CN110006563A - Distributed optical fiber decoupling measurement method for helicopter blade flapping and sway bending moment - Google Patents

Abstract

The invention discloses a distributed optical fiber decoupling measurement method for helicopter blade flapping and swaying bending moment, comprising the following steps: calculating the static moment of the helicopter blade section and calculating the blade section inertia moment; according to the Kirchhoff neutral layer Assuming the principle, preliminarily select the approximate decoupling node positions corresponding to the swaying/swing moment. The distributed fiber grating sensor is pasted along the spanwise direction of the blade at the position of the decoupling node; the strain value corresponding to any position in the chordwise direction of the blade is calculated by applying a flapping/swing vibration load to the helicopter blade; the strain value is fitted, Select the coordinate position corresponding to the zero strain value of the fitting curve. This coordinate position is selected as the decoupling node corresponding to the flapping/swing moment of the helicopter blade. The method can quickly and accurately determine the key position of the decoupling node of the helicopter blade flapping/swinging moment, and can accurately measure the strain of the helicopter blade in different directions under the action of different forms of flapping/swinging moment.

Description

Distributed optical fiber decoupling measurement method for helicopter blade flapping and sway bending moment

technical field

The invention belongs to the technical field of bending moment monitoring for structural health monitoring, and specifically proposes a method for decoupling a helicopter blade flapping and swaying bending moment and an optical fiber sensor layout method based on the Kirchhoff neutral layer principle.

Background technique

Blade machinery is the key equipment in the aerospace industry. In recent years, with the deepening of research, the performance of blade machinery has been continuously improved, and it has continued to develop in the direction of high speed, heavy load, high efficiency and high reliability. The requirements are also getting higher and higher. As the core component of the helicopter, the blade is the main load-bearing object. Most of the damage accidents (cracks, breaks, etc.) of the blade are caused by vibration, such as the waving motion and the swaying motion of the helicopter blade. Therefore, monitoring the state of helicopter blades is an important method to evaluate the design quality, dynamic monitoring of mechanical damage, mechanical fault diagnosis and prediction.

At present, the research on structural deformation state monitoring methods at home and abroad mainly includes photoelectric measurement systems and non-photoelectric measurement systems. Among them, photoelectric methods include photogrammetry, photogrammetry, laser trackers, etc.; Sensors, displacement sensors, fiber optic sensors, etc. The photoelectric measurement system is generally based on image capture and laser scanning tracking. The target point or light source is arranged at a given position of the target structure. The measurement system perceives the target point or light source in an imaging manner. According to the principle of distance correction, reflection principle and photogrammetry The principle is to obtain the spatial position of the target point through coordinate transformation, with fast response speed and high precision. However, this type of method is only suitable for structures with relatively flat shape and simple form, and has poor anti-interference ability. For flexible structures with large deformation, the light transmission is easily affected, and the reliability of the measurement system is difficult to guarantee. Non-photoelectric measurement systems are mainly based on contact measurement methods, based on strain information or curvature information, adapt to more complex structures, and can directly embed sensors into structural materials, with stable measurement performance. The contact measurement method is suitable for the deformation and reconstruction of more complex structures, and the measurement performance is relatively stable.

Therefore, in view of the harsh working environment of helicopter blades and the characteristics of more external interference, a non-photoelectric contact measurement method is selected, and a relatively advanced and increasingly mature fiber Bragg grating sensor is selected as the test element. Compared with traditional sensors, Fiber Bragg grating sensors have extremely high sensitivity and precision, and are intrinsically safe, strong in anti-electromagnetic interference, high in dielectric strength, small in size, and wide in frequency. Due to the asymmetry of the blade profile, there will be a coupling problem in the actual monitoring process of the bending moment and the swaying moment. In view of the shortcomings of the current monitoring method of the blade bending moment, it is necessary to study without a lot of prior knowledge. It can be applied to conventional fiber grating demodulators with lower sampling frequency, and is a new method that is simple, fast, convenient, reliable and practical.

