CN109733641A - A kind of aircraft full size structure part multiaxle fatigue experimental method - Google Patents
A kind of aircraft full size structure part multiaxle fatigue experimental method Download PDFInfo
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- CN109733641A CN109733641A CN201910057921.4A CN201910057921A CN109733641A CN 109733641 A CN109733641 A CN 109733641A CN 201910057921 A CN201910057921 A CN 201910057921A CN 109733641 A CN109733641 A CN 109733641A
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Abstract
The invention discloses a kind of aircraft full size structure part multiaxle fatigue experimental methods, are related to aircraft structure fatigue life test field, steps of the method are: (1) the stress concentration portion position of structural member is obtained using finite element simulation calculation；(2) foil gauge is pasted around stress concentration portion position or position；(3) the actual measurement spectrum of each foil gauge and strain rosette is recorded, and filters and obtains test strain spectrum；(4) to aircraft structure force analysis and simplified external force, according to simplified external force development test platform；(5) the multidirectional external force loading spectrum that the strain spectrum of actual measurement is transformed into testing stand is calculated using finite element optimization；(6) loading spectrum to rise and fall is loaded, whether compares the relative error for surveying strain spectrum on this loading spectrum strain value and aircraft less than 10%, satisfaction then carries out fatigue life test, and otherwise return step (4) re-starts force analysis.Test result illustrates that this method can reappear the multidirectional stressed situation of aircraft structure in actual use well.
Description
Technical field
Application field of the present invention is to belong to aircraft full size structure part fatigue test technology field, more particularly to refer in particular to one kind
Aircraft full size structure part multiaxle fatigue experimental method.
Background technique
The main failure forms of aircaft configuration are fatigue rupture, it is therefore desirable to be surveyed when aircraft is on active service using aircraft flight
Load history carries out the fatigue test of complete machine, and the fatigue life of complete machine is assessed with this, when advanceing to the longevity so as to aircraft, sends out in time
The weak link of existing aircaft configuration takes maintenance measure in advance.But the test period of full-scale fatigue test is longer, and for tool
The structure of body is lack of pertinence, and can not really reflect aircraft specific structure multidirectional (multiaxis) loading conditions complicated when being on active service,
Lead to the situation for occurring determining longevity inaccuracy, therefore the dedicated fatigue test technology invented for aircraft specific structure part has
Highly important practical meaning in engineering.
It is mostly at present single shaft fatigue test for the fatigue test method of aircraft structure, and aircraft structure is in practical clothes
Using as a servant in use process is mostly non-proportional loading state, will be unable to actual response using traditional single shaft fatigue test method at this time
For aircraft structure in commission stress, it is even more impossible to targetedly carry out accurately determining the longevity to aircraft structure.It is proposed
A kind of aircraft full size structure part multiaxle fatigue experimental method, can be completed in a short time with specific aim and accurately fly
Machine structural member fatigue test.
Summary of the invention
Present invention aims to meet the needs of aircraft structure fatigue life test, it is full-scale to propose a kind of aircraft
Structural member non-proportional loading life test method, this method are also applied for other mechanical structures and carry out fatigue life test.
The technical solution adopted by the present invention is a kind of aircraft full size structure part multiaxle fatigue experimental method, realizes this method
The steps include:
Step 1): determine that aircraft structure stress concentrates the position of most serious, record using finite element simulation calculation method
This position.It differs if there is the maximum value at multiple stress concentration portion positions and each position less than 20%, then records multiple positions；
Step 2): the stress concentration portion position recorded according to step 1), array is pasted in corresponding position on aircraft structure
Unidirectional strain gauge adhesion is concentrated root to measure the strain in stress maximum direction, in this foil gauge in stress by foil gauge
4 triaxial strain flowers are at least pasted in surrounding 20mm, measure the strain and stress maximum direction journey in stress maximum direction respectively
The strain in 90 ° of directions and strain with stress maximum direction 45 ° of directions of journey, if stress concentration portion position root can not paste strain
Piece then at least pastes 4 triaxial strain flowers in the 30mm around stress concentration portion position, measures answering for stress maximum direction respectively
Become, the strain with the strain in 90 ° of directions of stress maximum direction journey and with stress maximum direction 45 ° of directions of journey；
Step 3): record Typical Aircraft rise and fall in each foil gauge and strain rosette data, the strain in each direction filter simultaneously
Wave, and remove the side crops industry not damaged, it is chosen in all data risen and fallen and always damages maximum data of rising and falling as fatigue
Test alternating load course；
Step 4): the actual loading analysis of aircraft structure is carried out, the stress of structural member is simplified, at least reservation both direction
External force, according to the external force retained use actuator development test platform, retain actuator direction and aircraft structure
External force direction is consistent；
Step 5): establishing the finite element model of aircraft structure using FEM Simulation, using more in step 4)
Outward force is as boundary condition, and since first point of strain spectrum in step 3), the value of each foil gauge and strain rosette is made
For target value, multidirectional external force takes the efficient working range of actuator as independent variable, independent variable range, optimizes analysis and obtain
The value of multiple directions external force corresponding to entire strain spectrum, converts multiaxle fatigue experimental for the multiaxial strain load history of actual measurement
Loading spectrum；
Step 6): the loading spectrum obtained according to step 5) carries out the test risen and fallen with the Special test platform developed,
And the strain value of acquisition and same position on aircraft, compare the relative error of strain spectrum obtained in this strain value and step 3),
Fatigue test can be carried out if it is less than 10%, step 4) is returned to if it is greater than 10% and recalculates.
