High-reliability fiber bragg grating deformation measurement method

Abstract

The invention belongs to the technical field of optical fiber distributed sensing, and discloses a high-reliability fiber bragg grating deformation measurement method, which comprises the following steps of: arranging fiber bragg grating pairs extending along the wingspan direction on the side vertical surface or the upper and lower surfaces of a front beam of a wing to obtain strain information of each position. On the basis, the wing bending deformation is solved based on the bending moment-deformation relation and structural information in material mechanics, and wing bending deformation monitoring is achieved. According to the method, the problem of poor real-time performance of wing deformation flight verification by combining the fiber grating with the KO algorithm is solved, and the reliability and robustness of deformation measurement can be remarkably improved while the deformation measurement precision is ensured, so that the application requirements of aircraft wing deformation monitoring are met.

Description

Technical Field

[0001] This invention belongs to the field of fiber optic distributed sensing technology, specifically relating to a highly reliable fiber optic grating deformation measurement method. Background Technology

[0002] Fiber optic deformation sensing technology, based on distributed strain information provided by a fiber optic distributed sensing system and combined with deformation algorithms to calculate structural deformation, is an important application technology in the field of fiber optic sensing in recent years. In airfoil deformation measurement applications, it can monitor aircraft wing surface deformation in real time, effectively ensuring flight stability and safety, improving flight efficiency, and has broad application prospects.

[0003] Because airfoils have good rigidity and do not involve complex deformation, the mature wavelength division multiplexing (WDM) fiber optic grating (FGF) sensing technology is more suitable than optical frequency domain reflection (OFDR) technology for airfoil strain sensing, effectively improving the dynamics and reliability of strain information acquisition. Commonly used methods for deformation calculation include the Koebner phenomenon (KO), modal analysis, and neural network methods. The KO algorithm is based on the integral calculation of strain-deflection in material mechanics, with a clear theory but limited accuracy due to integration errors. The modal analysis method is based on the principle of modal superposition, offering good dynamic accuracy but complex modal matrix calibration and high requirements for sensor deployment. The neural network method is based on model training and optimization, achieving high deformation accuracy but incurring high training costs and generally poor real-time performance. Fiber optic deformation schemes using the above methods have been researched and experimentally verified. In engineering implementation, although the KO algorithm introduces some integration errors, it has advantages in fiber optic deployment, algorithm complexity, and adaptability flexibility, making it more suitable for engineering applications.

[0004] Currently, there are reports on using fiber Bragg gratings combined with the Koebner phenomenon (KO) algorithm for wing deformation flight verification. However, the double integration computation of the KO algorithm still limits the real-time performance of deformation calculation and reduces system dynamics. On the other hand, in the field of wing deformation measurement, related research focuses on describing and designing the algorithm performance itself, without analyzing the impact of sensor failure, laser mode hopping, and other factors on the system and conducting reliability design. Summary of the Invention

[0005] This invention addresses the problem of poor real-time performance in existing methods for wing deformation flight verification using fiber Bragg gratings combined with the KO algorithm. Based on the ease of implementation and reliability of wing deformation measurement in engineering, this invention provides a highly reliable fiber Bragg grating deformation measurement method.

