CN110006563B - Distributed optical fiber decoupling measurement method for blade flapping and shimmy bending moment of helicopter - Google Patents
Distributed optical fiber decoupling measurement method for blade flapping and shimmy bending moment of helicopter Download PDFInfo
- Publication number
- CN110006563B CN110006563B CN201910254397.XA CN201910254397A CN110006563B CN 110006563 B CN110006563 B CN 110006563B CN 201910254397 A CN201910254397 A CN 201910254397A CN 110006563 B CN110006563 B CN 110006563B
- Authority
- CN
- China
- Prior art keywords
- blade
- section
- strain
- shimmy
- flapping
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 238000005452 bending Methods 0.000 title claims abstract description 41
- 239000013307 optical fiber Substances 0.000 title claims abstract description 18
- 238000000691 measurement method Methods 0.000 title claims abstract description 11
- 239000000835 fiber Substances 0.000 claims abstract description 70
- 230000007935 neutral effect Effects 0.000 claims abstract description 45
- 230000003068 static effect Effects 0.000 claims abstract description 23
- 238000000034 method Methods 0.000 claims abstract description 22
- 102100022116 F-box only protein 2 Human genes 0.000 claims description 13
- 101000824158 Homo sapiens F-box only protein 2 Proteins 0.000 claims description 13
- 102100026080 F-box only protein 44 Human genes 0.000 claims description 12
- 102100024513 F-box only protein 6 Human genes 0.000 claims description 12
- 101000913298 Homo sapiens F-box only protein 44 Proteins 0.000 claims description 12
- 101001052796 Homo sapiens F-box only protein 6 Proteins 0.000 claims description 12
- NAWXUBYGYWOOIX-SFHVURJKSA-N (2s)-2-[[4-[2-(2,4-diaminoquinazolin-6-yl)ethyl]benzoyl]amino]-4-methylidenepentanedioic acid Chemical compound C1=CC2=NC(N)=NC(N)=C2C=C1CCC1=CC=C(C(=O)N[C@@H](CC(=C)C(O)=O)C(O)=O)C=C1 NAWXUBYGYWOOIX-SFHVURJKSA-N 0.000 claims description 3
- 230000008859 change Effects 0.000 claims description 3
- 238000005070 sampling Methods 0.000 claims description 3
- 238000005259 measurement Methods 0.000 abstract description 7
- 230000009471 action Effects 0.000 abstract description 3
- 238000012544 monitoring process Methods 0.000 description 7
- 230000035945 sensitivity Effects 0.000 description 4
- 230000005540 biological transmission Effects 0.000 description 3
- 238000011160 research Methods 0.000 description 3
- 230000004044 response Effects 0.000 description 3
- 238000012360 testing method Methods 0.000 description 3
- 230000008878 coupling Effects 0.000 description 2
- 238000010168 coupling process Methods 0.000 description 2
- 238000005859 coupling reaction Methods 0.000 description 2
- 230000007547 defect Effects 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 230000008054 signal transmission Effects 0.000 description 2
- 229920000049 Carbon (fiber) Polymers 0.000 description 1
- 241000764238 Isis Species 0.000 description 1
- 230000001133 acceleration Effects 0.000 description 1
- 230000006978 adaptation Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 239000004917 carbon fiber Substances 0.000 description 1
- 238000006243 chemical reaction Methods 0.000 description 1
- 239000008358 core component Substances 0.000 description 1
- 238000012937 correction Methods 0.000 description 1
- 238000013461 design Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 238000006073 displacement reaction Methods 0.000 description 1
- 230000036541 health Effects 0.000 description 1
- 238000003384 imaging method Methods 0.000 description 1
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 description 1
- 230000010355 oscillation Effects 0.000 description 1
Images
Classifications
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01L—MEASURING FORCE, STRESS, TORQUE, WORK, MECHANICAL POWER, MECHANICAL EFFICIENCY, OR FLUID PRESSURE
- G01L1/00—Measuring force or stress, in general
- G01L1/24—Measuring force or stress, in general by measuring variations of optical properties of material when it is stressed, e.g. by photoelastic stress analysis using infrared, visible light, ultraviolet
- G01L1/242—Measuring force or stress, in general by measuring variations of optical properties of material when it is stressed, e.g. by photoelastic stress analysis using infrared, visible light, ultraviolet the material being an optical fibre
- G01L1/246—Measuring force or stress, in general by measuring variations of optical properties of material when it is stressed, e.g. by photoelastic stress analysis using infrared, visible light, ultraviolet the material being an optical fibre using integrated gratings, e.g. Bragg gratings
Landscapes
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Length Measuring Devices By Optical Means (AREA)
Abstract
The invention discloses a distributed optical fiber decoupling measurement method for blade flapping and shimmy bending moment of a helicopter, which comprises the following steps of: calculating the static moment of the section of the helicopter blade and calculating the inertia moment of the section of the blade; according to the hypothesis principle of kirchhoff neutral layer, roughly decoupling node positions corresponding to the flapping/shimmy bending moment are preliminarily selected. Pasting a distributed fiber bragg grating sensor at the position of the decoupling node along the span direction of the paddle; calculating to obtain a strain value corresponding to any chord direction position of the blade by applying flapping/shimmy load to the helicopter blade; and fitting the strain value, and selecting the corresponding coordinate position when the strain value of the fitted curve is zero. And selecting the coordinate position as a decoupling node corresponding to the flapping/shimmy bending moment of the helicopter blade. The method can quickly and accurately determine the key position of the bending moment decoupling node of the flapping/shimmy of the helicopter blade, and can realize accurate measurement of the strain of the helicopter blade in different directions under the action of different forms of flapping/shimmy bending moments.
