CN114813003A - Multi-parameter measurement method for vibration fatigue damage of metal component of airplane - Google Patents
Multi-parameter measurement method for vibration fatigue damage of metal component of airplane Download PDFInfo
- Publication number
- CN114813003A CN114813003A CN202210733242.6A CN202210733242A CN114813003A CN 114813003 A CN114813003 A CN 114813003A CN 202210733242 A CN202210733242 A CN 202210733242A CN 114813003 A CN114813003 A CN 114813003A
- Authority
- CN
- China
- Prior art keywords
- metal component
- airplane
- damage
- trend
- vibration
- 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.)
- Pending
Links
- 239000002184 metal Substances 0.000 title claims abstract description 78
- 238000000691 measurement method Methods 0.000 title claims abstract description 11
- 230000004044 response Effects 0.000 claims abstract description 43
- 238000006073 displacement reaction Methods 0.000 claims abstract description 22
- 238000000034 method Methods 0.000 claims abstract description 22
- 238000009661 fatigue test Methods 0.000 claims abstract description 19
- 230000008859 change Effects 0.000 claims abstract description 14
- 238000012360 testing method Methods 0.000 claims description 13
- 230000003247 decreasing effect Effects 0.000 claims description 11
- 238000001931 thermography Methods 0.000 claims description 9
- 230000000977 initiatory effect Effects 0.000 claims description 5
- 238000012545 processing Methods 0.000 claims description 5
- 230000002159 abnormal effect Effects 0.000 claims description 3
- 230000001174 ascending effect Effects 0.000 claims description 3
- 238000001514 detection method Methods 0.000 claims description 3
- 238000001914 filtration Methods 0.000 claims description 3
- 230000000630 rising effect Effects 0.000 claims 1
- 238000005259 measurement Methods 0.000 abstract description 9
- 230000008569 process Effects 0.000 description 5
- 230000007547 defect Effects 0.000 description 4
- 238000010586 diagram Methods 0.000 description 3
- 230000035945 sensitivity Effects 0.000 description 3
- 230000001133 acceleration Effects 0.000 description 2
- 230000007613 environmental effect Effects 0.000 description 2
- 230000005284 excitation Effects 0.000 description 2
- 230000015572 biosynthetic process Effects 0.000 description 1
- 238000007689 inspection Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000011160 research Methods 0.000 description 1
Images
Classifications
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M7/00—Vibration-testing of structures; Shock-testing of structures
- G01M7/02—Vibration-testing by means of a shake table
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
- B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
- B64F5/60—Testing or inspecting aircraft components or systems
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01H—MEASUREMENT OF MECHANICAL VIBRATIONS OR ULTRASONIC, SONIC OR INFRASONIC WAVES
- G01H9/00—Measuring mechanical vibrations or ultrasonic, sonic or infrasonic waves by using radiation-sensitive means, e.g. optical means
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01J—MEASUREMENT OF INTENSITY, VELOCITY, SPECTRAL CONTENT, POLARISATION, PHASE OR PULSE CHARACTERISTICS OF INFRARED, VISIBLE OR ULTRAVIOLET LIGHT; COLORIMETRY; RADIATION PYROMETRY
- G01J5/00—Radiation pyrometry, e.g. infrared or optical thermometry
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M7/00—Vibration-testing of structures; Shock-testing of structures
- G01M7/02—Vibration-testing by means of a shake table
- G01M7/025—Measuring arrangements
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N29/00—Investigating or analysing materials by the use of ultrasonic, sonic or infrasonic waves; Visualisation of the interior of objects by transmitting ultrasonic or sonic waves through the object
- G01N29/14—Investigating or analysing materials by the use of ultrasonic, sonic or infrasonic waves; Visualisation of the interior of objects by transmitting ultrasonic or sonic waves through the object using acoustic emission techniques
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N3/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N3/02—Details
- G01N3/06—Special adaptations of indicating or recording means
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N3/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N3/02—Details
- G01N3/06—Special adaptations of indicating or recording means
- G01N3/068—Special adaptations of indicating or recording means with optical indicating or recording means
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N3/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N3/32—Investigating strength properties of solid materials by application of mechanical stress by applying repeated or pulsating forces