[0006] The technical solution of this invention is implemented as follows: A highly reliable method for measuring fiber optic grating deformation includes the following steps: Step 1: Horizontally lay two identical N-point fiber optic grating strings on both sides of the centerline of the front sparsity of the wing along the wingspan direction to form N pairs of strain measurement grating points. Step 2: The demodulation equipment has two channels connected to two fiber Bragg grating strings. It collects the center wavelength of each fiber Bragg grating and records the center wavelength of each grating point when the wing is at zero position as the zero position reference. At the same time, a wavelength range is preset for each grating point. A single fiber Bragg grating string has N preset wavelength ranges. Step 3: Update the wavelength range in real time according to the validity of each grating signal to ensure that there is at most one center wavelength in a single wavelength range. When the grating signal fails, the center wavelength value is 0. Step 4: Using the center wavelength of each grating point at the zero position of the wing as the reference wavelength, calculate the wavelength change Δλ of each subsequent grating point; subtract the wavelength changes of the upper and lower grating points at each position from each other to obtain the differential wavelength change δλ, introduce the calibration coefficient K to obtain the differential micro-strain δε at that position for deformation calculation. Step 5: For a pair of grating points effective for the wavelength, with a spacing h between the pair of grating points... i The differential microstrain δε is used as input to calculate the location X. i Angle θ at the point i ; Step 6: Discretize position X using the element length L as the step size. i Angle θ at the point i As input, interpolation is performed along the wingspan direction to calculate the rotation angles at various points on the wing; Step 7: Based on the rotation-deformation geometry model, calculate the airfoil deformation using a complex plane vector accumulation method; Step 8: Correct K using the actual deflection after one loading, and then repeat steps 4 to 7 with grating point wavelength demodulation to achieve dynamic measurement of wing deformation.

[0007] As a further aspect of the present invention: each pair of strain measurement grating points takes the wing root as the zero point and its position along the wingspan direction is X. i For i∈{1,2,…N}, the two grating points are equidistant from the centerline of the front wing sparsity, and are h apart. i , i∈{1,2,…N}.

[0008] As a further aspect of the present invention: In step one, for the case where the width of the front sparsity side facade of the wing contracts along the wingspan, the distance h between the upper and lower grating points along the wingspan direction... i The size can be gradually reduced depending on the available space, but it is necessary to ensure that the grating points are arranged parallel to the center line of the side facade.

[0009] As a further aspect of the present invention: in steps two and three, the wavelength ranges of the N grating points are adjacent to each other. The specific calculation method is to extend the center wavelength of the grating point to both sides by a certain amount of wavelength. Then, for the case of overlapping adjacent wavelengths, the average of the center wavelengths of the two grating points can be taken as the boundary of the two adjacent wavelength ranges.

[0010] As a further aspect of the present invention: in step three, if any wavelength in each pair of grating points is 0, then the grating signal of that pair of grating points is invalid and does not participate in the calculation in step four.

[0011] As a further aspect of the present invention: in steps two and three, after each acquisition of the grating point signal, the grating signal within the wavelength range of each grating point is calculated based on the wavelength range of the previous frame. For cases where the signal is damaged and has no center wavelength, the center wavelength is recorded as 0. For cases where the center wavelength increases, the signal is filtered according to the center wavelength spacing and peak intensity, retaining reliable grating point signals or directly recording them as 0. For reliable grating point center wavelengths, the wavelength range is updated.

[0012] As a further aspect of the present invention: in step four, a calibration coefficient K is introduced to correct the strain transfer coefficient, and the formula for calculating the differential micro-strain δε is as follows: δε=δλ*K*1.18; K is set to 1 by default and will be determined after calibration.

[0013] As a further aspect of the present invention: in step six, the rotation angle in the wingspan direction is interpolated by adding boundary constraints.

[0014] As a further aspect of the present invention: In step seven, a two-dimensional complex plane is established, with the span direction as the x-axis and the deflection direction as the y-axis, and the rotation angle θ on the length L of each element determined in step six is ​​used. i Calculate the orientation angle φ of each element. i ,in, ; Finally, the deflection y of the i-th element length in the wingspan direction i for: ; Each φ i The calculation can add an intermediate variable X. i By using recursion, the amount of computation can be reduced, and ; Where X0=θ0=0, then the final φ i for: ; Similarly, for y i Alternatively, you can use a similar method to iterate through and calculate all y values ​​at once.

[0015] As a further aspect of the present invention: in step eight, while ensuring process consistency, K at each grating point can be approximately the same, therefore, calibration can be completed with only one loading. K1 = K0 * y1 / y0; Where K0 is the current K coefficient, which defaults to 1, K1 is the calibrated coefficient, y0 is the deflection at the measurement point obtained by the algorithm after the wing is loaded, and y1 is the actual deflection at the measurement point.

