CN109815521A - A kind of appraisal procedure of the anti-FOD ability of blade of aviation engine - Google Patents
A kind of appraisal procedure of the anti-FOD ability of blade of aviation engine Download PDFInfo
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- CN109815521A CN109815521A CN201811464884.0A CN201811464884A CN109815521A CN 109815521 A CN109815521 A CN 109815521A CN 201811464884 A CN201811464884 A CN 201811464884A CN 109815521 A CN109815521 A CN 109815521A
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Abstract
The present invention relates to a kind of appraisal procedures of the anti-FOD ability of blade of aviation engine, the shock dynamics for carrying out foreign object damage to simulation blade numerical model in this method emulates, obtain notch gross feature and foreign object type, impact velocity, the relationship between angle of attack, and as data according to determining foreign object damage experimental condition, outer analogies damage test, the macro and micro feature of observation impact notch are carried out to simulation blade using air bubble；Using under blade inlet edge danger position service load static stress and dynamic stress as high-cycle fatigue test initial static load and dynamic load, high-cycle fatigue test is carried out to the simulation blade after damage, the high-cycle fatigue strength of blade is obtained by step-by-step method, according to the anti-FOD ability of high-cycle fatigue test outcome evaluation blade；Simulation foreign object damage is carried out to a small number of real blades and high-cycle fatigue test obtains its high-cycle fatigue strength, to verify the accordance of simulation blade and real blade test result.
Description
Technical field
The present invention relates to blade of aviation engine performance test fields.
Background technique
For aircraft during runway or type aircraft carrier deck landing, the aero-engine of high revolving speed operation can often suck stone, sand
When the hard object such as gravel, bolt, metal fragment, these foreign matters enter engine and the relative velocity of compressor blade is up to 100m/s-
350m/s can cause serious impact injury to damage to the leading edge of preceding what compressor blade especially blade.In the prior art,
The hard objects such as metal, sandstone are hit into the damage of engine bring and are referred to as " foreign object damage " (abbreviation FOD, Foreign
Object Damage)
If the blade being damaged finds and handles not in time, fatigue fracture failure may occur under high-frequency vibration, fly
Disconnected blade can punch the blade of what compressor after casing even destroys, and cause serious aircraft accident.Therefore foreign object is probed into
The affecting laws to blade material fatigue strength are damaged, the anti-foreign object damage ability of blade is assessed, is that blade of aviation engine is set
Very important a part in meter.
Domestic is mainly at present referring to such as U.S. army's table to the assessment of the anti-foreign object damage of blade of aviation engine (FOD) ability
Relevant regulations in the standards such as MIL-HDBK-1783B, examination engine cause blade to be equivalent to minimum fatigue after sucking foreign object
Notch factor KfWhen=3 damage, can engine work to two inspection cycles as defined in specification or hourage.However blade
It is that the damage location that impact is formed can have microfissure, stress is concentrated, remaining by the basic reason that can be broken after FOD
Stress and microstructure damage, damage position easily become tired source under high all service loads and fatigue fracture occur.And metal
There are biggish dispersibilities for high cycle fatigue failure, and the anti-foreign object damage of blade is examined only by the periodicity that can be worked after damage
The ability of wound, can not be effectively prevented from foreign object damage leads to the risk of blade failure.
Summary of the invention
Goal of the invention: the present invention simulates the high-cycle fatigue test of foreign object damage test, blade by air bubble, is damaged
The high cycle fatigue performance of blade, to assess the anti-FOD ability of blade.
