply a first preset pressure to the sample to be measured. The stress on the sample to be measured is changed from 0.1 kPa to 15 kPa by increasing the driving force, so as to verify the small stress monitoring range of the stress monitoring system. The first preset pressure is 0.1 kPa, 1.5 kPa, 3 kPa, 9 kPa or 15 kPa, and the driving force of the piezoelectric ceramic driver is changed from 0.05 N to 7.55 N with a step of 0.75 N.

[0092] When conducting a large stress test, first, use a weight to calibrate the second pressure sensor, and use the calibrated second pressure sensor to calibrate the driving force of the piezoelectric ceramic driver. Then, place the second pressure sensor between the driving end of the piezoelectric ceramic driver and the sample to be tested. Use the piezoelectric ceramic driver to apply a second preset pressure to the sample to be tested. By increasing the driving force, the stress on the sample to be tested changes from 30kPa to 1050kPa to verify the large stress monitoring range of the stress monitoring system. The second preset pressure is 30kPa, 100kPa, 400kPa, 600kPa, 800kPa or 1050kPa, and the driving force of the piezoelectric ceramic driver changes from 25N and increases to 525N in steps of 0.75N.

[0093] At each stress value, the stress value displayed by the monitoring system is recorded, and the stress value is obtained after fitting data processing and compared with the actual applied load.

[0094] In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "length", "width", "thickness", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", "clockwise", "counterclockwise", "axial", "radial", "circumferential" and the like to indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as limiting the present invention.

[0095] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of the technical features being referred to. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of the present invention, "plurality" means at least two, such as two, three, etc., unless otherwise specifically defined.

[0096] In the present invention, unless otherwise specified or limited, the terms "installed," "connected," "connect," "fixed," etc. should be understood in a broad sense. For example, they can refer to fixed connection, detachable connection, or integration; mechanical connection, electrical connection, or communication; direct connection or indirect connection through an intermediate medium; internal communication between two elements or interaction between two elements, unless otherwise specified. Those skilled in the art will understand the specific meanings of the above terms in the present invention based on specific circumstances.

[0097] In this disclosure, the terms "one embodiment", "some embodiments", "an example", "a specific example", or "some examples" mean that a particular feature, structure, material, or characteristic is included in at least one embodiment or example of the present disclosure. The illustrative appearances of the above-mentioned terms in various places in the specification are not necessarily referred to the same embodiment or example. Moreover, the particular features, structures, materials, or characteristics can be combined in any suitable manner in one or more embodiments or examples. Furthermore, the terminology "comprising" is used in the disclosure as comprising but not limited to, that is, it is open-ended and does not exclude the presence of additional features, structures, materials, or characteristics.

[0098] Although the embodiments of the present disclosure have been shown and described above, it is understood that the above-described embodiments are exemplary, and are not to be interpreted as limiting the present disclosure, and the ordinary skilled in the art can make changes, modifications, replacements, and variations to the above-described embodiments within the scope of the present disclosure.

Claims

1. A device for detecting micro stress of a spacecraft structure, characterized in that: include: The displacement amplification structure has two opposite ends fixedly mounted on one side of the spacecraft structure; A single-mode optical fiber is installed on one side of the displacement amplification structure; a reflective lens, mounted on the other side of the displacement amplification structure and arranged opposite to one end face of the single-mode optical fiber; a Fabry-Perot resonant cavity is formed between a side of the reflective lens close to the single-mode optical fiber and the end face of the single-mode optical fiber; a broadband light source connected to the single-mode optical fiber via a fiber coupler; a spectrum analyzer connected to the single-mode optical fiber via the optical fiber coupler; When the spacecraft structure is subjected to stress, the distance between the two opposite ends of the displacement amplification structure changes, thereby changing the cavity distance of the Fabry-Perot resonant cavity.

2. The spacecraft structure micro-stress detection device according to claim 1, characterized in that: The displacement amplification structure is rhombus-shaped as a whole, and includes: a first fixing seat, fixedly mounted on the spacecraft structure; a second fixing seat, arranged opposite to the first fixing seat and fixedly mounted on the spacecraft structure; An optical fiber mounting seat is provided between the first fixing seat and the second fixing seat; a clearance hole is formed on the optical fiber mounting seat; a sleeve, installed in the clearance hole, capable of allowing the end of the single-mode optical fiber to pass through; A lens adjustment seat is provided between the first fixing seat and the second fixing seat; the reflective lens is installed on a side surface of the lens adjustment seat close to the optical fiber mounting seat; a first connecting arm, one end of which is fixedly connected to the first fixing seat, and the other end of which is fixedly connected to the optical fiber mounting seat; a second connecting arm, one end of which is fixedly connected to the second fixing seat, and the other end of which is fixedly connected to the optical fiber mounting seat; a third connecting arm, one end of which is fixedly connected to the first fixing seat, and the other end of which is fixedly connected to the lens adjustment seat; A fourth connecting arm has one end fixedly connected to the second fixing seat, and the other end fixedly connected to the lens adjusting seat.

3. The spacecraft structure micro-stress detection device according to claim 2, characterized in that: The lens adjustment seat comprises: a first lens mounting seat, fixedly connected to the third connecting arm and the fourth connecting arm respectively; A second lens mounting seat is mounted on a side surface of the first lens mounting seat close to the optical fiber mounting seat; an elastic washer, mounted between the second lens mounting seat and the first lens mounting seat; An adjusting bolt is used to connect the second lens mounting seat, the elastic washer and the first lens mounting seat.

4. The spacecraft structure micro-stress detection device according to claim 2, characterized in that: The first fixing seat includes: There are two fixing blocks connected by connecting bolts; one of the fixing blocks is fixedly connected to the first connecting arm, and the other fixing block is fixedly connected to the third connecting arm; The structure of the second fixing base is the same as that of the first fixing base.

5. The spacecraft structure micro-stress detection device according to claim 2, characterized in that: The first fixing seat and the second fixing seat are respectively bonded to the spacecraft structure.

6. The spacecraft structure micro-stress detection device according to claim 2, characterized in that: The relief hole is a stepped hole.

7. The spacecraft structure micro-stress detection device according to claim 2, characterized in that: The reflective lens is mounted on the second lens mounting seat by bonding.

8. The spacecraft structure micro-stress detection device according to claim 2, characterized in that: The surface of the reflective lens is plated with a gold layer.

9. The spacecraft structure micro-stress detection device according to claim 2, characterized in that: The length of the first connecting arm, the second connecting arm, the third connecting arm and the fourth connecting arm is 145 mm, the arm thickness is 4 mm, and the angle between the extension direction and the horizontal direction is 2°.

10. A method for using the spacecraft structure micro-stress detection device according to any one of claims 1 to 9, characterized in that: include: The two ends of the displacement amplification structure of the spacecraft structure micro-stress detection device are respectively fixedly mounted on one side surface of the spacecraft structure; applying a first predetermined pressure to the spacecraft structure using an actuator; A second predetermined pressure is applied to the spacecraft structure using the actuator.

Images