Compared with prior art, the present invention has the advantages that.
The present invention has the advantages that proposing a kind of aircraft full size structure part multiaxle fatigue experimental method.This method energy
It is enough that the actual measurement load history of Aircraft Air is transformed on multiaxle fatigue experimental, it is true to react aircraft structure institute at work
The alternating load situation of receiving guarantees the authenticity and accuracy of aircraft structure fatigue test.More traditional airplane complete machine is tired
Labor test is compared, and this method can really go back stress condition of the origianl structure part in aircraft actual use, test result and reality
Situation is consistent, and with strong points, can accurately react its life situations for the weak structure position of aircraft.
Detailed description of the invention
The flow chart of Fig. 1 the method for the present invention realization aircraft components ground experiment.
Fig. 2 strain gauge adhesion position view of the present invention.
The structural member stress analysis schematic diagram of Fig. 3 the method for the present invention.
Specific embodiment
A specific embodiment of the invention is described with reference to the drawings.
The present invention is further illustrated the present invention by the fatigue test of aircraft full size structure part,
A kind of aircraft full size structure part multiaxle fatigue experimental method, circular are as follows:
Step 1): show that the aircraft structure stress concentrates the position of most serious, note using finite element simulation calculation method
Record this position；
Step 2): the stress concentration portion position recorded according to step 1), corresponding position is pasted foil gauge and is answered aboard
Become flower, the stress of the structural member concentrates root that can not paste foil gauge, pastes a survey master in this structural member stress concentration portion position
The foil gauge of Impact direction, surrounding paste 4 triaxial strain flowers, and strain and the stress for measuring stress maximum direction respectively are maximum
The strain in 90 ° of directions of direction journey and strain with stress maximum direction 45 ° of directions of journey, as shown in Figure 2；
Step 3): each foil gauge and strain rosette data that the record each typical case of aircraft rises and falls, the strain in each direction is simultaneously
Filtering, and remove the side crops industry not damaged, it is chosen in all data risen and fallen and always damages maximum data of rising and falling as tired
Labor tests strain spectrum；
Step 4): the actual loading analysis of aircraft structure is carried out, the stress of structural member is simplified, at least reservation both direction
External force, testing stand is built using actuator according to the external force retained, makes actuator direction and aircraft structure force analysis
The external force direction retained is consistent, as shown in Figure 3；
Step 5): establishing the finite element model of aircraft structure using FEM Simulation, using more in step 4)
Outward force is as boundary condition, since first point of strain history in step 3), by the value of each foil gauge and strain rosette
As target value, multidirectional external force takes the efficient working range of actuator, optimizes and analyze as independent variable, independent variable range
To the value of multiple directions external force corresponding to entire strain history, multiaxle fatigue experimental is converted by the multiaxial strain course of actual measurement
Loading spectrum；
Step 6): the loading spectrum obtained according to step 5) carries out the test risen and fallen with the testing stand developed, and adopts
The strain value of collection and same position on aircraft, the relative error for comparing strain spectrum obtained in this strain value and step 3) are less than
10%, carry out fatigue test.
The present invention has the advantages that proposing a kind of aircraft full size structure part multiaxle fatigue experimental method.This method energy
It is enough that the actual measurement load history of Aircraft Air is transformed into multiaxle fatigue experimental load spectrum, to truly react aircraft structure
Stress condition, it is ensured that the authenticity and accuracy of fatigue test.More traditional aircraft structure fatigue test is compared, we
Method can really go back loading conditions of the origianl structure part in aircraft actual use, and test result is consistent with actual conditions, and needle
It is strong to property, its life situations can accurately be reacted for the weak structure of aircraft.
In order to verify the effect of aircraft full size structure part multiaxle fatigue experimental method proposed by the present invention, this method is answered
It uses on aircraft full size structure part, by verification experimental verification, using the crack growth rate and reality of the fatigue test of this method
The crack growth rate relative error of border military service aircraft structure really reflects aircraft full size structure part in reality less than 3%
Life situations in the use of border.Therefore, a kind of aircraft full size structure part multiaxle fatigue experimental method of invention can be accurately
Carry out the non-proportional loading life test of aircraft full size structure.