SUMMARY OF THE INVENTION

Purpose of the invention: In order to overcome the deficiencies in the prior art, the present invention provides a method for decoupling the flapping and swaying bending moment of a helicopter blade and an optical fiber sensor based on Kirchhoff's neutral layer principle.

Technical scheme: In order to realize the above-mentioned purpose, the technical scheme adopted in the present invention is:

A distributed optical fiber decoupling measurement method for helicopter blade flapping and sway bending moment, comprising the following steps:

Step 1. Calculate the static moment of the elliptical part of the helicopter blade section according to the shape of the helicopter blade;

Step 2: Calculate the position of the centroid of the blade section according to the static moment theorem, determine the main coordinate system of the blade section according to the centroid position, and calculate the moment of inertia of the section according to the determined coordinate system;

Step 3. According to Kirchhoff's neutral layer hypothesis, preliminarily select the approximate decoupling node position corresponding to the waving/swing bending moment, and paste the fiber grating sensor along the span of the blade at the decoupling node position, and the fiber grating sensor axis The direction should be consistent with the spanwise direction of the blade;

Distributed fiber grating sensors are arranged on the blade structure, the root of the blade is fixed, a section is selected, and three fiber grating sensors are sequentially arranged on the blade surface passing through the main coordinate system of the section, which are fiber grating sensor-FBG1 and fiber grating sensor respectively. 2 FBG2, fiber grating sensor 3 FBG3, according to Kirchhoff’s neutral layer hypothesis, the structural neutral layer stress is zero, and the transverse axis and longitudinal axis of the main coordinate system of the section are approximately regarded as the initial neutral layer of the section, The intersection of the coordinate axis and the blade profile is the fiber grating sensor arrangement point, and the fiber grating sensor one FBG1, the fiber grating sensor two FBG2, and the fiber grating sensor three FBGs are serially connected by using fiber jumpers to form a wavelength division multiplexing type transmission. sense network;

Step 4: Load the helicopter blade with swaying vibration, measure the strain value of the finite point of its profile, and continuously collect the spanwise strain data of the discrete points of the blade to obtain different positions in the chordwise direction of the blade. The corresponding spanwise strain distribution curve equation, and then invert the strain value corresponding to any position in the chordwise direction of the blade;

Swing force and sway force are applied to the blade, and the reflected wavelength signal of the optical fiber on the three nodes is collected, which is converted into a strain value according to the principle of fiber grating strain sensing, and then the strain value is subjected to strain continuity to obtain the strain change. Equation, the strain value at any point of the section is obtained, and the relationship between strain and grating wavelength is shown in formula (6)

Δλ _B /λ _B =(1-P _e )Δε (6)

Among them, Δλ _B is the change of the center wavelength of the fiber grating, λ _B is the center wavelength of the fiber grating, P _e is the effective elastic-optical coefficient of the fiber grating, Δε is the change of the strain value, and the relationship between the strain value and the chord position is shown in Eq. (7 ) shown

Among them, Δλ _B is the change of the center wavelength of the fiber grating, λ _B is the center wavelength of the fiber grating, P _e is the effective elastic-optical coefficient of the fiber grating, Δε is the change of the strain value, and the relationship between the strain value and the chord position is shown in Eq. (7 ) shown

Among them, ε(γ) is the strain value, _εa is the chordwise initial strain value, εb is the chordwise end strain value, γ is the _chordwise position, _γa is the chordwise starting position, and Δγ is the chordwise length difference ;

Step 5: Fit the strain value corresponding to the discrete position of the blade chord direction measured in step 4, select the coordinate position corresponding to the strain value of the fitting curve when it is zero, and obtain the decoupling node of the swaying and swaying vibration. The position of the vibration decoupling node is corrected, and then the specific layout position of the blade profile fiber grating sensor is determined, that is, the swing measurement node and the swing vibration measurement node;

