CN113532304A - Wing skin structure health state monitoring method based on quasi-distributed fiber bragg grating - Google Patents
Wing skin structure health state monitoring method based on quasi-distributed fiber bragg grating Download PDFInfo
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01B—MEASURING LENGTH, THICKNESS OR SIMILAR LINEAR DIMENSIONS; MEASURING ANGLES; MEASURING AREAS; MEASURING IRREGULARITIES OF SURFACES OR CONTOURS
- G01B11/00—Measuring arrangements characterised by the use of optical techniques
- G01B11/16—Measuring arrangements characterised by the use of optical techniques for measuring the deformation in a solid, e.g. optical strain gauge
- G01B11/165—Measuring arrangements characterised by the use of optical techniques for measuring the deformation in a solid, e.g. optical strain gauge by means of a grating deformed by the object
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01K—MEASURING TEMPERATURE; MEASURING QUANTITY OF HEAT; THERMALLY-SENSITIVE ELEMENTS NOT OTHERWISE PROVIDED FOR
- G01K11/00—Measuring temperature based upon physical or chemical changes not covered by groups G01K3/00, G01K5/00, G01K7/00 or G01K9/00
- G01K11/32—Measuring temperature based upon physical or chemical changes not covered by groups G01K3/00, G01K5/00, G01K7/00 or G01K9/00 using changes in transmittance, scattering or luminescence in optical fibres
- G01K11/3206—Measuring temperature based upon physical or chemical changes not covered by groups G01K3/00, G01K5/00, G01K7/00 or G01K9/00 using changes in transmittance, scattering or luminescence in optical fibres at discrete locations in the fibre, e.g. using Bragg scattering
Abstract
A wing skin structure health state monitoring method based on a quasi-distributed fiber grating belongs to the technical field of composite material skin health monitoring. The invention aims to solve the problem that the real-time monitoring of the internal damage of the skin cannot be realized in the existing monitoring of the state of the composite material wing skin. The method comprises the following steps: calibrating the central wavelength of each sensing element; the method comprises the steps of driving broadband light into each sensing element, conducting optical signals collected by each sensing element to a fiber grating demodulator through optical fibers, calculating by the fiber grating demodulator according to optical center wavelength measured values of strain sensors, optical center wavelength measured values of temperature compensation sensors and calibration center wavelengths of corresponding sensing elements to obtain strain signals of skin positions of the strain sensors, and determining the health state of a skin structure according to the strain signals. The invention is used for monitoring the skin strain and temperature in real time.
Description
Technical Field
The invention relates to a method for monitoring the health state of a wing skin structure based on a quasi-distributed fiber grating, and belongs to the technical field of composite material skin health monitoring.
Background
Aircraft wings are one of the important structural components of an aircraft, and the wing skin structure is very easy to damage due to erosion of wind and rain and collision of birds during long-term flight. The most common damage forms are cracks, skin falling and the like; for composite skins, ply-to-ply debonding may also occur. These injuries can seriously affect the flight safety of the aircraft.
Therefore, the method has very important practical significance for monitoring the real-time health of the composite material wing skin of the airplane. At present, the common skin structure health state monitoring methods include an X-ray detection method, an infrared thermal imaging technology and the like, and the monitoring process of the methods is complicated and cannot be monitored in real time for a long time.
The fiber grating sensor has the characteristics of high sensitivity, small volume, multipoint measurement, corrosion resistance, strong electromagnetic interference resistance and the like, so that the fiber grating sensor is suitable for real-time monitoring of a wing skin structure. For example, the fiber grating sensor can be adhered to the outer surface of the wing to monitor the strain, bending moment and the like of the skin, but the real-time monitoring of the internal damage of the skin cannot be realized.
Disclosure of Invention
The invention provides a wing skin structure health state monitoring method based on a quasi-distributed fiber grating, aiming at the problem that the real-time monitoring of the internal damage of a skin cannot be realized in the existing composite material wing skin state monitoring.
