Aircraft wing impact positioning method based on fiber grating sensor

Abstract

The invention relates to a wing low-speed impact positioning method based on a distributed fiber bragg grating sensing network, and belongs to the technical field of sensing application of optical fiber sensors. Simulating and selecting a proper positioning index by using an inverse finite element method; according to the area of the structure, fiber grating sensors are laid along a wing and connected with a wavelength demodulator, and an initial signal of an impact event is collected through a computer and then processed; impact signals at different heights are obtained through the impact experiment platform, and the arrival time of the impact signals collected by the sensor is obtained. The signal processing flow comprises the following steps of: firstly, decomposing an impact signal through a variational mode decomposition algorithm aiming at the noise interference and wave aliasing problems of the impact signal, then selecting a reconstructed signal component through time-frequency domain matching and analysis of an original signal and each mode component, calculating an envelope line by utilizing a Hilbert algorithm to obtain signal arrival time, and finally, calculating the signal arrival time by utilizing the Hilbert algorithm. And finally, reconstructing the distribution condition of a strain field and a deformation field when an impact event occurs through an inverse finite element method to realize impact positioning. The purpose of wing structure health monitoring is achieved through the fiber bragg grating sensing network.

Description

Technical Field

[0001] The present invention relates to a novel sensing method and device for impact positioning of an aircraft wing structure, which uses a fiber optic grating sensor network to detect impact signals when an impact event occurs in real time online, and realizes single-point and multi-point positioning in the field of low-speed impact of the wing structure, belonging to the field of optical fiber sensing. Background Art

[0002] Carbon fiber-reinforced polymer (CFRP) composites have the characteristics of high strength, light weight, corrosion resistance and excellent design flexibility. They have been widely used in aviation, high-speed railway systems, civil engineering, wind power generation, shipbuilding and marine engineering. This requires reasonable design and control of the real-time and long-term performance of CFRP components. Structural health monitoring (SHM) technology can provide real-time information on the state of the structure. The main goal of SHM is to detect the behavior of composite structures and determine the failed or unhealthy structural state. Since impact events will cause damage that is difficult to observe in the long run and cause serious losses, the monitoring of impact events has certain research value. Summary of the invention

[0003] The fiber Bragg grating aircraft wing impact positioning system proposed in the present invention is a new type of online detection device, which has the following unique advantages: (1) The fiber Bragg grating aircraft wing impact positioning system completely uses optical signals in the sensing and transmission areas, and is electrically insulated, and has the characteristics of intrinsic safety, explosion-proof, lightning protection, and anti-electromagnetic interference; (2) The fiber Bragg grating sensor, that is, the fiber Bragg grating (FBG), is small in size and light in weight, and is easy to install on the outside of the wing structure or embed into the inside for installation and use; (3) The use of a new denoising algorithm can effectively eliminate environmental noise and the non-stationary and nonlinear characteristics of the impact signal, and can accurately estimate the arrival time, further improving the positioning accuracy. The following steps are included:

[0004] S1: Use Ansys Workbench to build a three-dimensional model of the wing structure; extract strain data and use the inverse finite element method to simulate and select appropriate positioning indicators.

[0005] S2: Determine the total number of fiber grating sensors based on the wing structure length and size as well as the spatial spacing.

[0006] S3: These fiber grating sensors are arranged around and installed at various marked areas on the same horizontal plane of the wing at regular intervals. Among them, the first fiber grating sensor is used for ambient temperature compensation, and the remaining fiber grating sensors are installed closely on the wing surface (or embedded) to monitor the vibration signal when an impact event occurs.

[0007] S4: The impact position is determined through the impact test platform. The fiber grating sensors are laid along the wing and connected to the wavelength demodulators respectively. Finally, they are connected to the computer and the back-end processing is performed on it.

[0008] S5: In general, it is necessary to calculate the arrival time of the signal and extract the strain value at that moment for positioning imaging. However, since the impact signal is usually a non-stationary and nonlinear signal with multimodal noise, the variational mode decomposition method is used, and the signal arrival time is calculated by Hilbert transform.

[0009] S6: For the signal processing part, first obtain the impact signal for time domain and frequency domain analysis, and set the number of variational modal decomposition layers to 6 to reduce the calculation time. Perform Fourier transform on different modal components for frequency domain analysis and compare them.