When the helicopter blade is rotating, it is mainly affected by the _centrifugal force F _c and the profile bending moment Me. According to the force balance, it can be obtained:

F _c = ∫ _A σdA = 0 (8)

Me = ∫ _{A zσdA} (9)

Among them, σ is the stress, A is the section area, z is the distance from the micro-element surface to the neutral axis, and there is a coordinate relationship Z _σ =Z ₀ +ΔZ in the blade section coordinate system;

When the blade is subjected to the bending moment load, a part of the section reaches the ultimate stress σ _m , and the relationship between the stress and Z _σ can be obtained according to the geometric relationship:

Combined with the coordinate relationship, substituting Equation (10) into Equation (8) can get

Substitute variables to get

Substitute equations (10) and (12) into equation (9) to get

Among them, l is the distance from the yielding part of the profile to the neutral axis, △z is the offset of the neutral axis, Me is the bending moment of the profile, _{Sk is} the static moment of the yielding part of the blade to the neutral axis of the profile, S _t is the static moment of the elastic part of the blade section to the neutral axis of the section, I _t is the moment of inertia of the elastic section of the blade section to the neutral axis of the section, A _k is the area of the yield part of the blade section, A _t is the area of the elastic part of the section, After the neutral axis offset △z is calculated, the neutral axis position of the blade is corrected, and the node position of the FBG sensor is finally determined.

Preferably: the static moment of the elliptical part of the helicopter blade section is:

Among them, S _y is the transverse (y-direction) static moment, S _z is the longitudinal (z-direction) static moment, bi and _hi are the transverse and longitudinal ellipse formula coefficients in the blade section coordinate system, respectively _{, and θ i is} the blade section The span occupied by the ellipse relative to the origin of the coordinate in the coordinate system, θ is the angle, and n is the number of figures divided by the section.

Preferably: in step 2, the coordinates y and z of the centroid of the section are respectively

Among them, y is the transverse coordinate position of the centroid of the section, z is the longitudinal coordinate position of the centroid of the section, A is the area of the section, A _i is the area of each part of the section, _yi is the transverse coordinate of the centroid of the shape of each section of the section, zi _i is the longitudinal coordinate of the centroid of the shape of each part of the section,

After determining the position of the centroid, the main coordinate system of the blade section is established with the centroid as the origin, and then the moment of inertia of the blade section is calculated according to the main coordinate system:

I=∫ _A ρ ² dA=∫ _A (y ² +z ² )dA=I _y +I _z (5)

Among them, I is the total inertia moment of the section, A is the section area, ρ is the distance from the micro-area to the coordinate origin, I _y is the transverse inertia moment, I _z is the longitudinal inertia moment, y and z are the coordinates of the blade micro-element section, respectively value.

Optimal: When the blade is deformed by bending and yielding, the neutral axis of the profile will be offset, so it is necessary to determine the offset Δz to correct the node position of the FBG sensor, thereby improving the decoupling accuracy.

Preferably: when the ratio of flapping strain to pendulum strain is less than the chord aspect ratio of the helicopter blade, it is considered that the decoupling of flapping and pendulum vibration can be achieved at this node;

The first stage: apply a waving load to the helicopter blade, and measure the strain values of each node of fiber grating sensor one FBG1, fiber grating sensor two FBG2, and fiber grating sensor three FBG3 are ε ₁ , ε ₂ , ε ₃ , and then give A swing load is applied to the blade, and the measured strain values of each node of fiber grating sensor one FBG1, fiber grating sensor two FBG2, and fiber grating sensor three FBG3 are ε ₁₁ , ε ₂₂ , and ε ₃₃ respectively;

The second stage: Set the FBG2 node of the fiber grating sensor as the estimated swing decoupling point. When the swing load is not more than half of the swing load, the ratio of the swing strain to the swing strain is compared with the chord length of the helicopter blade. Compared with the length ratio, if Then it is considered that this point meets the requirements of waving decoupling;

The third stage: set the node of fiber grating sensor 1 FBG1 and fiber grating sensor 3 FBG3 as the estimated swing decoupling point. In this case, it is necessary to compare the corresponding nodes of fiber grating sensor 1 FBG1 and fiber optic sensor 3 FBG3 respectively. The best decoupling accuracy is finally confirmed.