Description
Technical Field
The invention belongs to the technical field of bending moment monitoring of structural health monitoring, and particularly provides a kirchhoff neutral layer principle-based helicopter blade flapping and shimmy bending moment decoupling and optical fiber sensor layout method.
Background
Paddle class machinery is the key equipment of aerospace trade, and in recent years, along with the deepening of research, paddle class mechanical properties constantly improves, constantly develops to high speed, heavy load, high efficiency, high reliability direction, consequently also more and more high to mechanical system's requirement. The blade is taken as a core component of the helicopter and is a main force bearing object, and most of damage accidents (cracks, breakage and the like) of the blade are caused by vibration, such as flapping motion, shimmy motion and the like of the blade of the helicopter. Therefore, monitoring the blade state of the helicopter is an important method for evaluating the quality of design, dynamically monitoring mechanical damage, diagnosing and predicting mechanical faults.
At present, researches at home and abroad aiming at a structural deformation state monitoring method mainly comprise a photoelectric measuring system and a non-photoelectric measuring system, wherein the photoelectric measuring system comprises a photogrammetry method, a laser tracker and the like; the non-photoelectric type includes a strain sensor, an acceleration sensor, a displacement sensor, an optical fiber sensor and the like. The photoelectric measuring system generally takes image capture and laser scanning tracking as the main parts, arranges a target point or a light source at a given position of a target structure, senses the target point or the light source in an imaging mode, obtains the spatial position of the target point through coordinate conversion according to a distance correction principle, a reflection principle and a photogrammetry principle, and has high response speed and high precision. However, the method is only suitable for a structure with a flat appearance and a simple form, has poor external interference resistance, and is easy to influence light transmission and difficult to ensure the reliability of a measuring system for a large-deformation flexible structure. The non-photoelectric measuring system mainly adopts a contact type measuring mode, is suitable for a complex structure on the basis of strain information or curvature information, can directly embed a sensor into a structural material, and has stable measuring performance. The contact type measurement mode is suitable for deformation reconstruction of a complex structure, and the measurement performance is stable.
Therefore, aiming at the characteristics of severe working environment and more external interference of the helicopter blade, a non-photoelectric contact type measuring method is selected and a relatively advanced and technically mature fiber Bragg grating sensor is selected as a testing element. Due to the asymmetry of the blade section, the coupling problem exists in the actual bending moment state monitoring process of waving bending moment and shimmy bending moment, and aiming at the defects of the conventional blade bending moment state monitoring method, a large amount of priori knowledge is not needed for research, so that the method can be applied to a conventional fiber grating demodulator with lower sampling frequency, and is a new method which is simple, rapid, convenient, reliable and high in practicability.
Disclosure of Invention
The purpose of the invention is as follows: in order to overcome the defects in the prior art, the invention provides a kirchhoff neutral layer principle-based helicopter blade flap and shimmy bending moment decoupling and optical fiber sensor layout method.
The technical scheme is as follows: in order to achieve the purpose, the invention adopts the technical scheme that:
a distributed optical fiber decoupling measurement method for blade flapping and shimmy bending moment of a helicopter comprises the following steps:
calculating the static moment of an elliptical part of a section of a helicopter blade according to the shape of the helicopter blade;
secondly, calculating the centroid position of the blade section according to the static moment theorem, determining a blade section main coordinate system according to the centroid position, and calculating the section inertia moment according to the determined coordinate system;
preliminarily selecting an approximate decoupling node position corresponding to the waving/shimmy bending moment according to a kirchhoff neutral layer hypothesis principle, and pasting a fiber grating sensor at the decoupling node position along the spanwise direction of the paddle, wherein the axial direction of the fiber grating sensor is consistent with the spanwise direction of the paddle;
the method comprises the steps that distributed fiber grating sensors are arranged on a blade structure, the root of the blade is fixed, a section is selected, three fiber grating sensors, namely a first fiber grating sensor FBG1, a second fiber grating sensor FBG2 and a third fiber grating sensor FBG3 are sequentially arranged on the surface of the blade passing through a main coordinate system of the section, according to the hypothesis principle of a Kirchf neutral layer, the stress of a structural neutral layer is zero, the transverse axis and the longitudinal axis of the main coordinate system of the section are approximately regarded as an initial neutral layer of the section, the intersection point of the coordinate axis and the section of the blade is a fiber grating sensor arrangement point, and a first fiber grating sensor FBG1, a second fiber grating sensor FBG2 and the third fiber grating sensor FBG are connected in series by adopting a fiber jumper to form a wavelength;
step four: carrying out flap shimmy loading on the helicopter blade, measuring a strain value of a section finite point of the helicopter blade, carrying out serialization on collected spanwise strain data of discrete points of the blade, solving spanwise strain distribution curve equations corresponding to different positions of the blade in the chordwise direction, and further inversing the strain value corresponding to any position of the blade in the chordwise direction;
