CN112685937B - Airplane structure health monitoring method based on fiber grating sensor - Google Patents
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Abstract
The application belongs to the field of comprehensive strength design, and particularly relates to an aircraft structure health monitoring method based on a fiber grating sensor. The method comprises the following steps: determining a key part of an airplane structure, and installing a fiber bragg grating sensor at the key part of the airplane structure; determining the sampling rate of the fiber bragg grating sensor, acquiring strain history data of key parts of an airplane structure in the flight process through the fiber bragg grating sensor, and preprocessing the strain history data; performing rain flow counting on the preprocessed strain history data, and calculating equivalent damage of a key part of the aircraft structure of the current flight landing and landing according to an equivalent damage calculation formula; acquiring a test piece of a key part of an airplane structure, mounting a fiber grating sensor on the test piece of the key part of the airplane structure according to actual conditions, developing a fatigue test by adopting a full-airplane design load spectrum, and determining the reference damage of the key part of the airplane structure; and calculating the service life consumption and the residual service life according to the equivalent damage and the reference damage of the key part of the aircraft structure.
Description
Technical Field
The application belongs to the field of comprehensive strength design, and particularly relates to an aircraft structure health monitoring method based on a fiber grating sensor.
Background
With the continuous increase of the manufacturing cost of the new generation of airplanes and the stricter requirements of the users on the safety of the airplanes, the traditional regular maintenance cannot meet the requirements of economic repair and high safety and reliability, so the situation-based maintenance becomes the development trend of the maintenance of the new generation of airplanes, and the accurate assessment of the health condition of the airplane structure is the basis for realizing the situation-based maintenance.
At present, military aircrafts in China generally adopt a parameter type life monitoring method to carry out state evaluation on the whole aircraft, namely, the normal overload of the gravity center of the aircraft is taken as a calculation parameter, the integral equivalent damage of each landing aircraft is calculated, and the prediction of the life consumption of the aircraft is realized. However, the method is used for evaluating the overall health state of the airplane, the difference of key parts of each structure is not considered, and meanwhile, for the horizontal tail, the vertical tail and related structures, the linear relation does not exist between the structure loading and the gravity center normal overload, and the existing method cannot evaluate the health state of the structures. In addition, the traditional strain sensor has poor reliability and maintainability, and the weight of the machine body is increased due to the installation on the machine, so that the structural health monitoring requirement cannot be met.
Accordingly, a technical solution is desired to overcome or at least alleviate at least one of the above-mentioned drawbacks of the prior art.
Disclosure of Invention
The application aims to provide an aircraft structure health monitoring method based on a fiber grating sensor, so as to solve at least one problem existing in the prior art.
The technical scheme of the application is as follows:
an aircraft structure health monitoring method based on a fiber grating sensor comprises the following steps:
determining a key part of an airplane structure according to whole-airplane finite element analysis and engineering experience, and mounting a fiber bragg grating sensor on the key part of the airplane structure;
secondly, determining the sampling rate of the fiber grating sensor, acquiring strain history data of key parts of the aircraft structure in the flight process through the fiber grating sensor, and preprocessing the strain history data;
thirdly, performing rain flow counting on the preprocessed strain history data, and calculating equivalent damage of a key part of the airplane structure in current flight and landing according to an equivalent damage calculation formula;
step four, obtaining a test piece of the key part of the airplane structure, installing a fiber grating sensor on the test piece of the key part of the airplane structure according to actual conditions, developing a fatigue test by adopting a full-airplane design load spectrum, and determining the reference damage of the key part of the airplane structure;
and fifthly, calculating the life consumption of the key part of the aircraft structure according to the equivalent damage and the reference damage of the key part of the aircraft structure, and calculating the residual life of the key part of the aircraft structure according to the life consumption.
Optionally, in the first step, the aircraft structure key part includes: the butt-joint area of the middle fuselage and the outer wing, the butt-joint area of the rear fuselage and the vertical tail, the butt-joint area of the horizontal tail and the fuselage, and the butt-joint area of the outer wing and the middle fuselage.
Optionally, in step two, the sampling rate of the fiber grating sensor is set to 32 times/second.
Optionally, in the second step, the preprocessing of the strain history data includes artifact removal and filtering.