Preferably: when the swing load is not more than half of the swing load, compare the ratio of the swing strain to the swing strain at the two nodes, i.e. , compare the smaller of it with the ratio of the chord to the span of the helicopter blade, if it is less than It is considered that this point satisfies the requirement of decoupling of pendulum vibration.

Compared with the prior art, the present invention has the following beneficial effects:

Based on Kirchhoff's neutral layer assumption principle, this method can quickly and accurately determine the key positions of the helicopter blade flapping/swing moment decoupling nodes, without arranging a large number of strain gauges on the blade surface, and with the help of a large number of experiments, iteratively determined step by step Decouple node locations. Clever use of the unique strain-direction sensitivity characteristics of fiber grating sensors can accurately measure the strain of helicopter blades in different directions under the action of different forms of waving/swing bending moments, overcoming the poor sensitivity of traditional resistance strain sensors to the direction of strain response. limit. In addition, the distributed optical fiber strain sensing network does not require a large number of signal transmission cables like the resistance strain sensing system, which is not only conducive to information transmission under rotating conditions, but also avoids the need for the natural frequency of the blade structure under test. Mechanical properties are negatively affected.

Description of drawings

Fig. 1 Layout of distributed fiber Bragg grating sensor;

Figure 2 Blade section coordinate system;

Detailed ways

Below in conjunction with the accompanying drawings and specific embodiments, the present invention will be further clarified. It should be understood that these examples are only used to illustrate the present invention and are not used to limit the scope of the present invention. Modifications in the form of valence all fall within the scope defined by the appended claims of the present application.

A distributed optical fiber decoupling measurement method for helicopter blade flapping and sway bending moment. The method includes the following steps: step 1: deriving the static moment of the helicopter blade section; step 2: calculating the centroid position of the helicopter blade section, determining the main coordinate system of the blade section according to the centroid position, and calculating according to the determined coordinate system Blade profile moment of inertia; Step 3: According to Kirchhoff's neutral layer assumption, preliminarily select the approximate decoupling node positions corresponding to the flapping/swing moment. The fiber grating sensor is pasted along the spanwise direction of the blade at the position of the decoupling node, and the axial direction of the fiber grating sensor should be consistent with the spanwise direction of the blade; Step 4: By applying a waving/swinging load to the blade, a number of discrete cross sections are measured. The strain value of the position, and the collected spanwise strain data of discrete points of the blade are continuous, the spanwise strain distribution curve equation corresponding to different positions in the chordwise direction of the blade can be obtained, and then the blade can be obtained by inversion calculation. The strain value corresponding to any position in the chord direction; Step 5: Fit the strain values corresponding to several discrete positions in the chord direction of the blade measured in step 4, and select the coordinate position corresponding to the fitting curve when the strain value is zero. This coordinate position is selected as the decoupling node corresponding to the flapping/swing moment of the helicopter blade. Based on Kirchhoff's neutral layer assumption principle, this method can quickly and accurately determine the key positions of the helicopter blade flapping/swing moment decoupling nodes, without arranging a large number of strain gauges on the blade surface, and with the help of a large number of experiments, iteratively determined step by step Decouple node locations. Clever use of the unique strain-direction sensitivity characteristics of fiber grating sensors can accurately measure the strain of helicopter blades in different directions under the action of different forms of waving/swing bending moments, overcoming the poor sensitivity of traditional resistance strain sensors to the direction of strain response. limit. In addition, the distributed optical fiber strain sensing network does not require a large number of signal transmission cables like the resistance strain sensing system, which is not only conducive to information transmission under rotating conditions, but also avoids the need for the natural frequency of the blade structure under test. Mechanical properties are negatively affected.