[0010] S7: VMD establishes a variational problem and sets constraints to solve it:

[0011]

[0012] S8: Introduce the quadratic penalty factor α and the Lagrange multiplication operator λ(t) to transform the constrained variational problem into an unconstrained variational problem;

[0013]

[0014] S10: Compare the time and frequency domains of each modal component with the original signal to select the correct real signal Perform Hilbert transform H[x(t)];

[0015]

[0016] S11: Solve the complex analytical signal z(t);

[0017] S12: Its envelope is |z(t)|, and the signal arrival time is obtained.

[0018] S13: The maximum strain value at the time of impact is extracted using the signal arrival time, and provided to the inverse finite element method for strain field inversion for positioning.

[0019] S14: The maximum strain value at the time of impact is extracted using the signal arrival time, and provided to the inverse finite element method for strain field inversion for positioning.

[0020]

[0021] Among them, u e represents the displacement vector of each inverse shell element node, B m and B k represents the matrix containing the shape function, and ε xx ε yy γ xy It represents the strain in the x and y directions and the shear strain in the xy direction.

[0022] For low-speed impact positioning of aircraft wings, the present invention adopts wavelength division multiplexing technology, selects several groups of identical fiber gratings with different central wavelengths, and performs real-time monitoring at a sampling rate of 2000 Hz. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] In order to make the purpose, technical solution and beneficial effects of the present invention clearer, the present invention provides the following drawings for illustration:

[0024] Figure 1 This is a flow chart of the method for locating low-speed impact of a wing based on a fiber grating sensor according to the present invention;

[0025] Figure 2 This is a flow chart of the fiber Bragg grating optimization layout based on finite element simulation according to the present invention;

[0026] Figure 3 The present invention provides a flow chart of a signal processing method based on variational mode decomposition combined with Hilbert transform and an impact positioning method based on inverse finite element method. DETAILED DESCRIPTION

[0027] The embodiments of the present invention will be described in detail below with reference to the accompanying drawings.

[0028] The present invention discloses an aircraft wing impact positioning method based on a fiber grating sensor. First, a three-dimensional model of the corresponding wing structure is established using modeling software. After impact analysis of the three-dimensional model, the strain data is provided to the inverse finite element method to determine the number and location of the sensors, and the strain contour map is determined as a positioning indicator. A plurality of ribbon fiber grating sensors are attached to corresponding positions on the wing surface to monitor the vibration of the wing skin structure. The multimodal noise in the impact signal is removed by variational mode decomposition, and the accurate signal arrival time is estimated in combination with the Hilbert transform. The experimental strain value is provided to the inverse finite element method for imaging positioning. Compared with traditional electrical sensors, this method has better anti-electromagnetic interference and volume advantages.

[0029] like Figure 1As shown, the present invention mainly includes three parts, namely: sensor optimization layout, signal processing and impact positioning;

[0030] like Figure 2 As shown, the aircraft wing impact positioning method based on fiber grating sensor of the present invention specifically includes the following steps:

[0031] 1. Obtain the specific parameters of the structure or object to be monitored and establish a three-dimensional model

[0032] 1.1. Use ANSYS Workbench to perform finite element simulation on the model to obtain strain and deformation contour maps.

[0033] 1.2. Provide the strain data to the inverse finite element method for inversion and select appropriate imaging indicators.

[0034] 2. Sensor layout

[0035] 2.1. Determine the total number of fiber grating sensors based on the wing structure size and space spacing.

[0036] 2.2. Install these fiber Bragg grating sensors around the marked areas on the same horizontal plane of the wing at regular intervals. Use the optical fiber laid along the wing surface as the sensing element. Then use the fiber Bragg grating demodulator to demodulate the collected signal and save the demodulation result to the personal computer (PC).

[0037] like Figure 3 As shown, a wing low-speed impact positioning signal processing and impact positioning method based on fiber grating sensor includes the following steps:

[0038] 1. Impact signal preprocessing

[0039] 1.1、VMD establishes a variational problem and sets constraints to solve it. The quadratic penalty factor α and the Lagrange multiplication operator λ(t) are introduced to transform the constrained variational problem into an unconstrained variational problem.

[0040]

[0041] 1.2. Compare and analyze each modal component with the original signal in the time and frequency domain to select the correct real signal Perform Hilbert transform H[x(t)];

[0042]

[0043] 1.3 Solve the complex analytical signal z(t); Find its envelope |z(t)| and get the signal arrival time.