The invention relates to a wing skin structure health state monitoring method based on a quasi-distributed fiber grating, which comprises the following steps,
embedding 8 to 12 optical fibers in an interlayer between the 8 th layer and the 9 th layer or between the 9 th layer and the 10 th layer of each wing composite material skin structure to form a quasi-distributed fiber grating; n +2 sensing elements are sequentially arranged on each optical fiber at intervals of 20-30 cm, wherein the n +2 sensing elements comprise n strain sensors and two temperature compensation sensors, and the two temperature compensation sensors are positioned on two trisection points of the optical fibers; n is an integer greater than 9;
the monitoring method comprises the following steps:
calibrating the central wavelength of each sensing element;
the method comprises the steps of driving broadband light into each sensing element, conducting optical signals collected by each sensing element to a fiber grating demodulator through optical fibers, calculating by the fiber grating demodulator according to optical center wavelength measured values of strain sensors, optical center wavelength measured values of temperature compensation sensors and calibration center wavelengths of corresponding sensing elements to obtain strain signals of skin positions of the strain sensors, and determining the health state of a skin structure according to the strain signals.
According to the wing skin structure health state monitoring method based on the quasi-distributed fiber bragg gratings, different sensitive wave bands are set for each grating on each optical fiber; and determining the position of the corresponding grating of the corresponding sensing element according to the interval of the light center wavelength measurement value, and determining the health state of the skin structure at the corresponding grating according to the strain signal.
According to the method for monitoring the health state of the wing skin structure based on the quasi-distributed fiber bragg grating,
the method for calculating the strain signal of the skin position where each strain sensor is located comprises the following steps:
ε＝(Δλ1-Δλ2)/αε，
wherein epsilon represents strain signal obtained by calculation of fiber grating demodulator, delta lambda1Representing the difference, Δ λ, between the measured value of the optical centre wavelength of the strain sensor and the nominal centre wavelength2The difference value between the measured value of the optical center wavelength of the temperature compensation sensor and the calibrated center wavelength is represented, and the temperature compensation sensor is positioned on the same optical fiber as the current strain sensor and is closer to the current strain sensor; alpha is alphaεIndicating the strain sensitivity coefficient.
According to the method for monitoring the health state of the wing skin structure based on the quasi-distributed fiber bragg grating,
judging whether the strain sensors are missing or not according to all strain signals output by the fiber bragg grating demodulator, and if the number of the strain signals is n, enabling all the strain sensors to work normally; if the number of the strain signals is less than n, determining that the strain sensor is missing; and determining the optical fiber where the missing strain sensor is located and the grating position of the optical fiber where the missing strain sensor is located according to the section where the optical center wavelength measurement value is located.
According to the method for monitoring the health state of the wing skin structure based on the quasi-distributed fiber bragg grating,
and on the premise of judging that the strain sensors are not missing, if all the strain signals are in a preset strain stable interval, judging that the wings are in a normal working state.
According to the method for monitoring the health state of the wing skin structure based on the quasi-distributed fiber bragg grating,
if the strain signal exceeds a preset strain stable interval, judging the strain change speed of the corresponding strain sensor in different acquisition periods, and if the strain change speed exceeds a preset maximum strain change rate value interval and the strain signal is continuously increased or decreased, judging that the skin at the corresponding position is damaged;
and if the strain signal exceeding the preset strain stability interval returns to the strain stability interval after 10s, judging that the wing is in a normal working state.
According to the method for monitoring the health state of the wing skin structure based on the quasi-distributed fiber bragg grating,
if the strain signal exceeds a preset strain stability interval and the strain change speed is within a preset maximum strain rate value interval, judging whether the strain signal exceeds a maximum strain threshold value, and if so, judging that the skin material is in a failure state;
if the strain signal does not exceed the maximum strain threshold, judging whether the strain signal continuously increases or decreases in different acquisition periods, and if so, judging that the skin has a fault; and otherwise, judging whether the strain signal returns to a preset strain stability interval after 10s, if so, judging that the wing is in a normal working state, and otherwise, judging that the skin has a fault.
According to the method for monitoring the health state of the wing skin structure based on the quasi-distributed fiber bragg grating,
on each wing composite material skin structure, the head ends of all optical fibers are positioned on 8 to 12 equal points of a wing tail end skin contour line, and tail fibers of all the optical fibers are led out to a fiber grating demodulator from 8 to 12 equal points of the wing head end skin contour line.
According to the method for monitoring the health state of the wing skin structure based on the quasi-distributed fiber bragg grating,
the tail fibers of all the optical fibers on the two wings are connected to the fiber bragg grating demodulator after being collected by the line concentration box.
According to the method for monitoring the health state of the wing skin structure based on the quasi-distributed fiber bragg grating,
each optical fiber and the corresponding sensing element are packaged in the capillary tube and are embedded in the interlayer of the skin structure.
