A method for measuring the shape of a trailing edge of a variable camber wing using optical fibers

Abstract

The present application relates to a kind of variable camber wing trailing edge shape change of optical fiber measurement method, belong to the field of large wing trailing edge structure deformation monitoring such as aircraft. Including: the inner surface of the measured variable camber trailing edge structure is connected with the optical fiber sensing beam support structure through aviation special adhesive, optical fiber is installed in the inside of the upper and lower surface of sensing beam, sensing beam passes through the middle hole of optical fiber sensing beam support structure, and is fixed with optical fiber sensing beam support structure with screw mode, with optical fiber sensing beam support structure keeps free sliding. The tail end of optical fiber is connected with one end of single-core optical cable through optical fiber joint flange, the other end of single-core optical cable is connected with optical fiber grating center wavelength collector through FC / APC joint, and optical fiber grating center wavelength collector is connected with shape data processing module through network cable or wireless transmission module. The shape change history of the measured variable camber trailing edge structure can be detected in ground loading or flight process. Simple structure, strong anti-interference ability, long service life, low cost.

Description

TECHNICAL FIELD

[0001] The present application relates to a kind of variable camber wing trailing edge shape change optical fiber measurement method, belong to the large wing trailing edge structure deformation monitoring field of aircraft etc.. BACKGROUND

[0002] Variable camber wing trailing edge can change airfoil camber according to flight state, obtain optimal aerodynamic shape.Relative to rigid deformation wing trailing edge, its deformation curve is smoother, and weight is lighter, can significantly improve the lift-drag ratio of aircraft, reduce fuel consumption, enhance the stability and maneuverability of aircraft.If variable camber wing trailing edge shape change can be tested in real time during aircraft flight, the deformation mechanism of given target shape can be closed-loop controlled.The traditional laser interference method or binocular camera test deformation method, solving rate is slow, and it is difficult to install and debug in the aircraft, cannot be applied to dynamic shape test of aircraft, more not suitable for wing trailing edge shape online test under flight state.In addition, the method for measuring strain by using resistance strain gauge to deduce shape is also unable to realize accurate test of wing trailing edge shape because resistance strain gauge is easily affected by electromagnetic interference.The advanced electromagnetic interference resistant, and easily installed in the interior of variable camber wing trailing edge high-precision shape test means is the problem to be solved in the present application.Fiber grating takes optical signal as carrier, is not affected by electromagnetic interference, and structure is smart, sensitivity is high, can be packaged into shape test sensing beam.In order to solve the temperature strain cross-coupling problem existing in fiber grating sensor itself, the relative value of two symmetric gratings is designed to eliminate the influence of temperature.In order to solve the accuracy problem of optical fiber type deformation test, strain sensitivity coefficient calibration of sensing beam and key parameter control in installation process are proposed. SUMMARY

[0003] The present application aims at the demand that existing continuous variable camber trailing edge structure shape cannot be tested online, proposes a kind of variable camber wing trailing edge shape change optical fiber measurement method, solves the foregoing problems, and has important application value in variable camber wing trailing edge structure design and shape feedback field.

[0004] The present application aims at the demand that existing continuous variable camber trailing edge structure shape cannot be tested online, proposes a kind of variable camber wing trailing edge shape change optical fiber measurement method, solves the foregoing problems, and has important application value in variable camber wing trailing edge structure design and shape feedback field.

[0005] The idea of the present application is to drive the deformation of the optical fiber deformation beam by the deformation of the variable camber wing trailing edge structure. Two fiber gratings are designed and mounted on the upper and lower surfaces of the deformation beam by glue to form an optical fiber deformation beam. The optical fiber deformation beam is passed through the inner holes of several support structures. Only the root support structure is completely fixed with the optical fiber deformation beam by screws. The other support structures are as smooth as possible or wireless close to the optical fiber sensing beam without affecting the free sliding of the optical fiber sensing beam. The support structures are mounted on the skin inside the variable camber trailing edge structure by aviation special adhesive. Two fiber gratings are connected to the fiber grating center wavelength collector through a single-core optical cable. The center wavelength data acquisition during the shape change of the wing trailing edge is realized. The data is connected to the shape data processing module through a network cable or a wireless transmission module. The center wavelength data is processed to complete the shape test. The present application can realize the detection of the shape change process of the measured variable camber trailing edge structure during ground loading or flight.

[0006] The purpose of the present application is realized by the following technical solutions:

[0007] The present application relates to a kind of variable camber wing trailing edge deformation optical fiber measurement method, comprising: optical fiber sensing beam support structure, optical fiber, fiber grating, sensing beam, optical fiber joint flange, single-core optical cable, FC / APC joint, fiber grating center wavelength collector, network cable or wireless transmission module, shape data processing module.