Claims (4)
1. a kind of aircraft full size structure part multiaxle fatigue experimental method, it is characterised in that: realize steps of the method are,
Step 1): it determines that aircraft structure stress concentrates the position of most serious using finite element simulation calculation method, records this position
It sets.It differs if there is the maximum value at multiple stress concentration portion positions and each position less than 20%, then records multiple positions；
Step 2): the stress concentration portion position recorded according to step 1), array strain is pasted in corresponding position on aircraft structure
Unidirectional strain gauge adhesion is concentrated root to measure the strain in stress maximum direction, around this foil gauge in stress by piece
4 triaxial strain flowers are at least pasted in 20mm, measure strain and 90 ° of sides of stress maximum direction journey in stress maximum direction respectively
To strain and strain with stress maximum direction 45 ° of directions of journey, if stress concentration portion position root can not paste foil gauge,
4 triaxial strains flower is at least pasted in 30mm around stress concentration portion position, measure respectively stress maximum direction strain, with by
The strain in power maximum direction 90 ° of directions of journey and strain with stress maximum direction 45 ° of directions of journey；
Step 3): record Typical Aircraft rise and fall in each foil gauge and strain rosette data, the strain in each direction filter simultaneously, and
Remove the side crops industry not damaged, is chosen in all data risen and fallen and always damage maximum data of rising and falling as fatigue test and answer
Variable load history；
Step 4): carrying out the actual loading analysis of aircraft structure, simplifies the stress of structural member, and at least reservation both direction is outer
Power uses actuator development test platform, the external force for retaining actuator direction and aircraft structure according to the external force retained
Direction is consistent；
Step 5): establishing the finite element model of aircraft structure using FEM Simulation, using multidirectional outer in step 4)
Power is as boundary condition, since first point of strain spectrum in step 3), using the value of each foil gauge and strain rosette as mesh
Scale value, multidirectional external force take the efficient working range of actuator as independent variable, independent variable range, optimize analysis and obtain entirely
The value of multiple directions external force corresponding to strain spectrum, converts the multiaxial strain load history of actual measurement to the load of multiaxle fatigue experimental
Lotus spectrum；
Step 6): the loading spectrum obtained according to step 5) carries out the test risen and fallen with the Special test platform developed, and adopts
The strain value of collection and same position on aircraft, compares the relative error of strain spectrum obtained in this strain value and step 3), if
Fatigue test can be carried out less than 10%, returned to step 4) if it is greater than 10% and recalculated.
2. a kind of aircraft components dangerous position ground simulation according to claim 1 loaded test method in the air, feature
Be: the step 2) pastes array foil gauge in aircraft structure stress concentration portion position, and unidirectional strain gauge adhesion is existed
Stress concentrates root to measure the strain in stress maximum direction, at least pastes 4 triaxial strains in 20mm around this foil gauge
Flower, measure respectively stress maximum direction strain, with the strain in 90 ° of directions of stress maximum direction journey and with stress maximum direction journey
The strain in 45 ° of directions, if stress concentration portion position root can not paste foil gauge, in the 30mm around stress concentration portion position extremely
4 triaxial strain flowers, the strain for measuring stress maximum direction respectively, the strain with stress maximum direction 90 ° of directions of journey are pasted less
With the strain with stress maximum direction 45 ° of directions of journey.
3. a kind of aircraft components dangerous position ground simulation according to claim 1 loaded test method in the air, feature
Be: the step 4) carries out the actual loading analysis of aircraft structure, simplifies the stress of structural member, at least reservation both direction
External force, according to the external force retained use actuator development test platform, retain actuator direction and aircraft structure
External force direction is consistent.
4. a kind of aircraft components dangerous position ground simulation according to claim 1 loaded test method in the air, feature
Be: the step 5) establishes the finite element model of aircraft structure using FEM Simulation, using more in step 4)
Outward force is as boundary condition, and since first point of strain spectrum in step 3), the value of each foil gauge and strain rosette is made
For target value, multidirectional external force takes the efficient working range of actuator as independent variable, independent variable range, optimizes analysis and obtain
The value of multiple directions external force corresponding to entire strain spectrum, converts multiaxle fatigue experimental for the multiaxial strain load history of actual measurement
Loading spectrum.
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Cited By (4)
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| CN111624116A (en) * | 2020-05-09 | 2020-09-04 | 清华大学 | Fatigue life prediction method and device based on weighted average maximum shear stress plane |
| CN112179595A (en) * | 2020-09-25 | 2021-01-05 | 中国直升机设计研究所 | Helicopter body fairing vibration fatigue test verification method |
| CN113420366A (en) * | 2021-04-20 | 2021-09-21 | 中国直升机设计研究所 | Method for verifying bonding strength of blade anti-icing and deicing heating assembly |
| US20220180670A1 (en) * | 2019-08-28 | 2022-06-09 | Denso Corporation | Electric vertical takeoff and landing aircraft and control device for electric vertical takeoff and landing aircraft |
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| CN113420366A (en) * | 2021-04-20 | 2021-09-21 | 中国直升机设计研究所 | Method for verifying bonding strength of blade anti-icing and deicing heating assembly |
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Application publication date: 20190510 |