[0016] Compared with the prior art, the beneficial effects of the present invention are as follows: (1) The fiber optic deployment and sensor calibration process is simple and has strong engineering feasibility. It avoids a large amount of simulation and calibration of the modal method, and does not require a large amount of data and training of neural networks. After a single grating sensor fails and is replaced, it does not need to be recalibrated, which reduces maintenance costs.

[0017] (2) Compared with the traditional KO algorithm, the computer airfoil deformation is based on the rotation (curvature)-deformation geometry model. The complex plane vector accumulation method is used to improve the computational efficiency and simplify the complexity of two integral operations.

[0018] (3) Improved system reliability. Matched wavelength regions and signal filtering mechanisms were designed for each grating sensor, avoiding mismatch between grating and wavelength signal caused by sensor failure, and ensuring the stability of system operation.

[0019] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. Attached Figure Description

[0020] Figure 1 A schematic flowchart of the wing deformation method according to an embodiment of the present invention is shown.

[0021] Figure 2 A schematic diagram of the structure of two grating sensors arranged on the front spar of an airfoil according to an embodiment of the present invention is shown.

[0022] Figure 3 A schematic diagram of sensor failure according to an embodiment of the present invention is shown.

[0023] Figure 4 A schematic diagram of the structural local rotation (curvature)-deformation calculation method according to an embodiment of the present invention is shown.

[0024] Figure 5 The cantilever beam deformation calculated according to an embodiment of the present invention and its beneficial effects are shown.

[0025] Figure 6 This diagram illustrates the deflection measurement data of the wing tip of a certain UAV during flight, according to an embodiment of the present invention. Detailed Implementation

[0026] To make the objectives, technical solutions, and advantages of this invention clearer, the technical solutions in the embodiments of this invention will be described in more detail below with reference to the accompanying drawings.

[0027] In the accompanying drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all, of the embodiments of the present invention.

[0028] The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain the present invention, and should not be construed as limiting the present invention. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without inventive effort are within the scope of protection of the present invention.

[0029] The following is in conjunction with the appendix Figure 1-6 The embodiments of the present invention will be described in detail below.

[0030] Example 1 This invention provides a highly reliable method for measuring fiber optic grating deformation. It includes the following steps: Step 1: Install fiber Bragg grating sensors. Two N-point fiber Bragg grating strings are horizontally arranged along the wingspan on both sides of the centerline of the front sparsity of the wing. The two grating strings are of identical specifications, forming N pairs of strain measurement points. The position of each pair of measurement points along the wingspan is X, with the wing root as the zero point. i For i∈{1,2,…N}, the upper and lower gratings are equidistant from the centerline of the front beam side facade, and are separated by a distance h. i , i∈{1,2,…N}. Step 2: The demodulation device connects two channels to two fiber Bragg grating strings, acquiring the center wavelength of each fiber Bragg grating and recording the center wavelength of each grating sensor at the wing's zero position as the zero-position reference. Simultaneously, a wavelength range is preset for each grating sensor, with N preset wavelength ranges per fiber.

[0031] Step 3: Update the wavelength range in real time based on the validity of each grating signal to ensure that there is at most one center wavelength within a single wavelength region. When a grating signal fails, the center wavelength value is 0. Based on this, the number of center wavelengths participating in subsequent deformation calculations for a single optical fiber is always N, corresponding to the actual number of sensors.

[0032] Step 4: Using the center wavelength of each grating sensor at the zero position of the wing as a reference, calculate the wavelength change Δλ of each subsequent grating sensor. Subtract the wavelength changes of the upper and lower gratings pairwise for each position to obtain the differential wavelength change δλ, and then obtain the differential micro-strain δε at that position for deformation calculation. A calibration coefficient K (default 1, determined after calibration) is introduced to correct the strain transfer coefficient. Wherein, X at each position...i If one of the center wavelengths in a grating pair is 0, the differential micro-strain δε of that sensor pair will not be used as input for subsequent deformation calculations.