the method comprises the steps of applying waving force and oscillating force to a blade, converting optical fiber reflection wavelength signals on three nodes into strain values according to the fiber grating strain sensing principle by collecting the optical fiber reflection wavelength signals, carrying out strain serialization on the strain values, solving a strain change equation, obtaining the strain value of any point of the section, wherein the relation between the strain and the grating wavelength is shown as a formula (6)
ΔλB/λB＝(1-Pe)Δ (6)
wherein, △ lambdaBIs the variation of the central wavelength of the fiber grating, lambdaBIs the center wavelength of the fiber grating, Peis the effective elasto-optical coefficient of the fiber grating, the delta is the variation of the strain value, and the relation between the strain value and the chord direction position is shown in the formula (7)
wherein, △ lambdaBIs the variation of the central wavelength of the fiber grating, lambdaBIs the center wavelength of the fiber grating, Peis the effective elasto-optical coefficient of the fiber grating, the delta is the variation of the strain value, and the relation between the strain value and the chord direction position is shown in the formula (7)
Wherein (gamma) is a strain value,ais the value of the initial strain in the chord direction,bis the strain value of the chord-wise end, gamma is the chord-wise position, gammaaIs the chord-wise initial position, and delta gamma is the chord-wise length difference;
step five: fitting strain values corresponding to the chord direction discrete positions of the blade measured in the fourth step, selecting a coordinate position corresponding to a zero strain value of the fitted curve to obtain a flapping and shimmy decoupling node, and correcting the position of the flapping and shimmy decoupling node to further determine the specific layout positions of the blade section fiber grating sensor, namely the flapping measuring node and the shimmy measuring node;
the helicopter blades are mainly subjected to centrifugal force F in a rotating statecAnd section bending moment MeEffects, based on force balance, can be:
Fc＝∫AσdA＝0 (8)
Me＝∫AzσdA (9)
wherein, sigma is stress, A is section area, Z is distance from infinitesimal surface to neutral axis, and coordinate relation Z of blade section coordinate system is establishedσ＝Z0+ΔZ；
When the blade is subjected to bending moment load, a part of the section reaches ultimate stress sigmamThe stress and Z can be obtained from the geometric relationshipσThe relationship is as follows:
combining the coordinate relationship, the formula (10) can be substituted into the formula (8)
Alternative variables are available
And then substituting the formulas (10) and (12) for the formula (9)
wherein l is the distance from the yield part of the section to the neutral axis, △ z is the offset of the neutral axis, MeIs a section bending moment, SkIs the static moment of the yield part of the blade section to the neutral axis of the section, StIs the static moment of the elastic part of the blade section to the neutral axis of the sectiontIs the moment of inertia of the elastic region of the blade section to the neutral axis of the section, AkIs the area of the yield part of the blade section, Atand (3) calculating the neutral axis offset △ z for the area of the elastic part of the section, correcting the position of the neutral axis of the blade, and finally determining the node position of the FBG sensor.
Preferably: the static moment of the elliptical section of the helicopter blade is as follows:
wherein S isyIs the transverse (y-direction) static moment, SzIs the longitudinal (z direction) static moment, bi，hiAre respectively the coefficients of a transverse elliptic formula and a longitudinal elliptic formula in a blade section coordinate system, thetaiThe method is characterized in that the method is the span of an ellipse relative to the origin of coordinates in a blade section coordinate system, theta is an angle, and n is the number of graphs divided by a section.
Preferably: in the second step, the coordinates y and z of the centroid of the cross section are respectively
Wherein 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, AiIs the area of each part of the cross section, yiIs the transverse coordinate of the centroid of the shape of the sections of the section, ziIs the longitudinal coordinate of the centroid of each part shape of the section,
after the position of the center is determined, a main coordinate system of the blade section is established by taking the center as an origin, and the inertia moment of the blade section is calculated according to the main coordinate system:
I＝∫Aρ2dA＝∫A(y2+z2)dA＝Iy+Iz(5)
wherein I is total moment of inertia of the cross section, A is the area of the cross section, rho is the distance from the micro area to the origin of coordinates, IyIs a transverse moment of inertia, IzAnd y and z are coordinate values of the blade infinitesimal section respectively.
Preferably: when the blade is bent, yielding and deformed, the neutral axis of the section is deviated, so that the deviation amount delta z needs to be determined, and the node position of the FBG sensor is corrected, so that the decoupling precision is improved.
Preferably: when the ratio of the flap strain to the lag vibration strain is smaller than the chord-span ratio of the helicopter blade, the node is considered to be capable of realizing the flap and lag vibration decoupling;
the first stage is as follows: exerting a flapping load on the helicopter blade, and respectively measuring the strain values of each node of the first FBG1, the second FBG2 and the third FBG3 of the fiber grating sensor1、2、3And applying a shimmy load to the blade to measure strain values of nodes of the first FBG1, the second FBG2 and the third FBG3 of the first FBG sensor, the second FBG sensor and the third FBG sensor respectively11、22、33；
And a second stage: setting the node of the two FBG2 nodes of the fiber bragg grating sensor as an estimated flapping decoupling point, comparing the ratio of the pendulum vibration strain to the flapping strain with the ratio of the chord length to the span length of the blade of the helicopter when the flapping load is not more than half of the pendulum vibration load, and if the flapping load is not more than half of the pendulum vibration load, comparing the chord length to the span length of the blade of the helicopter
The point is considered to meet the flapping decoupling requirement;
and a third stage: setting the nodes of the first FBG1 and the third FBG3 of the fiber grating sensor as estimated shimmy decoupling points, wherein the decoupling accuracy corresponding to the nodes of the first FBG1 and the third FBG3 of the fiber grating sensor needs to be compared in the case, and finally, the optimal shimmy decoupling point is confirmed.