Optionally, in step three, performing rain flow counting on the preprocessed strain history data, and calculating the equivalent damage of the key part of the aircraft structure in the current takeoff and landing according to an equivalent damage calculation formula includes:
performing rain flow counting on the preprocessed strain history data to obtain peak-valley pair data (epsilon) max,i ，ε min,i ) And converting damage areas such as each cycle into a pulsating cycle by adopting an Otto conversion formula:
the damage of the key part of the aircraft structure corresponding to the equivalent strain is as follows:
D i ＝(ε equivalent weight, i ) m ×N i
Wherein, N i Is epsilon Equivalent weight, i M is the index of an S-N curve, and m is taken as 4 during calculation;
accumulating all full-cycle and half-cycle damages of the strain history of the key part of the aircraft structure in the current flight, landing and landing to obtain the equivalent damage of the key part of the aircraft structure in the jth flight, landing and landing:
D critical site, j ＝∑(ε Equivalent weight, i ) m ×N i 。
Optionally, in step four, the reference damage of the aircraft structure key part is:
wherein D is Spectrum assembly For total damage of the test piece, n is the number of spectral blocks performed when the test piece is damaged, T Spectrum block The corresponding time of flight for each spectral block.
Optionally, in step five, the calculating the life consumption of the aircraft structure key part according to the equivalent damage and the reference damage of the aircraft structure key part, and the calculating the remaining life of the aircraft structure key part according to the life consumption includes:
the life consumption of the key parts of the aircraft structure when the aircraft is taking off and landing in the next flight is as follows:
accumulating the past life consumption data of the key part of the aircraft structure, and calculating the remaining life of the key part of the aircraft structure by combining with a life index as follows:
wherein, T Total life time The total life of the critical parts of the aircraft structure.
The invention has at least the following beneficial technical effects:
the aircraft structure health monitoring method based on the fiber bragg grating sensor can accurately evaluate the damage and service life consumption conditions of the key parts of each aircraft structure in the flight process, can be used as data support for aircraft maintenance according to the situation, and is accurate in calculation.
Drawings
FIG. 1 is a flow chart of a method for monitoring the health of an aircraft structure based on fiber grating sensors according to an embodiment of the present application;
FIG. 2 is a schematic view of a fiber grating sensor installation according to an embodiment of the present application;
FIG. 3 is a schematic illustration of a test piece of a critical part of an aircraft structure according to an embodiment of the present application.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the drawings in the embodiments of the present application. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are a subset of the embodiments in the present application and not all embodiments in the present application. The embodiments described below with reference to the accompanying drawings are illustrative and intended to explain the present application and should not be construed as limiting the present application. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application. Embodiments of the present application will be described in detail below with reference to the accompanying drawings.
In the description of the present application, it is to be understood that the terms "center", "longitudinal", "lateral", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on those shown in the drawings, and are used merely for convenience in describing the present application and for simplifying the description, and do not indicate or imply that the referenced device or element must have a particular orientation, be constructed in a particular orientation, and be operated, and therefore should not be construed as limiting the scope of the present application.
The present application is described in further detail below with reference to fig. 1 to 3.
The application provides an aircraft structure health monitoring method based on a fiber grating sensor, which comprises the following steps:
determining a key part of an airplane structure according to the finite element analysis of a whole airplane and engineering experience, and installing a fiber grating sensor at the key part of the airplane structure;
secondly, determining the sampling rate of the fiber bragg grating sensor, acquiring strain history data of key parts of the aircraft structure in the flight process through the fiber bragg grating sensor, and preprocessing the strain history data;
thirdly, performing rain flow counting on the preprocessed strain history data, and calculating equivalent damage of a key part of the airplane structure of the current flight landing and landing according to an equivalent damage calculation formula;
acquiring a test piece of the key part of the airplane structure, mounting a fiber grating sensor on the test piece of the key part of the airplane structure according to actual conditions, carrying out fatigue test by adopting a full-airplane design load spectrum, and determining the reference damage of the key part of the airplane structure;
and fifthly, calculating the life consumption of the key part of the aircraft structure according to the equivalent damage and the reference damage of the key part of the aircraft structure, and calculating the residual life of the key part of the aircraft structure according to the life consumption.
In one embodiment of the present application, key parts of an aircraft structure are planned according to the whole aircraft finite element analysis result and fatigue analysis work of a certain type of aircraft, and engineering experience is combined, and a fiber grating sensor is installed at the corresponding key parts of the aircraft structure, see table 1.