Technical solution:
A kind of appraisal procedure of the anti-FOD ability of blade of aviation engine, which comprises the following steps:
(1), the numerical model of tested real blade is established, and selects real blade leading edge locus in the numerical model
A little it is used as dynamic stress rest point position；Extract the Modal Stress σ of this positiona1；Extract the mould of dynamic stress rest point position
State stress σ01, dynamic stress rest point position true stress σ0；Calculate the vibration stress σ under point position service loada:Meanwhile the static stress result σ of the point position is obtained in the numerical model of the real bladem；
(2), it is designed according to leading-edge radius R, the leading edge angulation θ of the blade inlet edge point position chosen in step (1)
And blade inlet edge simulation test piece is produced, and establish the numerical model of the blade inlet edge simulation test piece；
(3), different materials, various sizes of foreign object are simulated in dynamics analysis software with different impact velocitys, punching
Hit angle hit blade inlet edge simulation test piece numerical model process, obtain damage notch gross feature and foreign object type,
The simulation result of relationship between impact velocity, angle of attack；
(4), criterion boundary is repaired according to the damage data of aero-engine outfield blade statistics or engine blade correlation
Size determines the injury severity score index for needing to examine, and the injury severity score index of the corresponding examination, selects step
(3) meet foreign object type, the impact velocity, the angle of shock of the simulation foreign object damage test of the injury severity score index of the examination in
After degree, foreign object is emitted using air bubble, blade inlet edge simulation test piece is impacted and blade inlet edge simulation test piece is caused to damage；
(5), high-cycle fatigue test is carried out to the simulation blade after damage by high-cycle fatigue test equipment；With step (1)
In static stress σ under the leading edge danger position service load that is calculatedmAs the static load of high-cycle fatigue test, danger position work
Make the vibration stress σ under loadaAs high-cycle fatigue test dynamic load, service life N circulation is designed to the simulation blade after damage
High-cycle fatigue test；N is positive integer；
(6), the anti-FOD capability evaluation of blade: it is disconnected that the more than one piece as similar in degree of injury simulates blade generation fatigue in N is recycled
It splits, then illustrates that blade is not able to satisfy in the fatigue strength that the blade after being damaged by this kind of degree corresponded under the service life and start
The security requirement of machine service load；If fatigue fracture does not occur in N circulation, step-by-step method is used, each step test carries out N
Circulation, with dynamic load σ at this timeaDynamic load for initial dynamic load, each step increases compared with previous step, until fatigue fracture occurs, obtains
Fatigue strength after damage under test specimen specified life.
It further, further include step (7), real blade verification experimental verification: to the up-front point position of several real blades
Foreign object damage is carried out to test to obtain impact injury；High-cycle fatigue test is carried out to the real blade after damage and obtains its fatigue by force
Degree, to verify the accordance of simulation blade fatigue intensity results and real blade result.
Further, in step (5), N >=3 × 107。
Further, in step (1), utilizing works modeling software establishes the numerical value of the numerical model of tested real blade
Model carries out model analysis to tested real blade finite element model in finite element analysis software, to blade finite element mould
Type applies revolving speed identical with service load and boundary condition, calculates dynamic frequency and the vibration stress distribution of blade, extracts blade
Preceding fiber stress maximum position takes this position as dynamic stress rest point position to assess since this position vibration stress is maximum
The anti-foreign object damage ability of blade.
Further, in step (4), injury severity score index includes type of impairment, size, stress concentration degree.
The utility model has the advantages that the present invention carries out high-cycle fatigue test to the simulation blade after damage, blade is obtained by step-by-step method
High-cycle fatigue strength, according to high-cycle fatigue test as a result, it is possible to more accurately assess the anti-FOD ability of blade, so as to have
Foreign object damage is avoided to effect to lead to the risk of blade failure.
Detailed description of the invention
Fig. 1 is the flow chart of the appraisal procedure of the anti-FOD ability of blade of aviation engine of the present invention.
Fig. 2 is the structure chart of blade inlet edge simulation test piece.
Fig. 3 is the A-A sectional view of blade inlet edge simulation test piece in Fig. 2.
Specific embodiment
In the following, being described in further details in conjunction with attached drawing to the present invention.
As shown in Figure 1, the present invention discloses a kind of appraisal procedure of anti-FOD ability of blade of aviation engine, including following step
It is rapid:
A kind of appraisal procedure of the anti-FOD ability of blade of aviation engine, comprising the following steps:
(1), the Numerical-Mode of blade blade vibration stress and Static Strength Analysis: is established using UG or Other Engineering modeling software
Type carries out model analysis to real blade finite element model in ANSYS or other finite element analysis softwares, to blade finite element
Model applies revolving speed identical with service load and boundary condition, calculates dynamic frequency and the vibration stress distribution of blade.Extract leaf
Fiber stress maximum position before piece injures this position as sucked external hard object in engine operation since this position vibration stress is maximum
Danger the most is set, therefore takes this position as point position is tested to assess the anti-foreign object damage ability of blade.This test is extracted to survey
The Modal Stress σ of point positiona1；Extract the Modal Stress σ of dynamic stress rest point position01, dynamic stress rest point position is true
Stress σ0；Calculate the vibration stress σ under point position service loada:Meanwhile in ANSYS or other are limited
Static strength calculating is carried out to the finite element model of real blade in meta analysis software, identical with service load turn is applied to model
Speed and boundary condition calculate the static stress distribution of blade.Extract the static stress result σ under danger position service loadm。
(2), as shown in Figure 2 and Figure 3, according to the leading-edge radius R for the blade inlet edge point position chosen in step (1), leading edge
Angulation θ is designed and is produced blade inlet edge simulation test piece, and establishes the numerical model of the blade inlet edge simulation test piece；
The blade is suitable for high all tensile fatigue tests need to set other kinds of high-cycle fatigue test according to correlation test criterion
Leading edge geometrical characteristic can be characterized and be suitable for the testpieces of high-cycle fatigue test by counting other.