Preferably: when the flapping load is not more than half of the shimmy load, comparing the flapping at the two nodesThe ratio of strain to shimmy strain is
Is smaller than the ratio of the chord length to the span length of the helicopter blade
The point is deemed to meet the shimmy decoupling requirement.
Compared with the prior art, the invention has the following beneficial effects:
the method is based on the kirchhoff neutral layer hypothesis principle, the key position of the decoupling node of the flapping/shimmy bending moment of the helicopter blade can be quickly and accurately determined, a large number of strain gauges do not need to be arranged on the surface of the blade, and the position of the decoupling node is determined step by step in an iterative manner by means of a large number of tests. The unique strain-direction sensitivity characteristic of the fiber bragg grating sensor is ingeniously utilized, the accurate measurement of the strain of the helicopter blades in different directions under the action of different forms of flap/shimmy bending moments can be realized, and the limitation of poor strain response direction sensitivity of the traditional resistance strain type sensor is overcome. In addition, the distributed optical fiber strain sensing network is adopted, a large number of signal transmission cables are not needed like a resistance strain sensing system, information transmission under the rotating working condition is facilitated, and negative effects on mechanical properties such as the natural frequency of the measured paddle structure are avoided.
Drawings
FIG. 1 is a layout diagram of a distributed fiber Bragg grating sensor;
FIG. 2 blade section coordinate system;
Detailed Description
The present invention is further illustrated by the following description in conjunction with the accompanying drawings and the specific embodiments, it is to be understood that these examples are given solely for the purpose of illustration and are not intended as a definition of the limits of the invention, since various equivalent modifications will occur to those skilled in the art upon reading the present invention and fall within the limits of the appended claims.
A distributed optical fiber decoupling measurement method for blade flapping and shimmy bending moment of a helicopter is a method for blade flapping and shimmy bending moment decoupling and optical fiber sensor layout of a helicopter based on the kirchhoff neutral layer principle, and comprises the following steps: the method comprises the following steps: deducing the static moment of the section of the helicopter blade; step two: calculating the centroid position of the section of the helicopter blade, determining a blade section main coordinate system according to the centroid position, and calculating the blade section inertia moment according to the determined coordinate system; step three: according to the hypothesis principle of kirchhoff neutral layer, roughly decoupling node positions corresponding to the flapping/shimmy bending moment are preliminarily selected. A fiber grating sensor is pasted at the position of the decoupling node along the spanwise direction of the blade, and the axial direction of the fiber grating sensor is consistent with the spanwise direction of the blade; step four: strain values of a plurality of discrete positions of a section are measured by applying flapping/shimmy load to the blade, collected spanwise strain data of discrete points of the blade are serialized, spanwise strain distribution curve equations corresponding to different positions of the blade in the chord direction can be obtained, and then strain values corresponding to any positions of the blade in the chord direction can be obtained through inversion calculation; step five: and fitting the strain values corresponding to the chord direction of the blade measured in the step four at a plurality of discrete positions, and selecting the coordinate position corresponding to the strain value of the fitted curve when the strain value is zero. And selecting the coordinate position as a decoupling node corresponding to the flapping/shimmy bending moment of the helicopter blade. The method is based on the kirchhoff neutral layer hypothesis principle, the key position of the decoupling node of the flapping/shimmy bending moment of the helicopter blade can be quickly and accurately determined, a large number of strain gauges do not need to be arranged on the surface of the blade, and the position of the decoupling node is determined step by step in an iterative manner by means of a large number of tests. The unique strain-direction sensitivity characteristic of the fiber bragg grating sensor is ingeniously utilized, the accurate measurement of the strain of the helicopter blades in different directions under the action of different forms of flap/shimmy bending moments can be realized, and the limitation of poor strain response direction sensitivity of the traditional resistance strain type sensor is overcome. In addition, the distributed optical fiber strain sensing network is adopted, a large number of signal transmission cables are not needed like a resistance strain sensing system, information transmission under the rotating working condition is facilitated, and negative effects on mechanical properties such as the natural frequency of the measured paddle structure are avoided.
The method specifically comprises the following steps:
step one, deducing the static moment of the section of a helicopter blade, and adopting a helicopter rotor blade scaling model made of a carbon fiber material; the typical section of a helicopter blade can be generally regarded as a combination of two-dimensional figures such as an ellipse, a triangle and the like. Wherein, the static moment of the elliptical part of the section of the helicopter blade is as follows:
wherein S isyIs the transverse (y-axis direction) static moment, SzIs the longitudinal (z-axis direction) static moment, bi，hiAre respectively the coefficient of an elliptic formula in a blade section coordinate system, thetaiThe span of the ellipse relative to the origin of coordinates in the blade section coordinate system.
Step two, calculating the centroid position of the blade section, determining a blade section main coordinate system according to the centroid position, and calculating section inertia moment according to the determined coordinate system;
according to the theorem of statics moment, the centroid coordinates y and z of the section are respectively
After the position of the center is determined, a main coordinate system of the blade section can be established by taking the center as an origin, and the inertia moment of the blade section is calculated according to the main coordinate system
I＝∫Aρ2dA＝∫A(y2+z2)dA＝Iy+Iz(5)
Namely, the section inertia moment is the sum of the inertia moments of all parts of the section. Wherein I is total moment of inertia of the cross section, IyIs the transverse (y-axis direction) moment of inertia, IzIs the longitudinal (z-axis direction) moment of inertia, and p is the micro-area to seatAnd the distance of a standard origin, A is the area of the section, and y and z are coordinate values of the blade infinitesimal section respectively.