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01J—MEASUREMENT OF INTENSITY, VELOCITY, SPECTRAL CONTENT, POLARISATION, PHASE OR PULSE CHARACTERISTICS OF INFRARED, VISIBLE OR ULTRAVIOLET LIGHT; COLORIMETRY; RADIATION PYROMETRY
- G01J5/00—Radiation pyrometry, e.g. infrared or optical thermometry
- G01J2005/0077—Imaging
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N2203/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N2203/0001—Type of application of the stress
- G01N2203/0005—Repeated or cyclic
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N2203/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N2203/0058—Kind of property studied
- G01N2203/006—Crack, flaws, fracture or rupture
- G01N2203/0062—Crack or flaws
- G01N2203/0064—Initiation of crack
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N2203/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N2203/0058—Kind of property studied
- G01N2203/006—Crack, flaws, fracture or rupture
- G01N2203/0062—Crack or flaws
- G01N2203/0066—Propagation of crack
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N2203/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N2203/02—Details not specific for a particular testing method
- G01N2203/06—Indicating or recording means; Sensing means
- G01N2203/0641—Indicating or recording means; Sensing means using optical, X-ray, ultraviolet, infrared or similar detectors
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N2203/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N2203/02—Details not specific for a particular testing method
- G01N2203/06—Indicating or recording means; Sensing means
- G01N2203/0658—Indicating or recording means; Sensing means using acoustic or ultrasonic detectors
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N2203/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N2203/02—Details not specific for a particular testing method
- G01N2203/06—Indicating or recording means; Sensing means
- G01N2203/067—Parameter measured for estimating the property
- G01N2203/0682—Spatial dimension, e.g. length, area, angle
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N2203/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N2203/02—Details not specific for a particular testing method
- G01N2203/06—Indicating or recording means; Sensing means
- G01N2203/067—Parameter measured for estimating the property
- G01N2203/0694—Temperature
Landscapes
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- General Health & Medical Sciences (AREA)
- Immunology (AREA)
- Life Sciences & Earth Sciences (AREA)
- Chemical & Material Sciences (AREA)
- Analytical Chemistry (AREA)
- Biochemistry (AREA)
- Pathology (AREA)
- Health & Medical Sciences (AREA)
- Engineering & Computer Science (AREA)
- Spectroscopy & Molecular Physics (AREA)
- Manufacturing & Machinery (AREA)
- Transportation (AREA)
- Aviation & Aerospace Engineering (AREA)
- Acoustics & Sound (AREA)
- Investigating Or Analyzing Materials Using Thermal Means (AREA)
Abstract
The invention discloses a multi-parameter measurement method for vibration fatigue damage of an airplane metal component, which comprises the following steps: firstly, building a vibration fatigue test system; secondly, starting a vibration fatigue test system; measuring a movement displacement response signal of the metal component of the airplane, and marking a damage early warning signal of the metal component of the airplane; measuring an energy response signal of the airplane metal component, and marking the crack state in the airplane metal component; and fifthly, measuring the surface temperature value of the metal component of the airplane when the trend of the acoustic emission signal changes, and measuring the damage position. The method utilizes the change of the change trend of the motion displacement response signal as the early warning signal of the damage of the metal component of the airplane to observe whether the energy response trend and the counting trend are suddenly increased or not, and searches the maximum moment of the temperature standard deviation value by combining the change conditions of the surface temperature distribution and the temperature standard deviation value of the metal component of the airplane to comprehensively judge the occurrence and the position of the damage, thereby realizing the accurate and rapid multi-parameter measurement of the damage of the metal component of the airplane.
Description
Technical Field
The invention belongs to the technical field of vibration fatigue damage measurement of metal components of airplanes, and particularly relates to a vibration fatigue damage multi-parameter measurement method of metal components of airplanes.