[0044] 2. Inverse finite element method to reconstruct imaging positioning

[0045] 2.1. Extract appropriate strain data according to the signal arrival time to invert the strain field when the impact event occurs.

[0046]

[0047] Among them, u e represents the displacement vector of each inverse shell element node, B m and B k represents the matrix containing the shape function, and ε xx ε yy γ xy It represents the strain in the x and y directions and the shear strain in the xy direction. In this way, the strain field can be inverted.

[0048] 2.2. Usually, when an impact event occurs, the strain field at the impact point is the largest and is distributed in a circle around the impact point, so the impact point can be found.

[0049] Finally, it should be noted that the above preferred embodiments are only used to illustrate the technical solutions of the present invention rather than to limit it. Although the present invention has been described in detail through the above preferred embodiments, those skilled in the art should understand that various changes can be made in form and details without departing from the scope defined by the claims of the present invention.

Claims

1. A method for locating an aircraft wing impact based on a fiber grating sensor, characterized in that: The method comprises the following steps: Establish a three-dimensional model of the wing structure and use the inverse finite element method (iFEM) to simulate and select appropriate positioning indicators; Impact test platform, several fiber Bragg grating (FBG) sensors, and fiber Bragg grating sensor arrays are arranged on the wings; A wavelength demodulator and a computer. The impact test platform is used to realize impact events at different heights. Fiber Bragg grating sensors are laid along the wing surface and connected to the wavelength demodulator to collect impact signals. Processing the collected signals; The experimental data are provided to the inverse finite element method for strain field inversion and imaging positioning.

2. The method for locating an aircraft wing impact based on a fiber grating sensor as claimed in claim 1, characterized in that: The 3D model of the wing structure was established using ANSYS Workbench software, and an impact test was carried out on the 3D model of the wing to obtain the simulated strain data, which was provided to the inverse finite element method for positioning index selection.

3. The method for locating an aircraft wing impact based on a fiber grating sensor as claimed in claim 1, characterized in that: Arrange sensors on the platform: determine the total number of fiber grating sensors and the distributed layout of the fiber grating sensors based on the size of the wing and the spatial intervals during installation, and use optical fibers of the same specifications to make several groups of identical fiber gratings; connect and arrange these fiber grating sensors at certain spatial intervals and install them in various marking areas on the same horizontal plane of the wing, among which the first fiber grating sensor is used for ambient temperature compensation, and the remaining fiber grating sensors are installed close to the wing surface (or embedded) to collect impact signals.

4. The method for locating an aircraft wing impact based on a fiber grating sensor as claimed in claim 1, characterized in that: The impact position is determined by the impact test platform, and the fiber grating sensors are laid along the wing and connected to the wavelength demodulators respectively, and finally connected to the computer for back-end processing.

5. The method for locating an aircraft wing impact based on a fiber grating sensor as claimed in claim 1, characterized in that: Signal processing: The impact signal is modally decomposed by variation mode decomposition (VMD), and the appropriate modal components are selected for Hilbert envelope analysis to obtain the earliest arrival time of the impact signal to extract the global sensor strain value.

6. The method for locating an aircraft wing impact based on a fiber grating sensor as claimed in claim 5, characterized in that: The signal preprocessing is as follows: 1) VMD establishes the solution of variational problem. The algorithm can be divided into the construction and solution of variational problem, and its expression is: where u k ={u1,u2,...,u k } are the modal functions, ω k ={ω1,ω2,...,ω k } is the center frequency of each mode; 2) Introducing the quadratic penalty factor α and the Lagrange multiplication operator λ(t), the constrained variational problem is transformed into an unconstrained variational problem, expressed as: 3) Each mode function and its center frequency are updated as follows: 4) Select appropriate modal components for Hilbert transform: in For a given real signal, H[x(t)] represents the Hilbert transform of the signal.

7. The method for locating an aircraft wing impact based on a fiber grating sensor as claimed in claim 1, characterized in that: Inversion using the inverse finite element method: Among them, u e represents the displacement vector of each inverse shell element node, B m and B k represents the matrix containing the shape function, and {ε xx ε yy γ xy } represents the strain in the x and y directions and the shear strain in the xy direction.

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