And step three, according to the hypothesis principle of kirchhoff neutral layer, preliminarily selecting the approximate decoupling node position corresponding to the flapping/shimmy bending moment. A fiber grating sensor is pasted at the position of the decoupling node along the spanwise direction of the blade, and the axial direction of the fiber grating sensor is consistent with the spanwise direction of the blade;
as shown in fig. 1, distributed fiber grating sensors are arranged on a blade structure, the root of the blade is fixed, a section is selected, and three fiber grating sensors are sequentially arranged on the surface of the blade passing through a section main coordinate system. According to the hypothesis principle of the kirchhoff neutral layer, the stress of the structural neutral layer is zero, the x axis and the y axis of the main coordinate system of the section can be approximately regarded as the initial neutral layer of the section, the intersection point of the coordinate axis and the section of the blade is the arrangement point of the fiber bragg grating sensor, the FBGs 1, the FBG2 and the FBG3 are connected in series by adopting an optical fiber jumper to form a wavelength division multiplexing type sensing network, and the sampling frequency of a fiber bragg grating demodulator used by the sensing network is 1 KHz.
Step four: carrying out flapping, shimmy and loading on the helicopter blade, measuring a strain value of a section finite point of the helicopter blade, and carrying out serialization on collected spanwise strain data of discrete points of the blade so as to obtain spanwise strain distribution curve equations corresponding to different positions of the blade in the chordwise direction and further invert the strain value corresponding to any position of the blade in the chordwise direction;
the method comprises the steps of applying waving force and oscillating force to a blade, converting optical fiber reflection wavelength signals on three nodes into strain values according to the fiber grating strain sensing principle by collecting the optical fiber reflection wavelength signals, carrying out strain serialization on the strain values, solving a strain change equation, obtaining the strain value of any point of the section, wherein the relation between the strain and the grating wavelength is shown as a formula (6)
ΔλB/λB＝(1-Pe)Δ (6)
wherein, △ lambdaBIs the variation of the central wavelength of the fiber grating, lambdaBIs the center wavelength of the fiber grating, Pethe effective elasto-optical coefficient of the fiber grating, delta is the strain value, the relation between the strain value and the chord direction position is shown in the formula (7)
Step five: and fitting strain values corresponding to a plurality of discrete positions of the chord direction of the blade measured in the step four, and selecting a coordinate position corresponding to the strain value of the fitting curve when the strain value is zero to obtain a flapping oscillation decoupling node. And correcting the node positions, and further determining the specific layout positions of the blade section fiber grating sensors, namely a flapping measurement node and a shimmy measurement node.
When the blade is bent and deformed, the neutral axis of the section can deviate, so that the deviation amount delta z needs to be determined, and the node position of the FBG sensor is corrected, so that the decoupling precision is improved.
The helicopter blades are mainly subjected to centrifugal force F in a rotating statecAnd bending moment MeAction, obtainable from force balance
Fc＝∫AσdA＝0 (8)
Me＝∫AzσdA (9)
Where σ is the stress, A is the cross-sectional area, and z is the distance from the infinitesimal surface to the neutral axis. Establishing a blade section coordinate system as shown in FIG. 2, there is a coordinate relationship Zσ＝Z0+ΔZ。
When the blade is subjected to bending moment load, a part of the section reaches ultimate stress sigmamThe stress and Z can be obtained from the geometric relationshipσRelationships between
Combining the coordinate relationship, the formula (10) can be substituted into the formula (8)
Alternative variables are available
And then substituting the formulas (10) and (12) for the formula (9)
wherein l is the distance from the yield part of the section to the neutral axis, △ z is the offset of the neutral axis, MeIs a section bending moment, SkIs the static moment of the yield part of the blade section to the neutral axis of the section, StIs the static moment of the elastic part of the blade section to the neutral axis of the sectiontIs the moment of inertia of the elastic region of the blade section to the neutral axis of the section, AkIs the area of the yield part of the blade section, Atand calculating the neutral axis offset △ z, correcting the position of the neutral axis of the blade, and finally determining the node position of the FBG sensor.
When the blade is subjected to the shimmy force, if the ratio of the flap strain to the shimmy strain at the node is 0, the shimmy force is considered to be decoupled. However, this requirement is often difficult to achieve in practical engineering, and therefore, when the ratio of flap strain to edgewise strain is smaller than the chord-to-span ratio of the helicopter blade, it can be considered that flap and edgewise strain can be decoupled at this node.
The first stage is as follows: applying a flap load to the blade, and measuring the strain values of nodes of the FBGs 1, 2 and 3 as1、2、3. And applying a shimmy load to the blade, wherein the measured strain values of nodes of the FBGs 1, 2 and 3 are respectively11、22、33；
And a second stage: setting the node of the FBG2 as an estimated flapping decoupling point, comparing the ratio of the shimmy strain to the flapping strain with the ratio of the chord length to the span length of the helicopter blade when the flapping load is not more than half of the shimmy load, and if the shimmy load is not more than half of the shimmy load, comparing the ratio of the shimmy strain to the flapping strain
The point is considered to meet the flapping decoupling requirement;
and a third stage: setting the FBG1 and FBG3 nodes to be estimated shimmy solutionsAnd (3) coupling points, in this case, decoupling accuracies corresponding to nodes where the FBGs 1 and 3 are located respectively need to be compared, and finally, the optimal shimmy decoupling point is confirmed. When the flap load is not more than half of the pendulum vibration load, the ratio of the flap strain and the pendulum vibration strain at the two nodes is compared, namely
Is smaller than the ratio of the chord length to the span length of the helicopter blade
The point is deemed to meet the shimmy decoupling requirement.