Background
Structural vibration fatigue is a phenomenon in which a structure is fatigue-damaged by excitation such as vibration or noise. A significant difference from the conventional fatigue problem is that the excitation force is constantly changing and can be influenced by the structure dynamics. The vibration fatigue test is widely applied to the fields of aerospace, rail transit and the like as an important mode for verifying the dynamic strength of a structure, and how to accurately and quickly measure damage information in the vibration fatigue test process directly influences test criteria is a key research problem in the field of the vibration fatigue test.
At present, the mainstream method for measuring the vibration fatigue damage of the metal component of the airplane is to monitor a motion signal through an acceleration sensor, a laser displacement sensor or a laser vibration meter and the like, analyze the stress level of a test piece through a strain gauge, and judge and acquire damage information by integrating the information of the acceleration sensor, the laser displacement sensor or the laser vibration meter and the like. Although the method can find the change of the signal after the test piece is damaged, the method has two defects that the sensitivity is not high, and the damage information can be measured only when the state of the test piece is obviously changed; and secondly, hysteresis exists, the state of the test piece is required to be changed and then the test piece is continuously checked for a period of time to determine whether damage or interference of other factors occurs, and if the test piece is subjected to a condition which cannot be judged, shutdown inspection is required, so that damage formation and an evolution process of the test piece are inevitably influenced.
Disclosure of Invention
The invention aims to solve the technical problem of providing a multi-parameter measurement method for the vibration fatigue damage of the metal component of the airplane aiming at the defects in the prior art, which utilizes the change trend of a motion displacement response signal as an early warning signal for the damage of the metal component of the airplane, observes whether the energy response trend and the counting trend of the acoustic emission signal are suddenly increased, and the maximum time of the temperature standard deviation value is searched by combining the change conditions of the surface temperature distribution and the temperature standard deviation value of the metal component of the airplane, whether damage occurs or not is comprehensively judged, the damage position is determined, the multi-parameter accurate and rapid measurement of the damage of the metal component of the airplane is realized, a reliable basis is provided for the subsequent evolution process analysis of the damage of the metal component of the airplane, the vibration fatigue test is not required to be interrupted, the measurement of the damage is obtained through the combination of various parameters, and the method is convenient to popularize and use.
In order to solve the technical problems, the invention adopts the technical scheme that: a multi-parameter measurement method for vibration fatigue damage of an airplane metal component is characterized by comprising the following steps:
the method comprises the following steps of firstly, building a vibration fatigue test system, wherein the vibration fatigue test system comprises a vibration table for mounting an airplane metal component, a thermal infrared camera and a laser transmitter, the thermal infrared camera and the laser transmitter are arranged beside the vibration table and are used for collecting temperature signals and motion displacement response signals respectively, and an acoustic emission sensor is mounted on the airplane metal component;
the signal output end of the acoustic emission sensor is connected with an acoustic emission signal collector, the signal output end of the thermal infrared camera is connected with a thermal imaging analyzer, the signal output end of the laser transmitter is connected with a laser vibration meter, and the output ends of the acoustic emission signal collector, the thermal imaging analyzer and the laser vibration meter are all connected with a computer;
step two, starting a vibration fatigue test system;
measuring a movement displacement response signal of the airplane metal component by using a laser transmitter, processing the movement displacement response signal by using a laser vibration meter, transmitting the movement displacement response signal to a computer, continuously increasing the obtained movement displacement response value, and when the obtained movement displacement response value is changed from increasing to decreasing, indicating that the rigidity of the structure of the airplane metal component is decreased, the natural frequency is decreased and the resonance state is weakened, and marking the airplane metal component to form an internal damage early warning signal;
measuring an energy response signal of the metal component of the airplane by using the acoustic emission sensor, filtering and denoising the energy response signal by using the acoustic emission signal collector, transmitting the energy response signal to a computer, drawing an energy response trend curve and a counting trend curve by using the computer, and judging that the current moment is a crack initiation moment when the obtained energy response trend and the counting trend are increased; when the slope values of the energy response trend curve and the counting trend curve exceed a preset slope threshold value, marking that cracks in the airplane metal component have spread or the airplane metal component is in a failure state;
and step five, measuring the surface temperature value of the airplane metal component when the trend of the acoustic emission signal changes by using a thermal infrared camera, processing the surface temperature value by using a thermal imaging analyzer, transmitting the surface temperature value to a computer, drawing a temperature maximum value and temperature standard deviation value curve in each measuring point in the area to be monitored of the airplane metal component by using the computer, finding out the position of the measuring point where the temperature maximum value is located when the temperature standard deviation value of the area to be monitored of the airplane metal component is in an ascending trend, indicating that the temperature of the position of the measuring point is increased, indicating that the temperature difference of the position of the measuring point is maximum and starts to be reduced when the temperature of the position of the measuring point is changed from increasing to decreasing, and determining that the change trend of the temperature maximum value is abnormal, and the position of the measuring point where the temperature maximum value is located is the measured damage position.