(3), the Numerical Simulation Analysis of foreign object damage: design in UG or Other Engineering modeling software establishment step (1) is utilized
Blade inlet edge test point position blade numerical model, simulated in LS-DYNA or other dynamics analysis softwares different
Material, various sizes of foreign object hit the process of simulation blade with different impact velocitys, angle of attack, obtain damage notch
Gross feature and foreign object type, impact velocity, between angle of attack relationship simulation result.For the foreign object damage mould for simulating blade
Quasi- test provides data foundation.
(4), criterion boundary is repaired according to the damage data of aero-engine outfield blade statistics or engine blade correlation
Size determines the injury severity score index for needing to examine, and the injury severity score index of the corresponding examination, selects step
(3) meet foreign object type, the impact velocity, the angle of shock of the simulation foreign object damage test of the injury severity score index of the examination in
After degree, foreign object is emitted using air bubble, blade inlet edge simulation test piece is impacted and blade inlet edge simulation test piece is caused to damage；
Injury severity score index includes type of impairment, size, stress concentration degree.
(5), high-cycle fatigue test is carried out to the simulation blade after damage by high-cycle fatigue test equipment；With step (1)
In static stress σ under the leading edge danger position service load that is calculatedmAs the static load of high-cycle fatigue test, danger position work
Make the vibration stress σ under loadaAs high-cycle fatigue test dynamic load, service life N circulation is designed to the simulation blade after damage
High-cycle fatigue test；N is positive integer；Titanium alloy blade generally requires N=3 × 107, a small number of standard requirements N=10 at this stage9。
(6), the anti-FOD capability evaluation of blade: it is disconnected that the more than one piece as similar in degree of injury simulates blade generation fatigue in N is recycled
It splits, then illustrates that blade is not able to satisfy in the fatigue strength that the blade after being damaged by this kind of degree corresponded under the service life and start
The security requirement of machine service load；If fatigue fracture does not occur in N circulation, step-by-step method is used, each step test carries out N
Circulation, with dynamic load σ at this timeaDynamic load for initial dynamic load, each step increases compared with previous step, until fatigue fracture occurs, obtains
Fatigue strength after damage under test specimen specified life.
Step (7), real blade verification experimental verification: foreign object damage test is carried out to the up-front point position of several real blades
Obtain impact injury；High-cycle fatigue test is carried out to the real blade after damage and obtains its fatigue strength, to verify simulation blade
The accordance of fatigue strength result and real blade result.
Claims (5)
1. a kind of appraisal procedure of the anti-FOD ability of blade of aviation engine, which comprises the following steps:
(1), the numerical model of tested real blade is established, and selects real blade leading edge locus certain point in the numerical model
As dynamic stress rest point position；Extract the Modal Stress σ of this positiona1；The mode for extracting dynamic stress rest point position is answered
Power σ01, dynamic stress rest point position true stress σ0；Calculate the vibration stress σ under point position service loada:Meanwhile the static stress result σ of the point position is obtained in the numerical model of the real bladem；
(2), it designs and makes according to leading-edge radius R, the leading edge angulation θ of the blade inlet edge point position chosen in step (1)
Blade inlet edge simulation test piece is produced, and establishes the numerical model of the blade inlet edge simulation test piece；
(3), different materials, various sizes of foreign object are simulated in dynamics analysis software with different impact velocitys, the angle of shock
Degree hits the process of blade inlet edge simulation test piece numerical model, obtains the gross feature and foreign object type, impact of damage notch
The simulation result of relationship between speed, angle of attack；
(4), criterion boundary dimensions is repaired according to the damage data of aero-engine outfield blade statistics or engine blade correlation,
It determines the injury severity score index for needing to examine, and the injury severity score index of the corresponding examination, selects symbol in step (3)
After the foreign object type, impact velocity, angle of attack of closing the simulation foreign object damage test of the injury severity score index of the examination, adopt
Emit foreign object with air bubble, impacts blade inlet edge simulation test piece and blade inlet edge simulation test piece is caused to damage；
(5), high-cycle fatigue test is carried out to the simulation blade after damage by high-cycle fatigue test equipment；It is fallen into a trap with step (1)
Static stress σ under obtained leading edge danger position service loadmAs the static load of high-cycle fatigue test, danger position work is carried
Vibration stress σ under lotusaAs high-cycle fatigue test dynamic load, the height of service life N circulation is designed to the simulation blade after damage
All fatigue tests；N is positive integer；
(6), the anti-FOD capability evaluation of blade: the more than one piece as similar in degree of injury simulates blade and fatigue fracture occurs in N is recycled,
Then illustrate that blade corresponds to the fatigue strength under the service life in the blade after being damaged by this kind of degree and is not able to satisfy engine
The security requirement of service load；If fatigue fracture does not occur in N circulation, step-by-step method is used, each step test carries out N and follows
Ring, with dynamic load σ at this timeaDynamic load for initial dynamic load, each step increases compared with previous step, until fatigue fracture occurs, is damaged
Fatigue strength after wound under test specimen specified life.