The invention has the beneficial effects that: the invention combines the fiber grating measurement with the composite skin health monitoring technology. In the composite material skin structure with the pre-buried interior of quasi-distributed fiber grating sensor, can not only monitor the inside mechanical signal of covering, can also realize carrying out real-time supervision and early warning to typical damage such as the skin structure crackle damage and crack development.
The method of the invention arranges the sensor in the interlayer of the skin, can be used for monitoring the health state of the skin in real time, has accurate monitoring result, and can find the abnormity of the skin state in time, thereby ensuring the flight safety of the airplane.
The invention realizes the real-time monitoring of the strain and the temperature of the composite material skin of the wing, and can greatly improve the reliability and the safety of the skin structure of the airplane.
Drawings
FIG. 1 is a schematic diagram of the pre-embedding of a quasi-distributed fiber grating in an airfoil skin according to the present invention; in the figure 1 denotes the wing skin, 2 denotes the strain sensor, 3 denotes the temperature compensation sensor, 4 denotes the optical fibre, 5 denotes the header;
FIG. 2 is a cross-sectional view of an optical fiber embedded in a wing skin; in the figure, A represents the upper surface of the wing skin, B represents the lower surface of the wing skin, and C represents the optical fiber leading-out end;
FIG. 3 is a system diagram of wing skin structure health monitoring based on quasi-distributed fiber bragg gratings;
FIG. 4 is a flowchart of the operation of a specific implementation of the method of the present invention; in the figure, Y indicates YES and N indicates NO.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It should be noted that the embodiments and features of the embodiments may be combined with each other without conflict.
The invention is further described with reference to the following drawings and specific examples, which are not intended to be limiting.
First embodiment, referring to fig. 1 to 4, the present invention provides a method for monitoring a health status of a wing skin structure based on a quasi-distributed fiber grating, including,
embedding 8 to 12 optical fibers in an interlayer between the 8 th layer and the 9 th layer or between the 9 th layer and the 10 th layer of each wing composite material skin structure to form a quasi-distributed fiber grating; n +2 sensing elements are sequentially arranged on each optical fiber at intervals of 20-30 cm, wherein the n +2 sensing elements comprise n strain sensors and two temperature compensation sensors, and the two temperature compensation sensors are positioned on two trisection points of the optical fibers; n is an integer greater than 9;
the monitoring method comprises the following steps:
calibrating the central wavelength of each sensing element;
the method comprises the steps of driving broadband light into each sensing element, conducting optical signals collected by each sensing element to a fiber grating demodulator through optical fibers, calculating by the fiber grating demodulator according to optical center wavelength measured values of strain sensors, optical center wavelength measured values of temperature compensation sensors and calibration center wavelengths of corresponding sensing elements to obtain strain signals of skin positions of the strain sensors, and determining the health state of a skin structure according to the strain signals.
In the embodiment, the strain sensors are connected in series through optical fibers, are embedded in the composite material skin structure of the wing, and are connected with the grating demodulator through transmission optical fibers. The grating demodulator demodulates the optical signal of the strain sensor into a strain signal, and the strain signal can be transmitted to a pilot cab through a data line, so that flight personnel can monitor the health state of the wing composite material skin structure in real time.
By way of example, 12 strain sensors may be provided on each fiber, four in three groups, separated by two temperature compensation sensors. The strain sensor and the temperature compensation sensor are arranged in this way, so that not only can enough monitoring precision be guaranteed, but also the monitoring cost can be guaranteed not to be too high.
In this embodiment, the strain sensor forms a quasi-distributed fiber bragg grating sensor.
The working principle is as follows: after receiving the optical signal, the strain sensor reflects light with a specific central wavelength into the fiber grating demodulator. When the external environment changes, mainly the pulling and pressing of the external force and the change of the temperature can cause the central wavelength of the strain sensor and the central wavelength of the temperature compensation sensor to shift, and the strain signal can be calculated according to the shift. And then, demodulating the optical signal by using a complex division demodulation technology, and judging the position of the grating corresponding to the wavelength according to the interval where the reflection wavelength is located when the wave band of each grating on the single optical fiber is set to be different, so as to position the strain acquired by different gratings.
The composite material skin can be made of carbon fiber composite materials, the number of layers is 10-16, and the quasi-distributed fiber bragg grating is located between 8-10 layers of the composite material skin structure.
Calibrating central wavelength of fiber grating
Before the wing with the embedded sensor is put into use, the central wavelength of the fiber grating sensing element needs to be calibrated: and (3) standing the wing composite material skin structure for a period of time, and then recording the central wavelength of each strain fiber grating sensor and the central wavelength of each temperature compensation fiber grating sensor to finish the initial wavelength calibration.