The multi-parameter measurement method for the vibration fatigue damage of the metal component of the airplane is characterized by comprising the following steps of: the horizontal sliding table is arranged on the vibrating table, and the horizontal sliding table is provided with a test fixture for mounting a metal component of the airplane.
The multi-parameter measurement method for the vibration fatigue damage of the metal component of the airplane is characterized by comprising the following steps of: the detection range of the thermal infrared camera covers the whole airplane metal component.
The method has the advantages that the change of the change trend of the motion displacement response signal is used as an early warning signal of the damage of the metal component of the airplane, whether the energy response trend and the counting trend of the acoustic emission signal are suddenly increased or not is observed, the maximum time of the temperature standard deviation value is searched by combining the change conditions of the surface temperature distribution and the temperature standard deviation value of the metal component of the airplane, whether the damage occurs or not is comprehensively judged, the damage position is determined, the multi-parameter accurate and rapid measurement of the damage of the metal component of the airplane is realized, the reliable basis is provided for the subsequent evolution process analysis of the damage of the metal component of the airplane, the vibration fatigue test is not required to be interrupted, and the damage measurement is jointly obtained through multiple parameters; the method has the characteristics of high sensitivity, can monitor the whole life cycle from initiation to expansion to failure of the crack, can find the crack in the first time when the metal component of the airplane is damaged, eliminates the interference of more environmental factors, overcomes the defect of hysteresis in the traditional method, and is convenient to popularize and use.
The technical solution of the present invention is further described in detail by the accompanying drawings and embodiments.
Drawings
Fig. 1 is a schematic structural diagram of a vibration fatigue test system of the present invention.
FIG. 2 is a measurement schematic block diagram of the vibration fatigue test system of the present invention.
FIG. 3 is a block diagram of a method flow of the present invention.
Description of reference numerals:
1-a vibration table; 2-horizontal sliding table; 3-test fixture;
4-aircraft metal components; 5-acoustic emission sensor; 6-thermal infrared camera;
7-laser emitter; 8, an acoustic emission signal acquisition instrument; 9-thermographic analyzer;
10-laser vibrometer; 11-computer.