Specifically include the following steps:

Step 1. Derive the static moment of the helicopter blade section, and use the carbon fiber material helicopter rotor blade scale model; the typical section of the helicopter blade can generally be regarded as a combination of two-dimensional figures such as ellipses and triangles. Among them, the static moment of the elliptical part of the helicopter blade section is:

Among them, S _y is the transverse (y-axis direction) static moment, S _z is the longitudinal (z-axis direction) static moment, b _i and _{hi are the ellipse formula coefficients in the blade section coordinate system, respectively, and θ i is} the blade section coordinate The span occupied by the ellipse relative to the origin of the coordinates in the system.

Step 2: Calculate the centroid position of the blade section, determine the main coordinate system of the blade section according to the centroid position, and calculate the moment of inertia of the section according to the determined coordinate system;

According to the static moment theorem, the centroid coordinates y and z of the section are respectively

After determining the position of the centroid, the main coordinate system of the blade section can be established with the centroid as the origin, and then the moment of inertia of the blade section is calculated according to the main coordinate system.

I=∫ _A ρ ² dA=∫ _A (y ² +z ² )dA=I _y +I _z (5)

That is, the moment of inertia of the section is the sum of the moments of inertia of each part of the section. Among them, I is the total moment of inertia of the section, I _y is the transverse (y-axis direction) moment of inertia, I _z is the longitudinal (z-axis direction) moment of inertia, ρ is the distance from the micro area to the coordinate origin, A is the section area, y and z are the coordinate values of the blade micro-element profile, respectively.

Step 3: According to Kirchhoff's neutral layer hypothesis, preliminarily select the approximate decoupling node positions corresponding to the waving/swing bending moment. Paste the fiber grating sensor along the spanwise direction of the blade at the position of the decoupling node, and the axial direction of the fiber grating sensor should be consistent with the spanwise direction of the blade;

As shown in Figure 1, a distributed fiber grating sensor is arranged on the blade structure, the root of the blade is fixed, a section is selected, and three fiber grating sensors are sequentially arranged on the blade surface passing through the main coordinate system of the section. According to Kirchhoff's neutral layer hypothesis, the structural neutral layer stress is zero, and the x-axis and y-axis of the main coordinate system of the profile can be approximately regarded as the initial neutral layer of the profile, and the intersection of the coordinate axis and the blade profile It is the arrangement point of the fiber grating sensor. The FBG1, FBG2, and FBG3 are serially connected by fiber jumpers to form a wavelength division multiplexing sensor network. The sampling frequency of the fiber grating demodulator used in the sensor network is 1KHz.

Step 4: Load the helicopter blade with swaying vibration, measure the strain value at the finite point of its profile, and serialize the collected spanwise strain data at discrete points of the blade, so that the difference in the chordwise direction of the blade can be obtained. The spanwise strain distribution curve equation corresponding to the position, and then the strain value corresponding to any position in the chordwise direction of the blade can be inverted;

Swing force and sway force are applied to the blade, and the reflected wavelength signal of the optical fiber on the three nodes is collected, which is converted into a strain value according to the principle of fiber grating strain sensing, and then the strain value is subjected to strain continuity to obtain the strain change. Equation, the strain value at any point of the section is obtained, and the relationship between strain and grating wavelength is shown in formula (6)

Δλ _B /λ _B =(1-P _e )Δε (6)

Among them, Δλ _B is the variation of the center wavelength of the fiber grating, λ _B is the center wavelength of the fiber grating, P _e is the effective elastic-optical coefficient of the fiber grating, and Δε is the strain value. The relationship between strain value and chord position is shown in formula (7)

Step 5: Fit the strain values corresponding to several discrete positions in the blade chord direction measured in step 4, and select the coordinate position corresponding to the fitting curve when the strain value is zero to obtain the swinging vibration decoupling node. Then the node position is corrected, and then the specific layout position of the blade profile fiber grating sensor can be determined, that is, the swing measurement node and the sway measurement node.