Furthermore, setting different sensitive wave bands for each grating on each optical fiber; and determining the position of the corresponding grating of the corresponding sensing element according to the interval of the light center wavelength measurement value, and determining the health state of the skin structure at the corresponding grating according to the strain signal.
In this embodiment, the position of the sensing element corresponding to the grating refers to the position of the center point of the grating where the sensing element is located. That is, the position of the skin corresponding to the sensing element is not an interval, but a spatial coordinate point of the corresponding grating inside the skin.
Still further, the method for calculating the strain signal of the skin position where each strain sensor is located comprises the following steps:
ε＝(Δλ1-Δλ2)/αε，
wherein epsilon represents strain signal obtained by calculation of fiber grating demodulator, delta lambda1Representing the difference, Δ λ, between the measured value of the optical centre wavelength of the strain sensor and the nominal centre wavelength2The difference value between the measured value of the optical center wavelength of the temperature compensation sensor and the calibrated center wavelength is represented, and the temperature compensation sensor is positioned on the same optical fiber as the current strain sensor and is closer to the current strain sensor; alpha is alphaεIndicating the strain sensitivity coefficient.
Wherein Δ λ1And Δ λ2The sensor may be a positive value, which indicates that the sensor is under a tensile load, or a negative value, which indicates that the sensor is under a compressive load.
And when the distances between the two temperature compensation sensors on the same optical fiber and the current strain sensor are equal, selecting the average value of the measurement results of the two temperature compensation sensors as the measurement value of the current temperature compensation sensor.
Further, with reference to fig. 4, it is determined whether the strain sensors are missing according to all the strain signals output by the fiber grating demodulator, and if the number of the strain signals is n, all the strain sensors normally operate; if the number of the strain signals is less than n, determining that the strain sensor is missing; and determining the optical fiber where the missing strain sensor is located and the grating position of the optical fiber where the missing strain sensor is located according to the section where the optical center wavelength measurement value is located.
Further, referring to fig. 4, on the premise that it is determined that the strain sensors are not missing, if all the strain signals are in the preset strain stabilization interval, it is determined that the wing is in a normal working state.
Further, with reference to fig. 4, if the strain signal exceeds the preset strain stability interval, determining the strain change speed of the corresponding strain sensor in different acquisition periods, and if the strain change speed exceeds the preset maximum strain rate interval and the strain signal continuously increases or decreases, determining that the skin at the corresponding position is damaged;
and if the strain signal exceeding the preset strain stability interval returns to the strain stability interval after 10s, judging that the wing is in a normal working state.
Further, referring to fig. 4, if the strain signal exceeds the preset strain stability interval and the strain change speed is within the preset maximum strain change rate value interval, determining whether the strain signal exceeds a maximum strain threshold, and if so, determining that the skin material is in a failure state;
if the strain signal does not exceed the maximum strain threshold, judging whether the strain signal continuously increases or decreases in different acquisition periods, and if so, judging that the skin has a fault; and otherwise, judging whether the strain signal returns to a preset strain stability interval after 10s, if so, judging that the wing is in a normal working state, and otherwise, judging that the skin has a fault.
As shown in fig. 3, in this embodiment, the strain signal obtained by the fiber grating demodulator may be transmitted to the PC, the PC determines the skin state according to the range of the strain signal, and controls the alarm to alarm when the determination result is a fault. The alarm can carry out graded alarm according to different preset damage conditions.
Further, as shown in fig. 1, in each wing composite material skin structure, the head ends of all the optical fibers are located at the equal points 8 to 12 of the contour line of the wing tail end skin, and the tail fibers of all the optical fibers are led out to the fiber bragg grating demodulator from the equal points 8 to 12 of the contour line of the wing head end skin.
Still further, as shown in fig. 3, the pigtails of all the optical fibers on the two wings are connected to the fiber grating demodulator after being collected by the cable collection box.
The line concentration box and the fiber grating demodulator can be arranged under the cabin, the fiber grating demodulator is provided with a corresponding number of optical fiber interfaces, wherein one half of the optical fiber interfaces is connected with the left wing optical fiber, and the other half of the optical fiber interfaces is connected with the right wing optical fiber.
Still further, as shown in fig. 2, each optical fiber and the corresponding sensing element are packaged in a capillary tube, embedded in the interlayer of the skin structure.
In the invention, the monitoring component is required to be pre-embedded into the composite material skin when the composite material skin is produced.