[0033] Step 5: For wavelength-effective grating pairs, with grating pair spacing h... i The differential microstrain δε is used as input to calculate the location X. i Angle θ at the point i .

[0034] Step 6: Discretize position X using the element length L as the step size. i Angle θ at the point i As input, interpolate along the wingspan direction to calculate the rotation angle (curvature) at various points on the computer wing.

[0035] Step 7: Compute the airfoil deformation based on the rotation (curvature)-deformation geometry model, and improve the computational efficiency by using the in-plane vector accumulation method.

[0036] Step 8: Correct K using the actual deflection after a single loading. Then, using a grating sensor for wavelength demodulation, repeat steps 4 through 7 to achieve dynamic measurement of wing deformation.

[0037] Furthermore, in the design of the grating string parameters in step one, the N gratings can be evenly distributed, or the sensor spacing in the region of rapid strain change can be reduced to accommodate the static load of the wing. Sufficient wavelength spacing must be maintained between the center wavelengths of adjacent gratings, with the minimum wavelength spacing generally not less than 2 nm.

[0038] Furthermore, in step one, only the grating area and its extension portion of each grating are bonded to the structure using epoxy resin. For the case where the width of the front beam side facade contracts along the wingspan, the distance h between the upper and lower gratings along the wingspan direction... i The size can be gradually reduced depending on the available space, but it is necessary to ensure that the grating is installed parallel to the center line of the side facade.

[0039] Furthermore, in steps two and three, the wavelength ranges of the N gratings are adjacent to each other. The specific calculation method is to extend the center wavelength of the grating to both sides by several nm. Then, for the case of overlapping adjacent wavelengths, the average of the center wavelengths of the two gratings can be taken as the boundary of the two adjacent wavelength ranges.

[0040] Furthermore, in step three, after each sensor signal acquisition, the grating signal within each wavelength range is calculated based on the wavelength range of the previous frame. For signals that are damaged and lack a center wavelength, the center wavelength is recorded as 0. For signals with increased center wavelengths, the signals are filtered based on the center wavelength spacing and peak intensity, retaining reliable sensor signals or directly recording them as 0. For reliable grating center wavelengths, the wavelength range is updated.

[0041] Furthermore, in step five, a small-angle approximation is used, with the rotation angle θ... i In radians,

[0042] Furthermore, in step six, the rotation angle (curvature) in the wingspan direction is interpolated by adding boundary constraints. Cubic spline interpolation is more accurate than linear interpolation, but it increases the computation time.

[0043] Furthermore, in step seven, a two-dimensional complex plane is established, with the span direction as the x-axis and the deflection direction as the y-axis, and the rotation angle θ along the length L of each element determined in step six is ​​used. i Calculate the orientation angle φ of each element (two-dimensional vector). i ,in,

[0044] Finally, the deflection y of the i-th element length in the wingspan direction i for

[0045] Furthermore, in step seven, each φ i The calculation can add an intermediate variable X. i By using recursion, the amount of computation can be reduced, and

[0046] Where X0=θ0=0, then the final φ i for

[0047] Similarly, for yi, all y values ​​can be calculated in one iteration using a similar method.

[0048] Furthermore, in step eight, the calibration coefficient of K is mainly determined by the strain transfer characteristics caused by the bonding process between the grating and the structure. Under the condition of ensuring process consistency, K at each grating can be approximately the same. Therefore, calibration can be completed with only one loading operation. K1 = K0 * y1 / y0, where K0 is the current K coefficient (default is 1), K1 is the calibrated coefficient, y0 is the deflection at the measurement point obtained by the algorithm after wing loading, and y1 is the actual deflection at the measurement point.

[0049] Example 2 Reference Figure 1 , Figure 1 The flowchart of the wing deformation method according to an embodiment of the present invention is shown.