2. the appraisal procedure of the anti-FOD ability of blade of aviation engine according to claim 1, it is characterised in that: further include
Step (7), real blade verification experimental verification: several real blades up-front point position progress foreign object damage is tested and is impacted
Damage；High-cycle fatigue test is carried out to the real blade after damage and obtains its fatigue strength, blade fatigue intensity is simulated with verifying
As a result with the accordance of real blade result.
3. the appraisal procedure of the anti-FOD ability of blade of aviation engine according to claim 1, it is characterised in that: step (5)
In, N >=3 × 107。
4. the appraisal procedure of the anti-FOD ability of blade of aviation engine according to claim 1, it is characterised in that: step (1)
In, utilizing works modeling software establishes the numerical model of the numerical model of tested real blade, in finite element analysis software
Model analysis is carried out to tested real blade finite element model, identical with service load turn is applied to blade finite element model
Speed and boundary condition calculate dynamic frequency and the vibration stress distribution of blade, blade inlet edge stress maximum position are extracted, due to this position
Vibration stress maximum is set, this position is taken as dynamic stress rest point position and assess the anti-foreign object damage ability of blade.
5. the appraisal procedure of the anti-FOD ability of blade of aviation engine according to claim 1, it is characterised in that: step (4)
In, injury severity score index includes type of impairment, size, stress concentration degree.
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| CN201811464884.0A CN109815521B (en) | 2018-12-03 | 2018-12-03 | Method for evaluating FOD resistance of aero-engine blade |
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| CN201811464884.0A CN109815521B (en) | 2018-12-03 | 2018-12-03 | Method for evaluating FOD resistance of aero-engine blade |
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| CN110595709A (en) * | 2019-08-14 | 2019-12-20 | 南京航空航天大学 | Method for determining allowable amplitude of turbine engine blade |
| CN110884684A (en) * | 2019-12-04 | 2020-03-17 | 中国直升机设计研究所 | Design method for strength test of helicopter after impact of bearingless tail blade |
| CN110889172A (en) * | 2019-12-04 | 2020-03-17 | 中国直升机设计研究所 | Method for prefabricating impact damage of helicopter rotor system |
| CN110988002A (en) * | 2019-11-27 | 2020-04-10 | 南京航空航天大学 | Rapid imaging method for microcracks of damaged section of foreign object based on image recognition |
| CN110987389A (en) * | 2019-11-27 | 2020-04-10 | 南京航空航天大学 | Foreign object damage notch calibration method with stress concentration coefficient not less than 3 |
| CN111046490A (en) * | 2019-11-27 | 2020-04-21 | 南京航空航天大学 | Grid size inversion method in foreign object damage gap analysis |
| CN112881203A (en) * | 2021-01-15 | 2021-06-01 | 北京航空航天大学 | Preparation device and method for foreign object damage notch of blade leading edge simulation piece |
| CN113075066A (en) * | 2021-03-22 | 2021-07-06 | 中国人民解放军空军工程大学 | Laser strengthening blade edge performance test method |
| CN113378327A (en) * | 2021-07-02 | 2021-09-10 | 吉林重通成飞新材料股份公司 | Wind power blade trailing edge cracking resistance design method, device, equipment and storage medium |
| CN114171139A (en) * | 2021-10-20 | 2022-03-11 | 中国航发四川燃气涡轮研究院 | Material selection method for compressor blade |
| CN114492113A (en) * | 2022-01-05 | 2022-05-13 | 南京航空航天大学 | Impact damage numerical simulation optimization method based on laser mapping entity grid |