The above description is only of the preferred embodiments of the present invention, and it should be noted that: it will be apparent to those skilled in the art that various modifications and adaptations can be made without departing from the principles of the invention and these are intended to be within the scope of the invention.
Claims (6)
1. A distributed optical fiber decoupling measurement method for blade flapping and shimmy bending moment of a helicopter is characterized by comprising the following steps:
calculating the static moment of an elliptical part of a section of a helicopter blade according to the shape of the helicopter blade;
secondly, calculating the centroid position of the blade section according to the static moment theorem, determining a blade section main coordinate system according to the centroid position, and calculating the section inertia moment according to the determined coordinate system;
preliminarily selecting an approximate decoupling node position corresponding to the waving/shimmy bending moment according to a kirchhoff neutral layer hypothesis principle, and pasting a fiber grating sensor at the decoupling node position along the spanwise direction of the paddle, wherein the axial direction of the fiber grating sensor is consistent with the spanwise direction of the paddle;
the method comprises the steps that distributed fiber grating sensors are arranged on a blade structure, the root of the blade is fixed, a section is selected, three fiber grating sensors, namely a first fiber grating sensor FBG1, a second fiber grating sensor FBG2 and a third fiber grating sensor FBG3 are sequentially arranged on the surface of the blade passing through a main coordinate system of the section, according to the hypothesis principle of a Kirchf neutral layer, the stress of a structural neutral layer is zero, the transverse axis and the longitudinal axis of the main coordinate system of the section are approximately regarded as an initial neutral layer of the section, the intersection point of the coordinate axis and the section of the blade is a fiber grating sensor arrangement point, and a first fiber grating sensor FBG1, a second fiber grating sensor FBG2 and the third fiber grating sensor FBG are connected in series by adopting a fiber jumper to form a wavelength;
step four: carrying out flap shimmy loading on the helicopter blade, measuring a strain value of a section finite point of the helicopter blade, carrying out serialization on collected spanwise strain data of discrete points of the blade, solving spanwise strain distribution curve equations corresponding to different positions of the blade in the chordwise direction, and further inversing the strain value corresponding to any position of the blade in the chordwise direction;
the blade is exerted and waved and the force of pendulum vibration, through collecting the fiber reflection wavelength signal on three nodes, according to fiber grating strain sensing principle with it transform into strain value, carry out the serialization that meets an emergency with strain value again, try to get strain change equation, obtain this section arbitrary point strain value, strain and grating wavelength relation are as shown in formula (6):
wherein,△λ B is the variation of the central wavelength of the fiber grating,λ B is the central wavelength of the optical fiber grating,P e is the effective elastic-optical coefficient of the fiber grating,△the relation between the strain value and the chord direction position is shown as the formula (7) as the variable quantity of the strain value:
wherein,
in order to obtain the value of the strain,
is the value of the initial strain in the chord direction,
is the value of the strain at the end in the chord direction,
the position of the string is the position of the string,
is the starting position of the chord direction,
is the chord length difference;
step five: fitting strain values corresponding to the chord direction discrete positions of the blade measured in the fourth step, selecting a coordinate position corresponding to a zero strain value of the fitted curve to obtain a flapping and shimmy decoupling node, and correcting the position of the flapping and shimmy decoupling node to further determine the specific layout positions of the blade section fiber grating sensor, namely the flapping measuring node and the shimmy measuring node;
the helicopter blades are mainly subjected to centrifugal force in a rotating stateF c And section bending momentM eEffects, based on force balance, can be:
wherein,σin order to be the stress,Ain terms of the cross-sectional area,zestablishing coordinate relation of blade section coordinate system for distance from infinitesimal surface to neutral axis
，
；
When the blade is subjected to bending moment load, a part of the section reaches ultimate stress
The stress and the stress can be obtained according to the geometric relationshipZ σ The relationship is as follows:
combining the coordinate relation, substituting the formula (10) into the formula (8) to obtain:
replacing the variables to get:
then substituting the formulas (10) and (12) into the formula (9) to obtain:
wherein,
the distance of the yield portion of the profile to the neutral axis,△zin order to be the neutral axis offset,M e in order to obtain a section bending moment,S k the static moment of the yield part of the blade section to the neutral axis of the section,S t the static moment of the elastic part of the blade section to the neutral axis of the section,I t is the inertia moment of the elastic area of the blade section to the neutral axis of the section,A k the area of the yield part of the blade section,A t calculating the neutral axis offset for the cross-sectional area of the elastic portion△zAnd then, correcting the position of the neutral axis of the blade, and finally determining the node position of the FBG sensor.
2. The distributed fiber optic decoupling measurement method of blade flapping and drag bending moment of a helicopter of claim 1, comprising: the static moment of the elliptical section of the helicopter blade is as follows:
wherein,S y in order to obtain the transverse static moment,S z in order to obtain a longitudinal static moment,b i ，h i are respectively the coefficients of a transverse elliptic formula and a longitudinal elliptic formula in a blade section coordinate system,θ i the span of the ellipse relative to the origin of coordinates in the blade section coordinate system,
in order to be an angle, the angle is,
the number of the patterns divided by the cross section,yis the position of the transverse coordinate of the centroid of the section,zis the position of the longitudinal coordinate of the centroid of the section.