[0050] Step (1): Select two fiber Bragg grating strings 1 and 2 with identical parameters. Adhere the grating area to the side elevation 3 of the wing's front sparsity using epoxy resin DG-4. Each grating string contains 10 fiber Bragg gratings, distributed sequentially from the root to the wingtip. Spatially adjacent gratings have center wavelengths 3 nm apart in the C-band. The upper and lower gratings form 10 grating pairs. The grating area direction is parallel to the centerline of the wing's front sparsity side elevation, and the grating pairs are spaced hi apart, symmetrically distributed on both sides of the centerline. Figure 2 As shown.

[0051] In step (2), the two grating strings are connected to the grating wavelength demodulation device respectively. When the wing is at zero position, the initial 10 pairs of grating wavelengths are recorded as the reference wavelength. At the same time, 10 wavelength ranges are established for the 10 designed grating wavelengths. Each wavelength range is centered on the center wavelength of the corresponding sensor and extended 2nm on both sides.

[0052] Step (3): During each signal acquisition process, record the center wavelength and peak intensity of each reflection peak.

[0053] Then, the 10 wavelength ranges are matched sequentially. If there is only one center wavelength in the current wavelength range, the sensor is normal. If there is no center wavelength or two or more wavelengths appear, the sensor wavelength is set to 0 and the sensor reports a fault. Figure 3 This diagram illustrates sensor failure. A normal sensor's reflected signal spectrum is a Gaussian spectrum. However, due to factors such as bonding process defects, the Gaussian spectrum may split into two reflection peaks due to chirp effects at low temperatures or under load. For scanning lasers, mode hopping can even cause glitches in the sensor signal, all of which can lead to sensor signal failure. For failed sensors, the wavelength range is not updated; the previous valid range is retained. For normal sensors, the wavelength range is updated for use in determining the sensor's status after the next signal acquisition.

[0054] Step (4): For the grating pair in normal condition (both wavelengths are not 0), calculate the wavelength change of the upper and lower gratings relative to the reference wavelength, then subtract them to obtain the differential wavelength change δλ, and further multiply by the strain coefficient K to obtain the differential micro-strain δε.

[0055] Step (5): For grating pairs in normal condition, with grating pair spacing h... i The differential microstrain δε is used as input to calculate the location X. i Angle θ at the point i .like Figure 4 As shown, the height difference between the upper and lower gratings is h, and L is the element length for deformation calculation in the structure. When the structure bends, the upper and lower gratings stretch and contract respectively. When the differential microstrain of the upper and lower gratings is δε, since the element length L is small, this deformation can be approximated by a small angle and represented by an arc surface. According to geometric relationships, the bending angle...

[0056] Step (6): Based on the above calculation method, using discrete position X... i Angle θ at the point i Using the input as input, cubic spline interpolation is performed along the wingspan to obtain the rotation angle (curvature) at various points on the wing.

[0057] Step (7), according to Figure 4 As shown, a two-dimensional complex plane is established, with the span direction as the x-axis and the deflection direction as the y-axis. The rotation angle θ along the length L of each element, determined in step 6, is used. i Calculate the direction angle φ of each bending element (two-dimensional vector). i .according to Figure 4 In terms of geometric relationships, a curved element can be represented in a two-dimensional plane by the lines connecting its endpoints. In this case, the bending angle θ will cause subsequent elements to deflect by the same angle, while its own orientation angle will only change by θ / 2. Therefore, the orientation angles of each element are:

[0058] Finally, the deflection y of the i-th element length in the wingspan direction i for .

[0059] Step (8) applies a load to the wing. At this point, the actual deflection at the wingtip is y1. Under the condition of algorithm K=1, the calculated deflection is y0. After updating K=1 to K1 using the formula K1=K0*y1 / y0, the calibration is completed. The demodulation equipment dynamically measures the deflection at each point of the wing according to steps 4 to 7.

[0060] Figure 5 To test the effectiveness of the method, single-point, two-point, and multi-point loads were applied to the structure, and the deflection at the end of the cantilever beam was calculated. The maximum deviation between the calculated results and the actual measurements was controlled within 1 mm (the maximum deflection at the end was 106 mm). The above deformation algorithm was processed using a DSP28335, and the actual processing time was better than 2 ms.