TABLE 1
In this embodiment, the fiber grating sensor is installed along the stress direction according to the stress form of these key parts of the aircraft structure, as shown in fig. 2. Because the whole coverage of the sensor network cannot be realized due to the limiting factors such as the space on the machine, the data channel and the like, the key parts of the structure need to be determined according to the finite element analysis of the whole machine and the engineering experience, and the fiber grating sensor needs to be installed. When the fiber grating sensor is installed, attention needs to be paid to: the routing path of the sensor is ensured to be direct, and unified planning is performed according to the wiring harness, so that the wiring harness is prevented from being arranged on a movable part, and the wiring harness is prevented from being worn; the number of channels does not exceed the number of plans; the sensor should be installed in the structure high stress district, and simultaneously under the prerequisite that does not influence strain measurement, the sensor mounted position should leave operating space, the maintenance of being convenient for.
According to the aircraft structure health monitoring method based on the fiber grating sensor, in the second step, the sampling rate of the fiber grating sensor needs to be determined, the sampling rate is reduced as far as possible on the basis of ensuring the data integrity, the influence of high-frequency vibration on the measurement result is effectively avoided, and meanwhile, the storage space is saved. In the embodiment, the sampling rate is determined to be 32 times/second according to the analysis of the flight parameter data of the past airplane, so that the requirement on data integrity can be met, and meanwhile, the generated data volume does not influence storage. In addition, after strain history data of each aircraft structure key part in the flight process is collected through the fiber bragg grating sensor, preprocessing such as false removal and filtering needs to be carried out on the collected data, and data with no influence on the structural health state and error data are eliminated. Wherein, the filtering threshold value is set to be (20MPa/E) mu epsilon, and E is the elastic modulus of the selected material of the corresponding key part of the airplane structure.
The utility model provides an aircraft structure health monitoring method based on fiber grating sensor, in the third step, carry out rain flow meter to the strain history data after the preliminary treatment, calculate the equivalent damage of the aircraft structure key part of taking off and landing at present according to equivalent damage computational formula and include:
performing rain flow counting on the preprocessed strain history data to obtain peak-valley pair data (epsilon) max,i ，ε min,i ) And converting damage areas such as each cycle into a pulsating cycle by adopting an Otto conversion formula:
the damage of the key part of the airplane structure corresponding to the equivalent strain is as follows:
D i ＝(ε equivalent of i ) m ×N i
Wherein N is i Is epsilon Equivalent weight, i M is an index of an S-N curve, and m is taken to be 4 in calculation；
Accumulating all full-cycle and half-cycle damages of the strain histories of the key parts of the aircraft structure during the current time of flight, landing and taking off to obtain the equivalent damage of the key parts of the aircraft structure during the jth time of flight, landing and taking off:
D critical site, j ＝∑(ε Equivalent weight, i ) m ×N i 。
The aircraft structure health monitoring method based on the fiber bragg grating sensor comprises the fourth step of designing an element test piece and a box section test piece according to the actual structural condition of a sensor network, as shown in fig. 3, carrying out fatigue test by adopting a full-machine design load spectrum, carrying out fake removal, filtering and rain flow counting on strain history data acquired from the beginning to the end of the test, converting equivalent strain by adopting an Otton transformation formula, and calculating the total damage D of the test piece according to the calculation process of the third step Spectrum total And obtaining the reference damage of the corresponding key part of the airplane structure:
wherein D is Spectrum assembly For total damage of the test piece, n is the number of spectral blocks performed when the test piece is damaged, T Spectrum block The corresponding time of flight for each spectral block.
Finally, in the fifth step, calculating the life consumption of the key part of the aircraft structure according to the equivalent damage and the reference damage of the key part of the aircraft structure, and calculating the residual life of the key part of the aircraft structure according to the life consumption comprises:
the life consumption of the key parts of the aircraft structure when the aircraft flies and lands for the second time is as follows:
accumulating the past life consumption data of the key part of the aircraft structure, and calculating the remaining life of the key part of the aircraft structure by combining with a life index as follows:
wherein, T Total life time The total life of the critical parts of the aircraft structure.
In one embodiment of the present application, taking the key parts of the structure of the midbody as an example, the equivalent damage of the key parts of the midbody of the j-th flight takeoff and landing is obtained through the third step:
D middle fuselage j ＝∑(ε Equivalent weight, i ) m ×N i
Obtaining the benchmark damage of the key part of the middle fuselage according to the fourth step:
and finally, calculating to obtain the service life consumption of the key part of the middle fuselage during the rise and fall through the fifth step:
accumulating the past life consumption data of the part, and calculating the remaining life of the part by combining the life indexes:
according to the method, the residual life of the key part of each structure can be accurately grasped, and whether the health state meets the flight requirement or not and whether maintenance is needed or not is determined by combining the importance of the structure.