3. The distributed fiber optic decoupling measurement method of blade flapping and drag bending moment of a helicopter of claim 2, wherein: step two, the section centroid coordinatesyAndzare respectively
Wherein,yis the position of the transverse coordinate of the centroid of the section,zis the position of the longitudinal coordinate of the centroid of the section,
in terms of the cross-sectional area,
in order to describe the area of each part of the cross section,
is the transverse coordinate of the centroid of each part shape of the section,
is the longitudinal coordinate of the centroid of each part shape of the section,
after the position of the center is determined, a main coordinate system of the blade section is established by taking the center as an origin, and the inertia moment of the blade section is calculated according to the main coordinate system:
wherein,Iis the total moment of inertia of the cross section,
in terms of the cross-sectional area,ρthe distance of the micro-area to the origin of coordinates,I y in order to obtain a transverse moment of inertia,I z is the longitudinal moment of inertia.
4. The distributed fiber optic decoupling measurement method of blade flapping and drag bending moment of a helicopter of claim 3, wherein: when the ratio of the flap strain to the lag vibration strain is smaller than the chord-span ratio of the helicopter blade, the node is considered to be capable of realizing the flap and lag vibration decoupling;
the first stage is as follows: exerting a flapping load on the blade, and respectively measuring the strain values of nodes of the first FBG1, the second FBG2 and the third FBG3 of the fiber grating sensor as 1 、 2 、 3 And applying a shimmy load to the blade to measure strain values of nodes of the first FBG1, the second FBG2 and the third FBG3 of the first FBG sensor, the second FBG sensor and the third FBG sensor respectively 11 、 22 、 33 ；
And a second stage: setting the node of the two FBG2 nodes of the fiber bragg grating sensor as an estimated flapping decoupling point, comparing the ratio of the shimmy strain to the flapping strain with the ratio of the chord length to the span length of the helicopter blade when the flapping load is not more than half of the shimmy load, and if the shimmy load is not more than the shimmy load, comparing the ratio of the shimmy strain to the flapping strain, and if the shimmy load is not more
If so, the point is considered to meet the flapping decoupling requirement;
and a third stage: setting the nodes of the first FBG1 and the third FBG3 of the fiber grating sensor as estimated shimmy decoupling points, wherein the decoupling accuracy corresponding to the nodes of the first FBG1 and the third FBG3 of the fiber grating sensor needs to be compared in the case, and finally, the optimal shimmy decoupling point is confirmed.
5. The distributed fiber optic decoupling measurement method of blade flapping and drag bending moment of a helicopter of claim 4, wherein: when the flap load is not more than half of the pendulum vibration load, the ratio of the flap strain and the pendulum vibration strain at the two nodes is compared, namely
Is smaller than the ratio of the chord length to the span length of the helicopter blade
And then the point is considered to meet the shimmy decoupling requirement.
6. The distributed fiber optic decoupling measurement method of blade flapping and drag bending moment of a helicopter of claim 5, wherein: the sampling frequency of the fiber grating demodulator used by the sensing network is 1 KHz.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910254397.XA CN110006563B (en) | 2019-03-31 | 2019-03-31 | Distributed optical fiber decoupling measurement method for blade flapping and shimmy bending moment of helicopter |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910254397.XA CN110006563B (en) | 2019-03-31 | 2019-03-31 | Distributed optical fiber decoupling measurement method for blade flapping and shimmy bending moment of helicopter |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN110006563A CN110006563A (en) | 2019-07-12 |
| CN110006563B true CN110006563B (en) | 2020-09-18 |
Family
ID=67169090
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201910254397.XA Active CN110006563B (en) | 2019-03-31 | 2019-03-31 | Distributed optical fiber decoupling measurement method for blade flapping and shimmy bending moment of helicopter |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN110006563B (en) |
Families Citing this family (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110595373A (en) * | 2019-09-06 | 2019-12-20 | 沈阳天眼智云信息科技有限公司 | Power transformer winding structure strain detection method based on optical fiber sensing |
| CN111289020B (en) * | 2020-03-05 | 2021-07-13 | 山东大学 | Sensing method for large deformation space pose of flexible body |
| CN112432746B (en) * | 2020-10-21 | 2022-11-11 | 中国人民解放军总参谋部第六十研究所 | Method for determining swing load coupling degree of helicopter blade |
| CN112525066A (en) * | 2020-11-20 | 2021-03-19 | 中国直升机设计研究所 | Helicopter blade surface strain gauge pasting and wire wiring method |
| CN113029405B (en) * | 2021-02-04 | 2022-04-22 | 南京航空航天大学 | Blade bending moment decoupling and calibrating method based on optical fiber strain combined bridge circuit |
| CN114166141B (en) * | 2021-11-23 | 2023-05-26 | 中国直升机设计研究所 | Multi-circular mark detection method for helicopter high-speed rotating blade image |
| CN115165175A (en) * | 2022-04-29 | 2022-10-11 | 南京航空航天大学 | Bending moment decoupling method in rotor load identification based on fiber bragg grating sensor |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2008131303A3 (en) * | 2007-04-20 | 2009-02-05 | Hansen Medical Inc | Optical fiber shape sensing systems |