[0061] In some embodiments, when step 1 involves a gradual reduction in the height of the front beam side facade, the grating pair can gradually reduce the vertical spacing h according to the allowable spatial dimensions. i However, it is necessary to ensure that the direction of the grating area is parallel to the center line.

[0062] In some embodiments, based on updating the wavelength range in step 3 according to the effective sensor center wavelength within a 2nm range, the wavelength range can be further updated according to the average value of the center wavelengths of adjacent effective sensors to ensure the continuity between wavelength ranges.

[0063] In some embodiments, if two reflection peaks appear within a single wavelength range in step 3, a judgment condition can be added. If the center wavelengths of the two peaks are close and one is much stronger than the other, it can be considered that the sensor has a low degree of chirp (damage) and is still usable, and the center wavelength with higher intensity is retained.

[0064] In some embodiments, for applications where accuracy requirements are less stringent, such as wing flutter monitoring, in step 6, linear fitting can be used to fit the curvature in order to improve computational efficiency and system dynamics.

[0065] In some embodiments, the direction angle φ in step 7 i The calculation can add an intermediate variable X. i By using recursion, the amount of computation can be reduced, and

[0066] Where X0=θ0=0, then the final φ i for

[0067] Similarly, for y i Alternatively, you can use a similar method to iterate through and calculate all y values ​​at once.

[0068] In some embodiments, during the calibration process in step 8, the measurement value of deflection y in a single measurement may not be accurate due to sensor noise. In the calibration process, the y values ​​of multiple frames can be averaged continuously to calculate the calibration coefficient K.

[0069] Example 3 This invention provides a method for measuring the wing deformation of a certain unmanned aerial vehicle (UAV), specifically including: 1. System Composition: Fiber Bragg grating sensor: Select two custom fiber Bragg grating strings with the same parameters, each 3m long, each containing 10 gratings, arranged at equal intervals, with an wavelength interval of 3 nm between adjacent gratings (1530 nm, 1533 nm...1557 nm).

[0070] Demodulation equipment: DSP-based demodulation equipment (wavelength accuracy ±2 pm, sampling rate 100Hz), which connects two grating strings through two independent channels.

[0071] Processing unit: An embedded system based on the DSP28335 chip, used for real-time wavelength demodulation, deformation calculation and calibration processing, etc.

[0072] Loading and measuring equipment: servo loading device, laser rangefinder.

[0073] 2. Fiber Bragg grating sensor deployment and installation Deployment method (e.g.) Figure 2 (as shown) Two fiber grating strings are symmetrically attached to both sides of the centerline of the front spars of the wing, and are evenly distributed along the wingspan (from the wing root to the wingtip).

[0074] The upper and lower gratings are 20mm apart and symmetrical with respect to the center line.

[0075] The grating area is parallel to the centerline and is bonded with epoxy resin (model: DG-4). Only the grating area and the extension are bonded to avoid stress interference.

[0076] 3. Data Acquisition and Signal Processing Zero-point reference establishment (step two): With the wing unloaded, the initial center wavelength λ0 of each grating was recorded, and the wavelength range of each grating was set to λ. L —λ H .

[0077] Signal validity assessment (step three): In each sampling signal processing stage, within the wavelength range λ of each grating sensor... L —λ H Search for reflection peaks within the body: If there is no signal in the wavelength range or multiple reflection peaks appear (such as...) Figure 3 (As shown in the diagram, chirp failure or laser mode hopping) is recorded with a center wavelength of 0, marking the sensor as faulty, while retaining the wavelength range λ. L —λ H No updates will be performed.

[0078] The grating reflection peak is normal. The center wavelength λ is recorded for subsequent deformation calculations, and the wavelength range λ is updated based on λ. L —λ H Used to determine the validity of the next move.

[0079] Differential strain calculation (step four): For a pair of gratings where neither wavelength is zero, calculate the wavelength change of the upper and lower gratings relative to their initial center wavelength λ0, then subtract them to obtain the differential wavelength change δλ, and further multiply by the strain coefficient K (to be calibrated) to obtain the differential micro-strain δε.