According to the aircraft structure health monitoring method based on the fiber bragg grating sensors, the key parts of the aircraft structure are determined, the fiber bragg grating sensors which are light in weight and high in reliability and can realize multipoint measurement are adopted to form a structure state information acquisition network, and strain history data of the key parts of the structure in the flight process are acquired; in addition, an element test piece and a box section test piece are designed by combining the actual structure of the airplane to carry out fatigue test, and a mapping relation model of the strain history and the service life consumption of each structure key part is established; the strain history data of the key parts collected by the fiber bragg grating sensor are used as input, the life consumption and the residual life are calculated, the health state evaluation of the key parts of each structure is realized, and data support is provided for the maintenance according to the situation. The method and the device can be used for carrying out health monitoring on the specific aircraft structure key parts, and the calculation result is more accurate than that of the traditional single-machine service life monitoring method; meanwhile, the method is not limited to airplane structures, is also suitable for other mechanical structures, and has a great engineering application value and a wide application prospect.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (4)
1. An aircraft structure health monitoring method based on a fiber grating sensor is characterized by comprising the following steps:
determining a key part of an airplane structure according to finite element analysis of a whole airplane and engineering experience, and mounting a fiber grating sensor on the key part of the airplane structure;
secondly, determining the sampling rate of the fiber bragg grating sensor, acquiring strain history data of key parts of the aircraft structure in the flight process through the fiber bragg grating sensor, and preprocessing the strain history data;
thirdly, performing rain flow counting on the preprocessed strain history data, and calculating the equivalent damage of the key part of the aircraft structure of the current flight landing and landing according to an equivalent damage calculation formula, wherein the method comprises the following steps:
performing rain flow counting on the preprocessed strain history data to obtain peak-valley pair data (epsilon) max,i ，ε min,i ) Using the formula of the transformation of OttoThe damage such as each cycle is converted into a pulsating cycle:
the damage of the key part of the aircraft structure corresponding to the equivalent strain is as follows:
D i ＝(ε equivalent weight, i ) m ×N i
Wherein N is i Is epsilon Equivalent weight, i M is the index of an S-N curve, and m is taken as 4 during calculation;
accumulating all full-cycle and half-cycle damages of the strain histories of the key parts of the aircraft structure during the current time of flight, landing and taking off to obtain the equivalent damage of the key parts of the aircraft structure during the jth time of flight, landing and taking off:
D critical site, j ＝∑(ε Equivalent weight, i ) m ×N i ；
Step four, obtaining a test piece of the key part of the aircraft structure, installing a fiber grating sensor on the test piece of the key part of the aircraft structure according to actual conditions, carrying out fatigue tests by adopting a full-aircraft design load spectrum, and determining the reference damage of the key part of the aircraft structure, wherein the reference damage of the key part of the aircraft structure is as follows:
wherein D is Spectrum assembly For total damage of the test piece, n is the number of spectral blocks performed when the test piece is damaged, T Spectrum block A corresponding time of flight for each spectral block;
step five, calculating the life consumption of the key part of the aircraft structure according to the equivalent damage and the reference damage of the key part of the aircraft structure, and calculating the residual life of the key part of the aircraft structure according to the life consumption, wherein the method comprises the following steps:
the life consumption of the key parts of the aircraft structure when the aircraft flies and lands for the second time is as follows:
accumulating the past life consumption data of the key part of the aircraft structure, and calculating the remaining life of the key part of the aircraft structure by combining with a life index as follows:
wherein, T Total life time Is the total life of the critical parts of the aircraft structure.
2. The fiber bragg grating sensor-based aircraft structure health monitoring method as claimed in claim 1, wherein in the first step, the key parts of the aircraft structure comprise: the butt-joint area of the middle fuselage and the outer wing, the butt-joint area of the rear fuselage and the vertical tail, the butt-joint area of the horizontal tail and the fuselage, and the butt-joint area of the outer wing and the middle fuselage.
3. The fiber grating sensor-based aircraft structure health monitoring method as claimed in claim 1, wherein in the second step, the sampling rate of the fiber grating sensor is set to 32 times/second.
4. The fiber grating sensor-based aircraft structure health monitoring method as claimed in claim 3, wherein in the second step, the preprocessing of the strain history data includes artifact removal and filtering.
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