[0008] The present application provides a kind of variable camber wing trailing edge deformation optical fiber measurement method, optical fiber sensing beam is installed in the upper skin inside of the measured variable camber wing trailing edge structure by four optical fiber sensing beam support structures, the inner surface of the measured variable camber wing trailing edge structure is connected with optical fiber sensing beam support structure by aviation special adhesive, optical fiber sensing beam passes through the middle hole of optical fiber sensing beam support structure, and optical fiber sensing beam is completely fixed with the first optical fiber sensing beam support structure using bolt, is not fixed with other optical fiber sensing beam support structures, and the inner wall contact surface of the middle hole of other optical fiber sensing beam support structures is as smooth as possible, can freely slide, the installation gap of optical fiber sensing beam support structure and optical fiber sensing beam is accurately controlled by optical distance measuring system, the shape test of the measured variable camber wing trailing edge structure under loading state is realized;A pair of optical fibers are installed in the upper and lower surfaces of optical fiber sensing beam, four fiber gratings are engraved on one optical fiber, and four fiber gratings are engraved on another optical fiber, the fiber gratings on the upper and lower surfaces are symmetrically arranged, the tail end of the pair of optical fibers is connected with one end of single-core optical cable through optical fiber joint flange, the other end of single-core optical cable is connected with fiber grating center wavelength collector through FC / APC joint, fiber grating center wavelength collector is connected with shape data processing module through network cable or wireless transmission module;The center wavelength change is calculated by difference method, and the deformation of trailing edge structure is calculated by sensing beam deformation data analysis only by center wavelength-strain-shape test Frenet algorithm.

[0009] Advantageously, the optical fiber sensing beam support structure is made of plastic or aluminum alloy, and is installed on the inner surface of the upper skin of the measured variable-camber wing trailing edge structure by using an aviation special epoxy adhesive.

[0010] Advantageously, the high-precision installation and pre-stretching process control of the pair of optical fibers and the optical fiber sensing beam are performed, the optical fiber sensing beam is made of plastic, aluminum alloy or stainless steel, and semicircular grooves with a diameter of 0.3 mm are engraved on the upper and lower surfaces of the optical fiber sensing beam, the surface of the semicircular groove is very smooth, the optical fiber is installed in the semicircular groove, and a pre-stress is applied to the optical fiber by using a weight at both ends to keep the optical fiber horizontally stretched and prevent the optical fiber from being bent, and the position of the fiber Bragg grating on the optical fiber is marked, and then the semicircular groove is filled with liquid epoxy glue and cured at room temperature.

[0011] Advantageously, the gap between the optical fiber sensing beam and the middle hole of the optical fiber sensing beam support structure can be automatically measured, and an optical distance measuring module is installed on the optical fiber sensing beam support structure to ensure that the gap can be adjusted to an ideal state, avoid friction, and freely bend and deform along with the deformation of the trailing edge structure.

[0012] Advantageously, the center wavelength-strain calibration of the optical fiber sensing beam is performed by using a calibration beam testing device before installation to obtain the strain sensitivity coefficient of the fiber Bragg grating and ensure the strain test accuracy.

[0013] The working principle is as follows: the upper surface skin of the variable-camber wing trailing edge is deformed by bending, which drives the optical fiber deformation beam to bend and deform through the support structure, the fiber Bragg grating on the optical fiber deformation beam is stretched or compressed due to bending, which causes the center wavelength to increase or decrease. Through the change of the center wavelength of the fiber Bragg grating on the optical fiber deformation beam, the curvature change of the optical fiber sensing beam is deduced, and through curvature interpolation fitting and frenet shape algorithm, the shape change of the trailing edge skin can be deduced.

[0014] As a preferred embodiment, the position accuracy of the fiber Bragg grating, the installation and positioning of the fiber Bragg grating and the sensing beam, and the gap control during the installation of the sensing beam and the support structure need to be optimized to achieve high precision and easy installation.