| CN114526920A (en) * | 2022-02-21 | 2022-05-24 | 中国航发贵阳发动机设计研究所 | Method for testing asymmetric loading fatigue strength of fan blade in vacuum environment |
| CN115292925A (en) * | 2022-07-29 | 2022-11-04 | 中国航发沈阳发动机研究所 | Method for evaluating working blade of single crystal high-pressure turbine |
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| CN110595709A (en) * | 2019-08-14 | 2019-12-20 | 南京航空航天大学 | Method for determining allowable amplitude of turbine engine blade |
| CN110595709B (en) * | 2019-08-14 | 2020-07-17 | 南京航空航天大学 | Method for determining allowable amplitude of turbine engine blade |
| CN110988002A (en) * | 2019-11-27 | 2020-04-10 | 南京航空航天大学 | Rapid imaging method for microcracks of damaged section of foreign object based on image recognition |
| CN110987389A (en) * | 2019-11-27 | 2020-04-10 | 南京航空航天大学 | Foreign object damage notch calibration method with stress concentration coefficient not less than 3 |
| CN111046490A (en) * | 2019-11-27 | 2020-04-21 | 南京航空航天大学 | Grid size inversion method in foreign object damage gap analysis |
| CN111046490B (en) * | 2019-11-27 | 2024-07-02 | 南京航空航天大学 | Grid size inversion method in foreign object damage notch analysis |
| CN110889172A (en) * | 2019-12-04 | 2020-03-17 | 中国直升机设计研究所 | Method for prefabricating impact damage of helicopter rotor system |
| CN110884684A (en) * | 2019-12-04 | 2020-03-17 | 中国直升机设计研究所 | Design method for strength test of helicopter after impact of bearingless tail blade |
| CN110884684B (en) * | 2019-12-04 | 2022-04-12 | 中国直升机设计研究所 | Design method for strength test of helicopter after impact of bearingless tail blade |
| CN112881203A (en) * | 2021-01-15 | 2021-06-01 | 北京航空航天大学 | Preparation device and method for foreign object damage notch of blade leading edge simulation piece |
| CN112881203B (en) * | 2021-01-15 | 2022-03-15 | 北京航空航天大学 | Preparation device and method for foreign object damage notch of blade leading edge simulation piece |
| CN113075066A (en) * | 2021-03-22 | 2021-07-06 | 中国人民解放军空军工程大学 | Laser strengthening blade edge performance test method |
| CN113075066B (en) * | 2021-03-22 | 2023-10-31 | 中国人民解放军空军工程大学 | Laser reinforced blade edge performance test method |
| CN113378327B (en) * | 2021-07-02 | 2022-09-02 | 吉林重通成飞新材料股份公司 | Wind power blade trailing edge cracking resistance design method, device, equipment and storage medium |
| CN113378327A (en) * | 2021-07-02 | 2021-09-10 | 吉林重通成飞新材料股份公司 | Wind power blade trailing edge cracking resistance design method, device, equipment and storage medium |
| CN114171139B (en) * | 2021-10-20 | 2023-06-30 | 中国航发四川燃气涡轮研究院 | Material selecting method for compressor blade |
| CN114171139A (en) * | 2021-10-20 | 2022-03-11 | 中国航发四川燃气涡轮研究院 | Material selection method for compressor blade |
| CN114492113A (en) * | 2022-01-05 | 2022-05-13 | 南京航空航天大学 | Impact damage numerical simulation optimization method based on laser mapping entity grid |
| CN114492113B (en) * | 2022-01-05 | 2024-06-11 | 南京航空航天大学 | Impact damage numerical simulation optimization method based on laser mapping solid grids |
| CN114526920A (en) * | 2022-02-21 | 2022-05-24 | 中国航发贵阳发动机设计研究所 | Method for testing asymmetric loading fatigue strength of fan blade in vacuum environment |
| CN114526920B (en) * | 2022-02-21 | 2024-03-19 | 中国航发贵阳发动机设计研究所 | Method for testing asymmetric loading fatigue strength of fan blade in vacuum environment |
| CN115292925A (en) * | 2022-07-29 | 2022-11-04 | 中国航发沈阳发动机研究所 | Method for evaluating working blade of single crystal high-pressure turbine |
| CN118278113A (en) * | 2024-06-04 | 2024-07-02 | 南京航空航天大学 | CART-based aircraft structural component fatigue state detection method |
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