Among them, when the blade is deformed by bending and yielding, the neutral axis of the profile will be offset, so it is necessary to determine the offset △z and correct the node position of the FBG sensor, thereby improving the decoupling accuracy.

In the rotating state of the helicopter blade, it is mainly affected by the _centrifugal force F _c and the bending moment Me. According to the force balance, it can be obtained:

F _c = ∫ _A σdA = 0 (8)

Me = ∫ _{A zσdA} (9)

Among them, σ is the stress, A is the cross-sectional area, and z is the distance from the element surface to the neutral axis. The blade section coordinate system is established as shown in Figure 2, and there is a coordinate relationship Z _σ =Z ₀ +ΔZ.

When the blade is subjected to the bending moment load, a part of the section reaches the ultimate stress σ _m , and the relationship between the stress and Z _σ can be obtained according to the geometric relationship

Combined with the coordinate relationship, substituting Equation (10) into Equation (8) can get

Substitute variables to get

Substitute equations (10) and (12) into equation (9) to get

Among them, l is the distance from the yielding part of the profile to the neutral axis, △z is the offset of the neutral axis, Me is the bending moment of the profile, _{Sk is} the static moment of the yielding part of the blade to the neutral axis of the profile, S _t is the static moment of the elastic part of the blade section to the neutral axis of the section, I _t is the moment of inertia of the elastic section of the blade section to the neutral axis of the section, _Ak is the area of the yield part of the blade section, and At _is the area of the elastic part of the section. After the neutral axis offset △z is calculated, the neutral axis position of the blade is corrected, and the node position of the FBG sensor is finally determined.

When the blade is subjected to the sway force, if the ratio of the flapping strain to the sway strain at the node is 0, the sway force is considered to be decoupled. However, this requirement is often difficult to achieve in practical engineering. Therefore, when the ratio of the flapping strain to the swaying strain is less than the chord aspect ratio of the helicopter blade, it can be considered that the decoupling of flapping and swaying can be achieved at this node.

The first stage: a swing load is applied to the blade, and the measured strain values of each node of FBG1, FBG2, and FBG3 are ε ₁ , ε ₂ , and ε ₃ , respectively. A swing load is then applied to the blade, and the measured strain values of each node of FBG1, FBG2, and FBG3 are ε ₁₁ , ε ₂₂ , and ε ₃₃ , respectively;

The second stage: Set the FBG2 node as the estimated swing decoupling point. When the swing load is not more than half of the swing load, compare the ratio of the swing strain to the swing strain with the ratio of the chord to the extension of the helicopter blade compare, if Then it is considered that this point meets the requirements of waving decoupling;

The third stage: Set the nodes FBG1 and FBG3 as the estimated swing decoupling points. In this case, it is necessary to compare the decoupling accuracy corresponding to the nodes where FBG1 and FBG3 are located, and finally confirm the optimal swing decoupling point. When the swing load is not more than half of the swing load, compare the ratio of swing strain to swing strain at the two nodes, i.e. , compare the smaller of it with the ratio of the chord to the span of the helicopter blade, if it is less than It is considered that this point satisfies the requirement of decoupling of pendulum vibration.

The above is only the preferred embodiment of the present invention, it should be pointed out that: for those skilled in the art, without departing from the principle of the present invention, several improvements and modifications can also be made, and these improvements and modifications are also It should be regarded as the protection scope of the present invention.