Detailed Description
As shown in fig. 1 to 3, the method for measuring multiple parameters of the vibration fatigue damage of the metal component of the airplane comprises the following steps:
step one, a vibration fatigue test system is set up, the vibration fatigue test system comprises a vibration table 1 for mounting an airplane metal component 4, a thermal infrared camera 6 and a laser transmitter 7, the thermal infrared camera 6 and the laser transmitter are respectively arranged beside the vibration table 1 and used for collecting temperature signals and motion displacement response signals, and an acoustic emission sensor 5 is mounted on the airplane metal component 4;
the signal output end of the acoustic emission sensor 5 is connected with an acoustic emission signal collector 8, the signal output end of the thermal infrared camera 6 is connected with a thermal imaging analyzer 9, the signal output end of the laser emitter 7 is connected with a laser vibration meter 10, and the output ends of the acoustic emission signal collector 8, the thermal imaging analyzer 9 and the laser vibration meter 10 are all connected with a computer 11;
step two, starting a vibration fatigue test system;
measuring a movement displacement response signal of the airplane metal component 4 by using the laser transmitter 7, processing the movement displacement response signal by using the laser vibration meter 10, transmitting the data to the computer 11, continuously increasing the obtained movement displacement response value, and when the obtained movement displacement response value is changed from increasing to decreasing, indicating that the structure rigidity of the airplane metal component 4 is decreased, the natural frequency is decreased and the resonance state is weakened, and marking that the airplane metal component 4 forms an internal damage early warning signal;
measuring an energy response signal of the airplane metal component 4 by using the acoustic emission sensor 5, filtering and denoising the energy response signal by using the acoustic emission signal collector 8, transmitting the energy response signal to the computer 11, drawing an energy response trend curve and a counting trend curve by using the computer 11, and judging that the current moment is a crack initiation moment when the obtained energy response trend and the counting trend are increased; when the slope values of the energy response trend and the counting trend curve exceed a preset slope threshold value, marking that cracks in the airplane metal component 4 are expanded or the airplane metal component 4 is in a failure state;
measuring the surface temperature value of the airplane metal component 4 when the trend of the acoustic emission signal changes by using the thermal infrared camera 6, processing the surface temperature value by using the thermal imaging analyzer 9, transmitting the data to the computer 11, drawing a temperature maximum value and a temperature standard deviation curve in each measuring point in the area to be monitored of the airplane metal component 4 by using the computer 11, finding out the position of the measuring point where the temperature maximum value is located when the temperature standard deviation of the area to be monitored of the airplane metal component 4 is in an ascending trend, indicating that the temperature of the position of the measuring point is increased, indicating that the temperature difference of the position of the measuring point is maximum and begins to be reduced when the temperature of the position of the measuring point is changed from increasing to decreasing, and considering that the change trend of the temperature maximum value is abnormal, the internal damage of the airplane metal component 4 is generated and the position of the measuring point where the temperature maximum value is located is the measured damage position.
In this embodiment, a horizontal sliding table 2 is arranged on the vibrating table 1, and a test fixture 3 for installing an airplane metal component 4 is arranged on the horizontal sliding table 2.
In this embodiment, the detection range of the thermal infrared camera 6 covers the entire airplane metal component 4.
When the method is used, the change trend of the motion displacement response signal is used as an early warning signal of the damage of the metal component of the airplane, whether the energy response trend and the counting trend of the acoustic emission signal are suddenly increased or not is observed, the maximum time of the temperature standard deviation value is searched by combining the change conditions of the surface temperature distribution and the temperature standard deviation value of the metal component of the airplane, whether the damage occurs or not is comprehensively judged, the damage position is determined, the multi-parameter accurate and rapid measurement of the damage of the metal component of the airplane is realized, a reliable basis is provided for the subsequent evolution process analysis of the damage of the metal component of the airplane, the vibration fatigue test is not required to be interrupted, and the damage measurement is obtained through the combination of multiple parameters; the method has the characteristics of high sensitivity, can monitor the whole life cycle from the initiation to the expansion to the failure of the crack, can find the crack in the first time when the metal component of the airplane is damaged, eliminates the interference of more environmental factors, and overcomes the defect of hysteresis in the traditional method.
The above description is only a preferred embodiment of the present invention, and is not intended to limit the present invention, and all simple modifications, changes and equivalent structural changes made to the above embodiment according to the technical spirit of the present invention still fall within the protection scope of the technical solution of the present invention.