[0080] 4. Deformation Calculation Angle calculation (step five): Calculate the rotation angle θ at position X of the grating pair based on the differential micro-strain δε and the spacing h:

[0081] Where L = 50mm, is the minimum element length used in deformation calculation.

[0082] Interpolation and curvature calculation (step six): A continuous curvature distribution is obtained using cubic spline interpolation based on the discrete rotation angle θ and the length position X of the element. For faster response, linear interpolation can be used, sacrificing some accuracy to improve processing speed.

[0083] Complex plane vector accumulation method (step seven): Establish a two-dimensional complex plane with the wingspan direction as the x-axis and the deflection direction as the y-axis.

[0084] Calculate the direction angle φ of each bending element (two-dimensional vector) using the rotation angle θ along the length L of each element determined in step 6. Figure 4 Geometric relationships, the orientation angles of each element are:

[0085] Finally, the deflection yi of the i-th element length in the wingspan direction is .

[0086] 5. System Calibration and Dynamic Measurement Calibration process (step eight): When the wing is unloaded, the initial wavelength λ0 of each sensor is recorded. At this time, the wingtip deflection y=0 is calculated. This state is the initial zero position of the wing. At the same time, the laser rangefinder is zeroed.

[0087] The wing is loaded by a servo loading device, the laser rangefinder displacement is y1, and the wingtip deflection is y0 under the condition of K=1. After updating K=1 to K1 by the formula K1=K0*y1 / y0, the calibration is completed.

[0088] Dynamic measurement: Real-time acquisition of wavelength data for each grating; repeat steps four through seven to output the airfoil deformation curve.

[0089] 6. Experimental Verification and Result Analysis Experimental verification: Figure 6 For the deflection measurement data of the wingtip of a UAV equipped with the device of this invention during flight, at approximately frame 4000, due to sensor chirp failure, the deflection measurement experienced a jump, decreasing from about 700mm to about 650mm. The system automatically marked the failed sensor, and this point was ignored in the deformation calculation. The overall deflection curve remained stable and did not recover thereafter. This demonstrates that the present invention can still maintain normal operation after partial sensor failure, improving reliability in engineering applications.

[0090] Advantages of this method: Ease of implementation: Dual-string gratings are easy to deploy and require no complex calibration or simulation.

[0091] High reliability: The dynamic wavelength range update mechanism effectively addresses anomalies such as sensor failure and laser mode hopping.

[0092] Computational efficiency: Complex plane vector algorithm replaces integral operation, enabling fast processing by DSP28335 chip.

[0093] Simplified calibration: Calibration K is applied in one step to achieve deformation measurement.

[0094] Thus, the objective of this invention has been achieved.

[0095] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A high-reliability method for measuring fiber optic grating deformation, characterized in that, Includes the following steps: Step 1: Horizontally lay two identical N-point fiber optic grating strings on both sides of the centerline of the front sparsity of the wing along the wingspan direction to form N pairs of strain measurement grating points. Step 2: The demodulation equipment has two channels connected to two fiber Bragg grating strings. It collects the center wavelength of each fiber Bragg grating and records the center wavelength of each grating point when the wing is at zero position as the zero position reference. At the same time, a wavelength range is preset for each grating point. A single fiber Bragg grating string has N preset wavelength ranges. Step 3: Update the wavelength range in real time according to the validity of each grating signal to ensure that there is at most one center wavelength in a single wavelength range. When the grating signal fails, the center wavelength value is 0. Step 4: Using the center wavelength of each grating point at the zero position of the wing as the reference wavelength, calculate the wavelength change Δλ of each subsequent grating point; subtract the wavelength changes of the upper and lower grating points at each position from each other to obtain the differential wavelength change δλ, introduce the calibration coefficient K to obtain the differential micro-strain δε at that position for deformation calculation. Step 5: For a pair of grating points effective for the wavelength, with a spacing h between the pair of grating points... i The differential microstrain δε is used as input to calculate the location X. i Angle θ at the point i ; Step 6: Discretize position X using the element length L as the step size. i Angle θ at the point i As input, interpolation is performed along the wingspan direction to calculate the rotation angles at various points on the wing; Step 7: Based on the rotation-deformation geometry model, calculate the airfoil deformation using a complex plane vector accumulation method; Step 8: Correct K using the actual deflection after one loading, and then repeat steps 4 to 7 with grating point wavelength demodulation to achieve dynamic measurement of wing deformation.