Claims (7)

1. a method for measuring the distributed optical fiber decoupling of helicopter blade waving and sway bending moment, is characterized in that, comprises the following steps: Step 1. Calculate the static moment of the elliptical part of the helicopter blade section according to the shape of the helicopter blade; Step 2: Calculate the position of the centroid of the blade section according to the static moment theorem, determine the main coordinate system of the blade section according to the centroid position, and calculate the moment of inertia of the section according to the determined coordinate system; Step 3. According to Kirchhoff's neutral layer hypothesis, preliminarily select the approximate decoupling node position corresponding to the waving/swing bending moment, and paste the fiber grating sensor along the span of the blade at the decoupling node position, and the fiber grating sensor axis The direction should be consistent with the spanwise direction of the blade; Distributed fiber grating sensors are arranged on the blade structure, the root of the blade is fixed, a section is selected, and three fiber grating sensors are sequentially arranged on the blade surface passing through the main coordinate system of the section, which are fiber grating sensor-FBG1 and fiber grating sensor respectively. 2 FBG2, fiber grating sensor 3 FBG3, according to Kirchhoff’s neutral layer hypothesis, the structural neutral layer stress is zero, and the transverse axis and longitudinal axis of the main coordinate system of the section are approximately regarded as the initial neutral layer of the section, The intersection of the coordinate axis and the blade profile is the fiber grating sensor arrangement point, and the fiber grating sensor one FBG1, the fiber grating sensor two FBG2, and the fiber grating sensor three FBGs are serially connected by using fiber jumpers to form a wavelength division multiplexing type transmission. sense network; Step 4: Load the helicopter blade with swaying vibration, measure the strain value of the finite point of its profile, and continuously collect the spanwise strain data of the discrete points of the blade to obtain different positions in the chordwise direction of the blade. The corresponding spanwise strain distribution curve equation, and then invert the strain value corresponding to any position in the chordwise direction of the blade; Swing force and sway force are applied to the blade, and the reflected wavelength signal of the optical fiber on the three nodes is collected, which is converted into a strain value according to the principle of fiber grating strain sensing, and then the strain value is subjected to strain continuity to obtain the strain change. Equation, the strain value at any point of the section is obtained, and the relationship between strain and grating wavelength is shown in formula (6) Δλ _B /λ _B =(1-P _e )Δε (6) Among them, Δλ _B is the change of the center wavelength of the fiber grating, λ _B is the center wavelength of the fiber grating, P _e is the effective elastic-optical coefficient of the fiber grating, Δε is the change of the strain value, and the relationship between the strain value and the chord position is shown in Eq. (7 ) shown Among them, ε(γ) is the strain value, _εa is the chordwise initial strain value, εb is the chordwise end strain value, γ is the _chordwise position, _γa is the chordwise starting position, and Δγ is the chordwise length difference ; Step 5: Fit the strain value corresponding to the discrete position of the blade chord direction measured in step 4, select the coordinate position corresponding to the strain value of the fitting curve when it is zero, and obtain the decoupling node of the swaying and swaying vibration. The position of the vibration decoupling node is corrected, and then the specific layout position of the blade profile fiber grating sensor is determined, that is, the swing measurement node and the swing vibration measurement node; When the helicopter blade is rotating, it is mainly affected by the _centrifugal force F _c and the profile bending moment Me. According to the force balance, it can be obtained: F _c = ∫ _A σdA = 0 (8) Me = ∫ _{A zσdA} (9) Among them, σ is the stress, A is the section area, z is the distance from the micro-element surface to the neutral axis, and there is a coordinate relationship Z _σ =Z ₀ +ΔZ in the blade section coordinate system; When the blade is subjected to the bending moment load, a part of the section reaches the ultimate stress σ _m , and the relationship between the stress and Z _σ can be obtained according to the geometric relationship: Combined with the coordinate relationship, substituting Equation (10) into Equation (8) can get Substitute variables to get Substitute equations (10) and (12) into equation (9) to get Among them, l is the distance from the yielding part of the profile to the neutral axis, △z is the offset of the neutral axis, Me is the bending moment of the profile, _{Sk is} the static moment of the yielding part of the blade to the neutral axis of the profile, S _t is the static moment of the elastic part of the blade section to the neutral axis of the section, I _t is the moment of inertia of the elastic section of the blade section to the neutral axis of the section, A _k is the area of the yield part of the blade section, A _t is the area of the elastic part of the section, After the neutral axis offset △z is calculated, the neutral axis position of the blade is corrected, and the node position of the FBG sensor is finally determined. 