The pre-embedding process comprises the following steps: after the composite material prepreg is laid, the whole optical fiber and the sensor are packaged in the special capillary tube, the special capillary tube is laid between two layers of prepregs with the same direction as the carbon fiber direction of the lower layer of prepreg, the tail fiber is led out from the side face of the composite material laminated plate, and the tail fiber of the sensor can be protected by using a special material sleeve.
The following describes a specific implementation flow of the present invention with reference to fig. 4:
1) firstly, setting two threshold value intervals, namely a maximum strain value interval and a maximum strain change rate value interval, wherein the maximum strain value interval is set according to the maximum strain criterion of the composite material, namely-ec<ε<et. Wherein e istIs a tensile ultimate strain value of the matrix of the composite material, ecIs the compression limit strain value of the composite material matrix. The maximum strain rate interval is set according to a large amount of experimental and simulation data. Can set ecIs-300 mu epsilon/s, etIs 300. mu. epsilon/s. Then-300. mu. epsilon/s<ε<300. mu. epsilon/s. Then, all the strain data obtained by the demodulator are input into a program, and whether the monitoring strain value of the sensor of the missing part exists or not is judged. If there is a deficiency, find out the sensing of the deficiencyThe device number and the optical fiber where the device number and the optical fiber are located indicate that the optical fiber is broken to cause that part of the sensors cannot reflect back wavelength signals, and the main reason of the broken optical fiber is that the optical fiber is damaged due to overlarge external load, and even carbon fiber at the position of the optical fiber is broken, so that the device number and the optical fiber need to be subjected to alarm processing, and the alarm can be set to be a three-level alarm.
2) If the sensor is not in the absence, the next judgment can be performed, and at this time, whether each strain value is in the stable interval needs to be judged. During the operation of the airplane wing, the vibration phenomenon is inevitably generated, which is shown in that the strain value of the sensor fluctuates back and forth in a certain interval, the stable interval of the strain can be set to-50 mu epsilon <50 mu epsilon, and if the strain value is in the stable interval, the wing is in a normal working state.
3) If the strain value is not in the stable region, the change speed of the strain is determined. When the composite material skin is damaged by perforation, debonding and the like, a stress concentration phenomenon is generated at the damaged position, which is shown in that a strain value of the sensor is subjected to a sudden change phenomenon, namely the change speed of the strain value is greatly increased, if the change speed exceeds a set maximum strain rate value interval, whether the absolute value of the strain value has a trend of continuously increasing or decreasing or not needs to be judged, if yes, an alarm is needed, and the alarm is a first-level alarm. If not, judging whether the strain value returns to the normal interval after 10s, if not, alarming is needed, and the alarm is a first-level alarm. If the normal interval can be returned, the health state is achieved.
4) If the change speed does not exceed the set maximum strain rate value interval, whether the strain value exceeds the maximum strain value interval or not needs to be judged, if so, the material starts to lose efficacy, and an alarm needs to be given, wherein the alarm is a secondary alarm. If not, judging whether the absolute value of the strain has the trend of continuous increase or decrease, if so, alarming, wherein the alarming is a secondary alarming. If not, judging whether the strain value returns to the normal interval after 10s, if not, alarming is needed, and the alarming is primary alarming. If yes, the system is in a healthy state.
After all the determinations are carried out, all data need to be stored, and reference is conveniently provided for the subsequent health monitoring.
Although the invention herein has been described with reference to particular embodiments, it is to be understood that these embodiments are merely illustrative of the principles and applications of the present invention. It is therefore to be understood that numerous modifications may be made to the illustrative embodiments and that other arrangements may be devised without departing from the spirit and scope of the present invention as defined by the appended claims. It should be understood that features described in different dependent claims and herein may be combined in ways different from those described in the original claims. It is also to be understood that features described in connection with individual embodiments may be used in other described embodiments.
Claims (10)
1. A wing skin structure health state monitoring method based on quasi-distributed fiber bragg grating is characterized by comprising the following steps,
embedding 8 to 12 optical fibers in an interlayer between the 8 th layer and the 9 th layer or between the 9 th layer and the 10 th layer of each wing composite material skin structure to form a quasi-distributed fiber grating; n +2 sensing elements are sequentially arranged on each optical fiber at intervals of 20-30 cm, wherein the n +2 sensing elements comprise n strain sensors and two temperature compensation sensors, and the two temperature compensation sensors are positioned on two trisection points of the optical fibers; n is an integer greater than 9;
the monitoring method comprises the following steps:
calibrating the central wavelength of each sensing element;
the method comprises the steps of driving broadband light into each sensing element, conducting optical signals collected by each sensing element to a fiber grating demodulator through optical fibers, calculating by the fiber grating demodulator according to optical center wavelength measured values of strain sensors, optical center wavelength measured values of temperature compensation sensors and calibration center wavelengths of corresponding sensing elements to obtain strain signals of skin positions of the strain sensors, and determining the health state of a skin structure according to the strain signals.