2. The high-reliability fiber optic grating deformation measurement method according to claim 1, characterized in that, Each pair of strain measurement grating points has its zero point at the wing root and its position along the span direction as X. i For i∈{1,2,…N}, the two grating points are equidistant from the centerline of the front wing sparsity, and are h apart. i , i∈{1,2,…N}.

3. The high-reliability fiber optic grating deformation measurement method according to claim 2, characterized in that, In step one, for the case where the width of the front sparsity side facade of the wing contracts along the wingspan, the distance h between the upper and lower grating points along the wingspan direction is... i The size can be gradually reduced depending on the available space, but it is necessary to ensure that the grating points are arranged parallel to the center line of the side facade.

4. The high-reliability fiber optic grating deformation measurement method according to claim 1, characterized in that, In steps two and three, the wavelength ranges of the N grating points are adjacent to each other. The specific calculation method is to extend the center wavelength of the grating point to both sides by a certain amount of wavelength. Then, for the case of overlapping adjacent wavelengths, the average of the center wavelengths of the two grating points can be taken as the boundary of the two adjacent wavelength ranges.

5. The high-reliability fiber optic grating deformation measurement method according to claim 1, characterized in that, In step three, if any wavelength in each pair of grating points is 0, then the grating signal of that pair of grating points is invalid and will not participate in the calculation in step four.

6. The high-reliability fiber optic grating deformation measurement method according to claim 1, characterized in that, In steps two and three, after each acquisition of the grating point signal, the grating signal within the wavelength range of each grating point is calculated based on the wavelength range of the previous frame. For cases where the signal is damaged and has no center wavelength, the center wavelength is recorded as 0. For cases where the center wavelength increases, the signal is filtered according to the center wavelength spacing and peak intensity, retaining reliable grating point signals or directly recording them as 0. For reliable grating point center wavelengths, the wavelength range is updated.

7. The high-reliability fiber optic grating deformation measurement method according to claim 1, characterized in that, In step four, a calibration coefficient K is introduced to correct the strain transfer coefficient, and the formula for calculating the differential micro-strain δε is as follows: δε=δλ*K*1.18; K is set to 1 by default and will be determined after calibration.

8. The high-reliability fiber optic grating deformation measurement method according to claim 1, characterized in that, In step six, the rotation angle in the wingspan direction is interpolated by adding boundary constraints.

9. The high-reliability fiber optic grating deformation measurement method according to claim 1, characterized in that, In step seven, a two-dimensional complex plane is established, with the span direction as the x-axis and the deflection direction as the y-axis, and the rotation angle θ along the length L of each element determined in step six is ​​used. i Calculate the orientation angle φ of each element. i ,in, ； Finally, the deflection y of the i-th element length in the wingspan direction i for: ； Each φ i The calculation can add an intermediate variable X. i By using recursion, the amount of computation can be reduced, and ； Where X0=θ0=0, then the final φ i for: ； Similarly, for y i Alternatively, you can use a similar method to iterate through and calculate all y values ​​at once.

10. The high-reliability fiber optic grating deformation measurement method according to claim 1, characterized in that, In step eight, while ensuring process consistency, K at each grating point can be approximately the same. Therefore, calibration can be completed with only one loading. K1 = K0 * y1 / y0; Where K0 is the current K coefficient, which defaults to 1, K1 is the calibrated coefficient, y0 is the deflection at the measurement point obtained by the algorithm after the wing is loaded, and y1 is the actual deflection at the measurement point.

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