Claims (3)
1. A multi-parameter measurement method for vibration fatigue damage of an airplane metal component is characterized by comprising the following steps:
the method comprises the following steps of firstly, building a vibration fatigue test system, wherein the vibration fatigue test system comprises a vibration table (1) for mounting an airplane metal component (4), a thermal infrared camera (6) and a laser transmitter (7), the thermal infrared camera (6) and the laser transmitter are arranged beside the vibration table (1) and are used for collecting temperature signals and collecting motion displacement response signals respectively, and an acoustic emission sensor (5) is mounted on the airplane metal component (4);
the signal output end of the acoustic emission sensor (5) is connected with an acoustic emission signal collector (8), the signal output end of the thermal infrared camera (6) is connected with a thermal imaging analyzer (9), the signal output end of the laser emitter (7) is connected with a laser vibration meter (10), and the output ends of the acoustic emission signal collector (8), the thermal imaging analyzer (9) and the laser vibration meter (10) are connected with a computer (11);
step two, starting a vibration fatigue test system;
measuring a movement displacement response signal of the airplane metal component (4) by using the laser transmitter (7), processing the data by using the laser vibration meter (10), transmitting the data to the computer (11), continuously increasing the obtained movement displacement response value, and when the obtained movement displacement response value is changed from increasing to decreasing, indicating that the structure rigidity of the airplane metal component (4) is decreased, the natural frequency is decreased and the resonance state is weakened, and marking that the airplane metal component (4) forms an internal damage early warning signal;
measuring an energy response signal of the airplane metal component (4) by using the acoustic emission sensor (5), filtering and denoising the energy response signal by using the acoustic emission signal collector (8), and transmitting the energy response signal to the computer (11), drawing an energy response trend and counting trend curve by using the computer (11), and judging that the current moment is a crack initiation moment when the obtained energy response trend and counting trend are increased; when the slope values of the energy response trend curve and the counting trend curve exceed a preset slope threshold value, marking that cracks in the airplane metal component (4) have been expanded or the airplane metal component (4) is in a failure state;
measuring the surface temperature value of the airplane metal component (4) when the trend of the acoustic emission signal changes by using the thermal infrared camera (6), and the data is transmitted to a computer (11) after being processed by a thermal imaging analyzer (9), the computer (11) draws a curve of the maximum temperature value and the standard temperature deviation value in each measuring point in the region to be monitored of the metal component (4) of the airplane, when the standard deviation value of the temperature of the area to be monitored of the airplane metal component (4) is in an ascending trend, the position of a measuring point where the maximum temperature value is located is found out, the temperature of the position of the measuring point is increased, when the temperature of the position of the measuring point is changed from rising to falling, the temperature difference of the position of the measuring point reaches the maximum and begins to reduce, the change trend of the maximum temperature value is abnormal, and the measured position of the measuring point where the maximum temperature value is located is considered to be the measured damage position when the internal damage of the metal component (4) of the airplane is generated.
2. The multi-parameter measurement method for the vibration fatigue damage of the metal component of the airplane as claimed in claim 1, wherein the method comprises the following steps: the test bed is characterized in that a horizontal sliding table (2) is arranged on the vibrating table (1), and a test fixture (3) for installing an airplane metal component (4) is arranged on the horizontal sliding table (2).