2. The method for monitoring the health state of the wing skin structure based on the quasi-distributed fiber bragg gratings according to claim 1, wherein a different sensitive waveband is set for each grating on each optical fiber; and determining the position of the corresponding grating of the corresponding sensing element according to the interval of the light center wavelength measurement value, and determining the health state of the skin structure at the corresponding grating according to the strain signal.
3. The wing skin structure health state monitoring method based on the quasi-distributed fiber bragg grating as claimed in claim 1 or 2,
the method for calculating the strain signal of the skin position where each strain sensor is located comprises the following steps:
ε＝(Δλ1-Δλ2)/αε，
wherein epsilon represents strain signal obtained by calculation of fiber grating demodulator, delta lambda1Representing the difference, Δ λ, between the measured value of the optical centre wavelength of the strain sensor and the nominal centre wavelength2The difference value between the measured value of the optical center wavelength of the temperature compensation sensor and the calibrated center wavelength is represented, and the temperature compensation sensor is positioned on the same optical fiber as the current strain sensor and is closer to the current strain sensor; alpha is alphaεIndicating the strain sensitivity coefficient.
4. The wing skin structure health status monitoring method based on the quasi-distributed fiber grating according to claim 3,
judging whether the strain sensors are missing or not according to all strain signals output by the fiber bragg grating demodulator, and if the number of the strain signals is n, enabling all the strain sensors to work normally; if the number of the strain signals is less than n, determining that the strain sensor is missing; and determining the optical fiber where the missing strain sensor is located and the grating position of the optical fiber where the missing strain sensor is located according to the section where the optical center wavelength measurement value is located.
5. The wing skin structure health status monitoring method based on the quasi-distributed fiber grating according to claim 4,
and on the premise of judging that the strain sensors are not missing, if all the strain signals are in a preset strain stable interval, judging that the wings are in a normal working state.
6. The method for monitoring the health state of the wing skin structure based on the quasi-distributed fiber bragg grating is characterized in that if a strain signal exceeds a preset strain stability interval, the strain change speed of a corresponding strain sensor in different acquisition periods is judged, and if the strain change speed exceeds a preset maximum strain change rate value interval and the strain signal is continuously increased or decreased, the skin at a corresponding position is judged to be damaged;
and if the strain signal exceeding the preset strain stability interval returns to the strain stability interval after 10s, judging that the wing is in a normal working state.
7. The wing skin structure health status monitoring method based on the quasi-distributed fiber grating according to claim 6,
if the strain signal exceeds a preset strain stability interval and the strain change speed is within a preset maximum strain rate value interval, judging whether the strain signal exceeds a maximum strain threshold value, and if so, judging that the skin material is in a failure state;
if the strain signal does not exceed the maximum strain threshold, judging whether the strain signal continuously increases or decreases in different acquisition periods, and if so, judging that the skin has a fault; and otherwise, judging whether the strain signal returns to a preset strain stability interval after 10s, if so, judging that the wing is in a normal working state, and otherwise, judging that the skin has a fault.
8. The wing skin structure health status monitoring method based on the quasi-distributed fiber grating according to claim 7,
on each wing composite material skin structure, the head ends of all optical fibers are positioned on 8 to 12 equal points of a wing tail end skin contour line, and tail fibers of all the optical fibers are led out to a fiber grating demodulator from 8 to 12 equal points of the wing head end skin contour line.
9. The wing skin structure health status monitoring method based on the quasi-distributed fiber grating according to claim 8,
the tail fibers of all the optical fibers on the two wings are connected to the fiber bragg grating demodulator after being collected by the line concentration box.
10. The wing skin structure health status monitoring method based on the quasi-distributed fiber grating according to claim 9,
each optical fiber and the corresponding sensing element are packaged in the capillary tube and are embedded in the interlayer of the skin structure.
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| CN114623776A (en) * | 2022-05-16 | 2022-06-14 | 四川省公路规划勘察设计研究院有限公司 | Tunnel damage prediction method based on tunnel deformation monitoring |
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