[0015] Advantages

[0016] Compared with the prior art, the present application has the following characteristics:

[0017] 1. The optical fiber sensing beam support structure is installed inside the wing trailing edge, which does not affect the appearance of the wing, has a simple structure, and is easy to operate, and can be installed together during wing assembly;

[0018] 2. The high-precision installation and pre-stretching process control of the optical fiber and the sensing beam enable the fiber Bragg grating engraved on the optical fiber to be precisely positioned and to withstand a certain compression deformation;

[0019] 3、The gap between the sensing beam and the hole of the support structure can be automatically measured, the optical distance measuring module is installed on the support structure, the gap can be adjusted to the ideal state, the friction is avoided, and the sensing beam can be freely deformed along with the deformation of the trailing edge structure;

[0020] 4、The fiber Bragg gratings on the upper and lower surfaces of the sensing beam are symmetrically arranged, the difference method is used to calculate the center wavelength change, the influence of the strain temperature cross coupling of the fiber Bragg grating can be ignored, and the test precision is high;

[0021] 5、The center wavelength-strain-shape test algorithm Frenet algorithm is used, the measured trailing edge structure does not need to be pre-stressed and deformed, and the deformation of the trailing edge structure can be calculated only by analyzing the deformation data of the sensing beam, the calculation is simple, and the engineering applicability and generalization are strong. BRIEF DESCRIPTION OF DRAWINGS

[0022] Figure 1 It is a schematic diagram of the measuring system of the application;

[0023] Figure 2 It is a schematic diagram of the fiber wiring mode;

[0024] Figure 3 It is a schematic diagram of the deformation of the fiber;

[0025] Figure 4 It is a schematic diagram of the installation section of the fiber and the sensing beam;

[0026] Figure 5 The flow chart of the measuring algorithm of the application. DETAILED DESCRIPTION

[0027] The application will be described in detail below in combination with the drawings and examples.

[0028] As Figures 1-2As shown, two optical fibers 4, 5 are designed, four fiber gratings 16, 17, 18, 19 are engraved on one optical fiber 4, four fiber gratings 20, 21, 22, 23 are engraved on the other optical fiber 5, the fiber gratings on the upper and lower surfaces are symmetrically arranged, and the optical fibers 4, 5 are mounted inside the upper and lower surfaces of the deformation beam to form a fiber sensing beam 3. The fiber sensing beam 3 passes through the internal holes of several fiber sensing beam support structures 2, 6, 7, 8. Only the root fiber sensing beam support structure 2 is completely fixed with the fiber sensing beam 3 by screws, and the other fiber sensing beam support structures 6, 7, 8 are as smooth as possible or infinitely close, without affecting the free sliding of the fiber sensing beam 3. The fiber sensing beam support structures 2, 6, 7, 8 are mounted inside the upper skin of the measured variable-camber wing trailing edge structure 1 by using aviation special adhesive. The two optical fibers 4, 5 are connected to the fiber grating central wavelength collector 13 through the optical fiber joint flange 9, single-core optical cable 10, 11 and FC / APC joint 12, respectively, to realize the central wavelength data collection during the shape change of the wing trailing edge. The data is connected to the shape data processing module 15 through the network cable or wireless transmission module 14, and the central wavelength data is processed to complete the shape test.

[0029] The working principle of the variable-camber wing trailing edge deformation optical fiber measurement method is that the upper surface skin of the variable-camber wing trailing edge is deformed by bending, which drives the fiber sensing beam 3 to bend and deform by stretching or compression through the fiber sensing beam support structures 2, 6, 7, 8. The fiber 4, 5 on the fiber sensing beam 3 is stretched or compressed, resulting in an increase or decrease in the central wavelength. Through the change of the fiber grating central wavelength on the fiber sensing beam 3, the curvature change of the fiber sensing beam 3 is deduced, and the fiber grating curvature measurement principle is as shown in Figure 3

[0030] When the fiber is bent, the upper grating is subjected to stress ε1, and the lower grating is subjected to stress ε2, so ε1=d1 / R; ε2=-d2 / R, the negative sign represents the compressive stress, then:

[0031] (1)

[0032] The core distance d is determined, and the curvature 1 / R of a specific point on the multi-core optical fiber can be calculated by demodulating the strain value through the wavelength shift.

[0033] Then the shape change of the trailing edge skin can be deduced through curvature interpolation fitting and frenet shape algorithm. The flow chart of the curvature reconstruction algorithm based on fiber sensing is as shown in Figure 4

[0034] The material of the fiber sensing beam 3 is PE polyethylene plastic, the size is 500mm*10mm*3mm, the upper and lower surfaces are engraved with semicircular grooves with a diameter of 1mm, and the four edges are chamfered with a diameter of 5mm.​​

[0035] Two optical fibers 4, 5 each with four fiber gratings are installed in the groove and are pasted with glue, as shown in the figure. Figure 4 The fiber sensing beam support structure 2, 6, 7, 8 is made of ABS or other light materials, and has a size of 14mm*7mm*7mm, a through hole of 11mm*7mm*4mm at the center position, and four corners with a diameter of 5mm inside, so that the fiber sensing beam 3 can be put in and longitudinally slide freely.