3. The multi-parameter measurement method for the vibration fatigue damage of the metal component of the airplane as claimed in claim 1, wherein the method comprises the following steps: the detection range of the thermal infrared camera (6) covers the whole airplane metal component (4).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210733242.6A CN114813003A (en) | 2022-06-27 | 2022-06-27 | Multi-parameter measurement method for vibration fatigue damage of metal component of airplane |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210733242.6A CN114813003A (en) | 2022-06-27 | 2022-06-27 | Multi-parameter measurement method for vibration fatigue damage of metal component of airplane |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114813003A true CN114813003A (en) | 2022-07-29 |
Family
ID=82522681
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210733242.6A Pending CN114813003A (en) | 2022-06-27 | 2022-06-27 | Multi-parameter measurement method for vibration fatigue damage of metal component of airplane |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN114813003A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117622499A (en) * | 2024-01-25 | 2024-03-01 | 中国飞机强度研究所 | Aircraft damage visualization system |
Citations (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB0513901D0 (en) * | 2005-07-06 | 2005-08-10 | Airbus Uk Ltd | Method and apparatus for measuring the structural integrity of a safe-life aircraft component |
| CN102426197A (en) * | 2011-08-19 | 2012-04-25 | 北京航空航天大学 | Method for identifying damage of aircraft structural parts based on acoustic emission detection |
| CN102692188A (en) * | 2012-05-08 | 2012-09-26 | 浙江工业大学 | Dynamic crack length measurement method for machine vision fatigue crack propagation test |
| CN102809611A (en) * | 2011-06-02 | 2012-12-05 | 中国人民解放军装甲兵工程学院 | System and method for detecting damage of metal component nondestructively |
| JP2013140185A (en) * | 2013-04-19 | 2013-07-18 | National Institute For Materials Science | Cryogenic temperature ultrasonic fatigue nondestructive test evaluation apparatus |
| CN104634879A (en) * | 2015-02-04 | 2015-05-20 | 北京科技大学 | Metallic material fatigue loading testing and fatigue damage nondestructive testing analytical method |
| US20170052150A1 (en) * | 2015-08-20 | 2017-02-23 | U.S.A., as represented by the Administrator of the National Aeronautics and Space Administration | System and Method for Progressive Damage Monitoring and Failure Event Prediction in a Composite Structure |
| CN108548646A (en) * | 2018-03-28 | 2018-09-18 | 中国航发北京航空材料研究院 | The quantitative measuring method of damage development overall process in a kind of vibration fatigue test |
| CN112288715A (en) * | 2020-10-28 | 2021-01-29 | 湖南大学 | Method, device and equipment for evaluating fatigue damage of metal component and storage medium |
| CN113252794A (en) * | 2021-06-03 | 2021-08-13 | 沈阳工业大学 | Acoustic emission crack monitoring method and system |
| CN113465733A (en) * | 2021-08-13 | 2021-10-01 | 重庆大学 | Vibration table structure displacement response prediction method and device based on EEMD-DNN |
| CN113607580A (en) * | 2021-08-10 | 2021-11-05 | 江苏徐工工程机械研究院有限公司 | Metal component fatigue test method and residual life prediction method |
-
2022
- 2022-06-27 CN CN202210733242.6A patent/CN114813003A/en active Pending
Patent Citations (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB0513901D0 (en) * | 2005-07-06 | 2005-08-10 | Airbus Uk Ltd | Method and apparatus for measuring the structural integrity of a safe-life aircraft component |
| CN102809611A (en) * | 2011-06-02 | 2012-12-05 | 中国人民解放军装甲兵工程学院 | System and method for detecting damage of metal component nondestructively |
| CN102426197A (en) * | 2011-08-19 | 2012-04-25 | 北京航空航天大学 | Method for identifying damage of aircraft structural parts based on acoustic emission detection |
| CN102692188A (en) * | 2012-05-08 | 2012-09-26 | 浙江工业大学 | Dynamic crack length measurement method for machine vision fatigue crack propagation test |
| JP2013140185A (en) * | 2013-04-19 | 2013-07-18 | National Institute For Materials Science | Cryogenic temperature ultrasonic fatigue nondestructive test evaluation apparatus |
| CN104634879A (en) * | 2015-02-04 | 2015-05-20 | 北京科技大学 | Metallic material fatigue loading testing and fatigue damage nondestructive testing analytical method |