2. according to the described helicopter blade of claim 1, it is characterized in that: the static moment of the helicopter blade section ellipse part is: Among them, S _y is the transverse static moment, S _z is the longitudinal static moment, b _i , _{hi are the transverse and longitudinal ellipse formula coefficients in the blade section coordinate system, respectively, θ i is} the relative coordinate origin of the ellipse in the blade section coordinate system Occupy the span, θ is the angle, and n is the number of figures divided by the section. 3. according to the described helicopter blade waving and swaying bending moment distributed optical fiber decoupling measurement method of claim 2, it is characterized in that: in step 2, the profile centroid coordinates y and z are respectively Among them, y is the transverse coordinate position of the centroid of the section, z is the longitudinal coordinate position of the centroid of the section, A is the area of the section, A _i is the area of each part of the section, _yi is the transverse coordinate of the centroid of the shape of each section of the section, zi _i is the longitudinal coordinate of the centroid of the shape of each part of the section, After determining the position of the centroid, the main coordinate system of the blade section is established with the centroid as the origin, and then the moment of inertia of the blade section is calculated according to the main coordinate system: I=∫ _A ρ ² dA=∫ _A (y ² +z ² )dA=I _y +I _z (5) Among them, I is the total moment of inertia of the section, A is the area of the section, ρ is the distance from the micro-area to the origin of the coordinates, I _y is the transverse moment of inertia, and I _z is the longitudinal moment of inertia. 4. The distributed optical fiber decoupling measurement method of helicopter blade waving and sway bending moment according to claim 3, characterized in that: when the blade undergoes bending and yielding deformation, the neutral axis of the profile will be offset, so it is necessary to determine The offset △z corrects the node position of the FBG sensor, thereby improving the decoupling accuracy. 5. The method for measuring the distributed optical fiber decoupling between the swinging and swaying bending moment of a helicopter blade according to claim 4, is characterized in that: when the ratio of swinging strain and swinging strain is less than the chord aspect ratio of the helicopter blade, it is considered that the node can be Realize the decoupling of swing and swing; The first stage: apply a waving load to the propeller, and measure the strain values of each node of fiber grating sensor 1 FBG1, fiber grating sensor 2 FBG2, and fiber grating sensor 3 FBG3, respectively ε ₁ , ε ₂ , ε ₃ , and then apply the propeller A swing load is applied to the blade, and the measured strain values of each node of fiber grating sensor 1 FBG1, fiber grating sensor 2 FBG2, and fiber grating sensor 3 FBG3 are ε ₁₁ , ε ₂₂ , ε ₃₃ , respectively; The second stage: Set the FBG2 node of the fiber grating sensor as the estimated swing decoupling point. When the swing load is not more than half of the swing load, the ratio of the swing strain to the swing strain is compared with the chord length and spread of the helicopter blade. Compared with the length ratio, if Then it is considered that this point meets the requirements of waving decoupling; The third stage: set the node of fiber grating sensor 1 FBG1 and fiber grating sensor 3 FBG3 as the estimated swing decoupling point. In this case, it is necessary to compare the corresponding nodes of fiber grating sensor 1 FBG1 and fiber optic sensor 3 FBG3 respectively. The best decoupling accuracy is finally confirmed. 6. The method for measuring the distributed optical fiber decoupling of helicopter blade flapping and swaying bending moment according to claim 5, characterized in that: when the flapping load is not more than half of the swaying load, compare the flapping strain and the swaying strain at the two nodes The ratio is , compare the smaller of it with the ratio of the chord to the span of the helicopter blade, if it is less than It is considered that this point satisfies the requirement of decoupling of pendulum vibration. 7 . The distributed optical fiber decoupling measurement method for helicopter blade flapping and sway bending moment according to claim 6 , wherein the sampling frequency of the fiber grating demodulator used in the sensing network is 1KHz. 8 .