[0036] The above is only a preferred embodiment of the present application, and the present application should not be limited to the content disclosed in the embodiment and the drawings. Any equivalent or modification made without departing from the disclosed spirit falls within the protection scope of the present application.

Claims

1. A fiber optic method for measuring trailing edge deformation of a variable camber airfoil, characterized in that: The fiber optic sensing beam (3) is installed inside the upper skin of the trailing edge structure (1) of the variable camber wing under test through four fiber optic sensing beam support structures (2, 6, 7, 8). The inner surface of the trailing edge structure (1) of the variable camber wing under test is connected to the fiber optic sensing beam support structures (2, 6, 7, 8) with aviation-grade adhesive. The fiber optic sensing beam (3) passes through the middle hole of the fiber optic sensing beam support structures (2, 6, 7, 8). The fiber optic sensing beam (3) is completely fixed to the first fiber optic sensing beam support structure (2) by bolts, but not fixed to the other fiber optic sensing beam support structures (6, 7, 8). The contact surface between the fiber optic sensing beam support structures (2, 6, 7, 8) and the inner wall of the middle hole of the other fiber optic sensing beam support structures (6, 7, 8) should be as smooth as possible and can slide freely. The installation gap between the fiber optic sensing beam support structures (2, 6, 7, 8) and the fiber optic sensing beam (3) is precisely controlled by an optical ranging system, realizing the shape measurement of the trailing edge structure (1) of the variable camber wing under test in the loaded state. Test; A pair of optical fibers (4, 5) are respectively installed inside the upper and lower surfaces of the optical fiber sensing beam (3). Four fiber gratings (16, 17, 18, 19) are etched on one optical fiber (4), and four fiber gratings (20, 21, 22, 23) are etched on the other optical fiber (5). The fiber gratings on the upper and lower surfaces are symmetrically arranged. The tail ends of the pair of optical fibers (4, 5) are respectively connected to one end of a single-core optical cable (10, 11) through an optical fiber connector flange (9). The other end of the single-core optical cable (10, 11) is connected to the fiber grating center wavelength acquisition device (13) through an FC / APC connector (12). The fiber grating center wavelength acquisition device (13) is connected to the shape data processing module (15) through a network cable or a wireless transmission module (14). The change in center wavelength is calculated by the difference method. The Frenet algorithm for center wavelength-strain-shape testing is used. The deformation of the trailing edge structure is calculated only by analyzing the deformation data of the sensing beam.

2. The fiber optic measurement method for trailing edge deformation of a variable camber wing according to claim 1, characterized in that: The fiber optic sensing beam support structure (2, 6, 7, 8) is made of plastic or aluminum alloy and is installed on the inner surface of the upper skin of the trailing edge structure (1) of the variable camber wing under test using aviation-grade epoxy adhesive.

3. The fiber optic measurement method for trailing edge deformation of a variable camber wing according to claim 1, characterized in that: The pair of optical fibers (4, 5) and the optical fiber sensing beam (3) are subjected to high-precision installation and pre-stretching process control. The optical fiber sensing beam (3) is made of plastic, aluminum alloy or stainless steel. The upper and lower surfaces are engraved with semi-circular grooves with a diameter of 0.3 mm. The surface of the semi-circular grooves must be very smooth. The optical fiber is installed in the semi-circular groove. At the same time, prestress is applied at both ends with weights to keep the optical fiber horizontally stretched and prevent the optical fiber from bending. The fiber gratings (16, 17, 18, 19, 20, 21, 22, 23) on the optical fiber are marked with positions. Then, the semi-circular grooves are filled with liquid epoxy glue and cured at room temperature.

4. The fiber optic measurement method for trailing edge deformation of a variable camber wing according to claim 1, characterized in that: The gap between the fiber optic sensing beam (3) and the intermediate hole of the fiber optic sensing beam support structure (2, 6, 7, 8) can be automatically measured. An optical ranging module is installed on the fiber optic sensing beam support structure (2, 6, 7, 8) to ensure that the gap can be adjusted to an ideal state, avoid friction, and bend and deform freely with the deformation of the trailing edge structure.

5. The fiber optic measurement method for trailing edge deformation of a variable camber wing according to claim 1, characterized in that: Before the fiber optic sensing beam (3) is installed, the center wavelength-strain calibration is performed using a calibration beam testing device to obtain the fiber optic grating strain sensitivity coefficient and ensure the accuracy of strain testing.

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