| US20170052150A1 (en) * | 2015-08-20 | 2017-02-23 | U.S.A., as represented by the Administrator of the National Aeronautics and Space Administration | System and Method for Progressive Damage Monitoring and Failure Event Prediction in a Composite Structure |
| CN108548646A (en) * | 2018-03-28 | 2018-09-18 | 中国航发北京航空材料研究院 | The quantitative measuring method of damage development overall process in a kind of vibration fatigue test |
| CN112288715A (en) * | 2020-10-28 | 2021-01-29 | 湖南大学 | Method, device and equipment for evaluating fatigue damage of metal component and storage medium |
| CN113252794A (en) * | 2021-06-03 | 2021-08-13 | 沈阳工业大学 | Acoustic emission crack monitoring method and system |
| CN113607580A (en) * | 2021-08-10 | 2021-11-05 | 江苏徐工工程机械研究院有限公司 | Metal component fatigue test method and residual life prediction method |
| CN113465733A (en) * | 2021-08-13 | 2021-10-01 | 重庆大学 | Vibration table structure displacement response prediction method and device based on EEMD-DNN |
Non-Patent Citations (4)
| Title |
|---|
| 彭艳涛等: "铝合金振动疲劳同步测量技术研究", 《航空工程进展》 * |
| 李斌等: "基于红外和声发射的复合材料疲劳损伤实时监测", 《机械科学与技术》 * |
| 王锟等: "平纹编织C/SiC复合材料拉―拉疲劳特性的试验研究", 《机械强度》 * |
| 申雅峰等: "铝合金疲劳试验的声发射滤波技术研究", 《计测技术》 * |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117622499A (en) * | 2024-01-25 | 2024-03-01 | 中国飞机强度研究所 | Aircraft damage visualization system |
| CN117622499B (en) * | 2024-01-25 | 2024-04-16 | 中国飞机强度研究所 | Aircraft damage visualization system |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US20110213569A1 (en) | Method and device for detecting cracks in compressor blades | |
| CN107063679B (en) | Method and device for quickly detecting gear defects by structure tuned resonance | |
| CN107607247B (en) | Explosive explosion impulse and wind pressure combined test method | |
| CN111521681A (en) | Concrete internal damage assessment method based on piezoelectric ceramic shear wave energy loss | |
| CN114813003A (en) | Multi-parameter measurement method for vibration fatigue damage of metal component of airplane | |
| CN111307487A (en) | Rotating mechanical vibration measurement method based on micro-motion amplification | |
| CN112213092A (en) | Measuring method for testing internal force increment of arch bridge suspender by adopting inertia method | |
| Checkel et al. | Testing a third derivative knock indicator on a production engine | |
| US11353034B2 (en) | Method and device for determining an indicator for a prediction of an instability in a compressor and use thereof | |
| KR101997993B1 (en) | Crack inspection device and crack inspecting method using the same | |
| CN104913988A (en) | Hopkinson principle-based concrete axial tensile strength measuring method | |
| US6923046B2 (en) | Arrangement and method to measure cylinder pressure in a combustion engine | |
| KR102247305B1 (en) | Gas turbine combustion instability diagnosis system based on zero crossing rate and method thereof | |
| CN112097717B (en) | Gap detection system and method based on collision vibration | |
| CN108414217B (en) | Gear box noise test system | |
| CN114235951B (en) | Crack fault acoustic diagnosis method and device for engine air inlet casing support plate | |
| US9970373B1 (en) | Method and system for detecting and eliminating knocking | |
| CN202362112U (en) | Detonation sensor testing arrangement | |
| US11624687B2 (en) | Apparatus and method for detecting microcrack using orthogonality analysis of mode shape vector and principal plane in resonance point | |
| KR20140139955A (en) | System and method of detecting failure of engine | |
| CN212110560U (en) | Stator blade natural frequency detection device under complicated path | |
| CN110988138B (en) | Weld assembly quality detection device and method | |
| Tamura et al. | Non-contact vibration measurement of the rotor blades that play a pivotal role in the reliability of gas turbines | |
| CN106370726B (en) | A kind of damage detection system and its detection method of Two-dimensional Composites | |
| CN111319787A (en) | Helicopter moving part vibration monitoring data validity evaluation method |
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 | ||
| RJ01 | Rejection of invention patent application after publication |
Application publication date: 20220729 |
|
| RJ01 | Rejection of invention patent application after publication |