| EP2351169A1 (en) * | 2008-10-24 | 2011-08-03 | Cymer, Inc. | System method and apparatus for selecting and controlling light source bandwidth |
| CN102426457A (en) * | 2011-11-20 | 2012-04-25 | 西北工业大学 | Flight control navigation system of miniature flapping-wing flying vehicle |
| CN103010485A (en) * | 2012-12-21 | 2013-04-03 | 南京航空航天大学 | Simulation modeling method for tilt-rotor unmanned plane and system thereof |
| CN104648688A (en) * | 2013-11-25 | 2015-05-27 | 中国直升机设计研究所 | Paddle strain gauge arranging and decoupling method |
| EP2915752A1 (en) * | 2012-11-05 | 2015-09-09 | Kabushiki Kaisha Towani | Aeroplane crushing device, and aeroplane demolition method |
| CN108120592A (en) * | 2017-11-29 | 2018-06-05 | 中国直升机设计研究所 | A kind of test method of helicopter blade static strength |
| CN108387183A (en) * | 2018-01-31 | 2018-08-10 | 天津大学 | Servo-actuated real-time measurement apparatus and its method for rotating vane whole audience dynamic deformation |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US10209275B2 (en) * | 2017-06-26 | 2019-02-19 | Corad Technology Inc. | Detachable probe card interface |
-
2019
- 2019-03-31 CN CN201910254397.XA patent/CN110006563B/en active Active
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2008131303A3 (en) * | 2007-04-20 | 2009-02-05 | Hansen Medical Inc | Optical fiber shape sensing systems |
| EP2351169A1 (en) * | 2008-10-24 | 2011-08-03 | Cymer, Inc. | System method and apparatus for selecting and controlling light source bandwidth |
| CN102426457A (en) * | 2011-11-20 | 2012-04-25 | 西北工业大学 | Flight control navigation system of miniature flapping-wing flying vehicle |
| EP2915752A1 (en) * | 2012-11-05 | 2015-09-09 | Kabushiki Kaisha Towani | Aeroplane crushing device, and aeroplane demolition method |
| CN103010485A (en) * | 2012-12-21 | 2013-04-03 | 南京航空航天大学 | Simulation modeling method for tilt-rotor unmanned plane and system thereof |
| CN104648688A (en) * | 2013-11-25 | 2015-05-27 | 中国直升机设计研究所 | Paddle strain gauge arranging and decoupling method |
| CN108120592A (en) * | 2017-11-29 | 2018-06-05 | 中国直升机设计研究所 | A kind of test method of helicopter blade static strength |
| CN108387183A (en) * | 2018-01-31 | 2018-08-10 | 天津大学 | Servo-actuated real-time measurement apparatus and its method for rotating vane whole audience dynamic deformation |
Non-Patent Citations (2)
| Title |
|---|
| Junsong Yu.Short-time Energy Characterization of Low Velocity Impact Localization Method on CFRP Using FBG Sensors.《The 6th Asia Pacific Optical Sensors Conference 2016 ©》.2016, * |
| 一种高效率的桨叶载荷标定方法;余询;《直升机技术》;20011231(第2期);9-11 * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN110006563A (en) | 2019-07-12 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN110006563B (en) | Distributed optical fiber decoupling measurement method for blade flapping and shimmy bending moment of helicopter | |
| WO2020192623A2 (en) | Rotating blade non-contact dynamic strain field measurement method and system | |
| WO2020192621A1 (en) | Measurement method for rotor blade dynamic strain field based on blade tip timing and system thereof | |
| CN110127078B (en) | Optical fiber monitoring method for strain-deflection-bending moment state of helicopter blade structure | |
| CN109883380B (en) | Rotor blade displacement field measuring method and system based on blade end timing | |
| CN105136359B (en) | Method based on the beam surface strain values evaluation work load that fibre optical sensor measures | |
| CN109766617B (en) | Displacement field reconstruction method based on strain sensor | |
| CN114754691B (en) | Distributed optical fiber monitoring and inversion method for helicopter blade bending form | |
| CN113029405B (en) | Blade bending moment decoupling and calibrating method based on optical fiber strain combined bridge circuit | |
| JPH1090087A (en) | Estimation method for load acting on structure | |
| CN108225720A (en) | Optical fiber aerodynamics force measurement balance and fiber optic strain gage installation method | |
| CN107505477A (en) | A kind of three-dimensional fiber Bragg grating wind speed wind direction sensor and system | |
| CN203688576U (en) | Fiber Bragg grating dynamic anemoclinograph | |
| CN105758602A (en) | Truss girder bridge section buffeting force synchronous measurement method | |
| CN111649907A (en) | Rib strip type micro-axial force annular two-component strain balance | |
| CN207675408U (en) | Optical fiber aerodynamics force measurement balance applied to hypersonic low density wind tunnel | |
| CN113970298B (en) | FBG-based overall deformation detection method for large-scale aerospace flexible structure | |
| CN106525277A (en) | Unstressed encapsulated fiber grating temperature sensing string | |
| CN206311247U (en) | A kind of sensor device of power and displacement measurement based on distributed fibre optic sensing | |
| CN107701379B (en) | Sensor calibration method and device for wind generating set blade | |
| CN114199152A (en) | Wing shape variation measuring method and device | |
| CN108267264A (en) | Method, device and equipment for calibrating fiber grating sensor of fan blade | |
| CN106525301A (en) | Force and displacement measurement method and sensor based on distributed optical fiber sensing | |
| CN110823120B (en) | Compensation method for measurement error of surface-mounted fiber grating strain sensor | |
| WO2024073899A1 (en) | Inhaul cable tension identification algorithm considering semi-rigid constraints at two ends |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |