CN109815521B - A method for evaluating the anti-FOD capability of aero-engine blades - Google Patents
A method for evaluating the anti-FOD capability of aero-engine blades Download PDFInfo
- Publication number
- CN109815521B CN109815521B CN201811464884.0A CN201811464884A CN109815521B CN 109815521 B CN109815521 B CN 109815521B CN 201811464884 A CN201811464884 A CN 201811464884A CN 109815521 B CN109815521 B CN 109815521B
- Authority
- CN
- China
- Prior art keywords
- blade
- damage
- test
- stress
- impact
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
Images
Landscapes
- Investigating Strength Of Materials By Application Of Mechanical Stress (AREA)
- Testing Of Devices, Machine Parts, Or Other Structures Thereof (AREA)
Abstract
本发明涉及一种航空发动机叶片抗FOD能力的评估方法，该方法中对模拟叶片数值模型进行外物损伤的冲击动力学仿真，得到缺口宏观特征与外物类型、冲击速度、冲击角度之间的关系，并以此为数据依据确定外物损伤试验条件，利用空气炮对模拟叶片进行外模拟物损伤试验，观察冲击缺口的宏观微观特征；以叶片前缘危险位置工作载荷下的静应力与动应力作为高周疲劳试验的初始静载与动载，对损伤后的模拟叶片进行高周疲劳试验，通过步进法得到叶片的高周疲劳强度，根据高周疲劳试验结果评估叶片抗FOD能力；对少数的真实叶片进行模拟外物损伤及高周疲劳试验得到其高周疲劳强度，以验证模拟叶片与真实叶片试验结果的符合性。
The invention relates to a method for evaluating the anti-FOD capability of an aero-engine blade. In the method, the impact dynamics simulation of foreign object damage is performed on a simulated blade numerical model, and the relationship between the macroscopic characteristics of the notch and the type of the foreign object, the impact speed and the impact angle is obtained. Based on this, the external object damage test conditions were determined, and the simulated blade was subjected to an external simulated damage test using an air gun to observe the macroscopic and microscopic characteristics of the impact notch. Stress is used as the initial static load and dynamic load of the high-cycle fatigue test. The high-cycle fatigue test is performed on the damaged simulated blade, and the high-cycle fatigue strength of the blade is obtained by the step-by-step method, and the FOD resistance of the blade is evaluated according to the high-cycle fatigue test results. The high-cycle fatigue strength was obtained by simulating foreign object damage and high-cycle fatigue tests on a few real blades to verify the compliance of the simulated blades with the test results of the real blades.
Description
Technical Field
The invention relates to the field of aero-engine blade performance testing.
Background
In the taking-off and landing process of an airplane on a runway or an aircraft carrier deck, an aircraft engine running at a high rotating speed often sucks hard objects such as stones, gravels, bolts, metal fragments and the like, the relative speed of the foreign objects entering the engine and the blades of the air compressor can reach 100-350 m/s, and serious impact damage can be caused to the blades of the air compressor in the previous stages, particularly the front edges of the blades. In the prior art, the Damage caused by hard objects such as metal, sand and stone impacting the engine is called as Foreign Object Damage (FOD for short, Foreign Object Damage)
If the damaged blade is not found and treated in time, fatigue fracture failure may occur under high-frequency vibration, and the flying-off blade may break through a casing and even destroy the blades of the later stages of the compressor, thereby causing serious flight accidents. Therefore, the influence rule of foreign object damage on the fatigue strength of the blade material is explored, and the foreign object damage resistance of the blade is evaluated, so that the evaluation method is a part which is not ignored in the design of the blade of the aero-engine.
At present, the Foreign Object Damage (FOD) resistance of the blades of the aero-engine is evaluated at home mainly by referring to relevant regulations in standards such as American military computer MIL-HDBK-1783B and the like, and the minimum fatigue notch coefficient K equivalent to the minimum is evaluated on the blades after the foreign objects are sucked into the blades of the aero-enginefIf the engine is operating for a damage of 3, the engine can be operated for two checking periods or hours specified by the specifications. However, the root cause of the blade fracture after FOD is that the damage part formed by impact has micro-cracks, stress concentration, residual stress and microstructure damage, and the damage part is easy to become a fatigue source to generate fatigue fracture under high-cycle working load. And the high cycle fatigue failure of metal has larger dispersity, the capability of resisting foreign object damage of the blade is examined only through the number of cycles which can work after damage, and the risk of blade failure caused by foreign object damage cannot be effectively avoided.
Disclosure of Invention
The purpose of the invention is as follows: according to the invention, the high cycle fatigue performance of the damaged blade is obtained through an air cannon simulation foreign object damage test and a high cycle fatigue test of the blade, so that the FOD resistance of the blade is evaluated.
The technical scheme is as follows:
an assessment method for FOD resistance of an aircraft engine blade is characterized by comprising the following steps:
(1) establishing a numerical model of the tested real blade, and selecting a certain point at the front edge position of the real blade in the numerical model as a dynamic stress testA point position; extracting the modal stress sigma of the positiona1(ii) a Extracting modal stress sigma of dynamic stress test point position01And the true stress sigma of the dynamic stress test point position0(ii) a Calculating the vibration stress sigma under the working load of the position of the measuring pointa：
Meanwhile, obtaining the static stress result sigma of the measuring point position in the numerical model of the real bladem；
(2) Designing and manufacturing a blade leading edge simulation test piece according to the leading edge radius R of the blade leading edge measuring point position selected in the step (1) and the angle theta formed by the leading edge, and establishing a numerical model of the blade leading edge simulation test piece;
(3) simulating the process that foreign objects with different materials and different sizes impact the numerical model of the blade leading edge simulation test piece at different impact speeds and impact angles in dynamics analysis software to obtain a simulation result of the relationship between the macroscopic characteristics of the damage notch and the types, impact speeds and impact angles of the foreign objects;
(4) determining a damage severity index needing to be assessed according to damage data statistics of an aeroengine outfield blade or a related maintenance criterion boundary size of the engine blade, selecting a foreign object type, an impact speed and an impact angle of a simulated foreign object damage test which accords with the assessed damage severity index in the step (3) according to the assessed damage severity index, and then launching the foreign object by using an air cannon to impact a blade front edge simulation test piece to damage the blade front edge simulation test piece;
(5) carrying out high-cycle fatigue test on the damaged simulated blade by high-cycle fatigue test equipment; the static stress sigma under the working load of the dangerous position of the leading edge calculated in the step (1) is usedmVibration stress sigma under working load at dangerous position as static load of high cycle fatigue testaPerforming a high cycle fatigue test of a design life N cycle on the damaged simulation blade as a high cycle fatigue test dynamic load; n is a positive integer;
(6) and evaluating the FOD resistance of the blade: such as similar damage degreeThe fatigue fracture of a plurality of simulated blades in the N cycle indicates that the fatigue strength of the damaged blades at the corresponding service life can not meet the safety requirement of the working load of the engine; if no fatigue fracture occurs in the N cycles, the test is performed in each step by a step method for N cycles with the dynamic load σ at that timeaAnd (4) for initial dynamic load, increasing the dynamic load of each step compared with the dynamic load of the previous step until fatigue fracture occurs, and obtaining the fatigue strength of the damaged test piece under the specified service life.
Further, the method also comprises the step (7) of verifying a real blade test: carrying out a foreign object damage test on the measuring point positions of the front edges of the real blades to obtain impact damage; and carrying out high-cycle fatigue test on the damaged real blade to obtain the fatigue strength of the damaged real blade so as to verify the conformity of the simulated blade fatigue strength result and the real blade result.
Further, in the step (5), N is more than or equal to 3 multiplied by 107。
Further, in the step (1), a numerical model of the tested real blade is established by utilizing engineering modeling software, modal analysis is carried out on the finite element model of the tested real blade in finite element analysis software, the rotating speed and boundary conditions which are the same as the working load are applied to the finite element model of the blade, the dynamic frequency and vibration stress distribution of the blade is calculated, the position with the maximum stress of the front edge of the blade is extracted, and the position with the maximum vibration stress is taken as the position of a dynamic stress test point to evaluate the foreign object damage resistance of the blade.
Further, in the step (4), the damage severity index includes damage type, size and stress concentration degree.
Has the advantages that: according to the method, the high cycle fatigue test is carried out on the damaged simulated blade, the high cycle fatigue strength of the blade is obtained through a stepping method, and the FOD resistance of the blade can be more accurately evaluated according to the high cycle fatigue test result, so that the risk of blade failure caused by foreign object damage can be effectively avoided.
Drawings
FIG. 1 is a flow chart of the method for evaluating the FOD resistance of an aircraft engine blade according to the invention.
FIG. 2 is a block diagram of a blade leading edge simulation test piece.
FIG. 3 is a sectional view taken along the line A-A of the leading edge simulation test piece of the blade of FIG. 2.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings.
As shown in FIG. 1, the invention discloses a method for evaluating the FOD resistance of an aircraft engine blade, which comprises the following steps:
an assessment method for FOD resistance of an aircraft engine blade comprises the following steps:
(1) analyzing the vibration stress and the static strength of the blade: and (3) establishing a numerical model of the blade by utilizing UG (Unigraphics) or other engineering modeling software, carrying out modal analysis on a real blade finite element model in ANSYS or other finite element analysis software, applying the rotating speed and boundary conditions which are the same as the working load to the blade finite element model, and calculating the dynamic frequency and vibration stress distribution of the blade. And extracting the position with the maximum stress at the front edge of the blade, wherein the position with the maximum vibration stress is most dangerous to damage if a foreign hard object is sucked in during the operation of the engine, and taking the position as a test point position to evaluate the foreign object damage resistance of the blade. Extracting the modal stress sigma of the position of the test pointa1(ii) a Extracting modal stress sigma of dynamic stress test point position01And the true stress sigma of the dynamic stress test point position0(ii) a Calculating the vibration stress sigma under the working load of the position of the measuring pointa：
Meanwhile, static strength calculation is carried out on the finite element model of the real blade in ANSYS or other finite element analysis software, the rotating speed and boundary conditions which are the same as the working load are applied to the model, and static stress distribution of the blade is calculated. Extracting static stress result sigma under working load of dangerous positionm。
(2) Designing and manufacturing a blade leading edge simulation test piece according to the leading edge radius R of the blade leading edge measuring point position selected in the step (1) and the angle theta formed by the leading edge, and establishing a numerical model of the blade leading edge simulation test piece, as shown in FIGS. 2 and 3; the blade is suitable for high cycle tensile fatigue tests, and for other types of high cycle fatigue tests, other test pieces which can represent the geometrical characteristics of the leading edge and are suitable for the high cycle fatigue tests need to be designed according to relevant test criteria.
(3) And numerical simulation analysis of foreign object damage: and (3) establishing a numerical model of the blade at the position of the testing point at the front edge of the blade designed in the step (1) by utilizing UG or other engineering modeling software, and simulating the process that foreign objects with different materials and different sizes impact the simulated blade at different impact speeds and impact angles in LS-DYNA or other dynamics analysis software to obtain a simulation result of the relationship between the macroscopic characteristics of the damage gap and the types, impact speeds and impact angles of the foreign objects. And data basis is provided for a foreign object damage simulation test for simulating the blade.
(4) Determining a damage severity index needing to be assessed according to damage data statistics of an aeroengine outfield blade or a related maintenance criterion boundary size of the engine blade, selecting a foreign object type, an impact speed and an impact angle of a simulated foreign object damage test which accords with the assessed damage severity index in the step (3) according to the assessed damage severity index, and then launching the foreign object by using an air cannon to impact a blade front edge simulation test piece to damage the blade front edge simulation test piece; the damage severity index includes damage type, size, stress concentration degree.
(5) Carrying out high-cycle fatigue test on the damaged simulated blade by high-cycle fatigue test equipment; the static stress sigma under the working load of the dangerous position of the leading edge calculated in the step (1) is usedmVibration stress sigma under working load at dangerous position as static load of high cycle fatigue testaPerforming a high cycle fatigue test of a design life N cycle on the damaged simulation blade as a high cycle fatigue test dynamic load; n is a positive integer; titanium alloy blades generally require N to be 3 × 107At present, the requirement of few standards is that N is 109。
(6) And evaluating the FOD resistance of the blade: if fatigue fracture occurs in N cycles of a plurality of simulated blades with similar damage degrees, the damaged blades with the damage degrees are indicatedThe fatigue strength under the corresponding service life can not meet the safety requirement of the working load of the engine; if no fatigue fracture occurs in the N cycles, the test is performed in each step by a step method for N cycles with the dynamic load σ at that timeaAnd (4) for initial dynamic load, increasing the dynamic load of each step compared with the dynamic load of the previous step until fatigue fracture occurs, and obtaining the fatigue strength of the damaged test piece under the specified service life.
Step (7), verifying a real blade test: carrying out a foreign object damage test on the measuring point positions of the front edges of the real blades to obtain impact damage; and carrying out high-cycle fatigue test on the damaged real blade to obtain the fatigue strength of the damaged real blade so as to verify the conformity of the simulated blade fatigue strength result and the real blade result.
Claims (4)
1. An assessment method for FOD resistance of an aircraft engine blade is characterized by comprising the following steps:
(1) establishing a numerical model of the tested real blade, and selecting a certain point of the front edge position of the real blade in the numerical model as a dynamic stress test point position; extracting the modal stress sigma of the positiona1(ii) a Extracting modal stress sigma of dynamic stress test point position01And the true stress sigma of the dynamic stress test point position0(ii) a Calculating the vibration stress sigma under the working load of the position of the measuring pointa：
Meanwhile, the static stress sigma of the measuring point position is obtained in the numerical model of the real bladem(ii) a The position of the measuring point is selected as the position with the maximum vibration stress, namely the position with the maximum vibration stress is damaged most as a danger, so the position of the measuring point is taken as a front edge danger position;
in the step, a numerical model of a tested real blade is established by utilizing engineering modeling software, modal analysis is carried out on the finite element model of the tested real blade in finite element analysis software, the rotating speed and boundary conditions which are the same as the working load are applied to the finite element model of the blade, the dynamic frequency and the vibration stress distribution of the blade are calculated, the maximum position of the stress of the front edge of the blade is extracted, and the position is taken as the position of a dynamic stress test point to evaluate the foreign object damage resistance of the blade as the position of the maximum vibration stress of the position is selected;
(2) designing and manufacturing a blade leading edge simulation test piece according to the leading edge radius R of the blade leading edge measuring point position selected in the step (1) and the angle theta formed by the leading edge, and establishing a numerical model of the blade leading edge simulation test piece;
(3) simulating the process that foreign objects with different materials and different sizes impact the numerical model of the blade leading edge simulation test piece at different impact speeds and impact angles in dynamics analysis software to obtain a simulation result of the relationship between the macroscopic characteristics of the damage notch and the types, impact speeds and impact angles of the foreign objects;
(4) determining a damage severity index needing to be assessed according to damage data statistics of an aeroengine outfield blade or a related maintenance criterion boundary size of the engine blade, selecting a foreign object type, an impact speed and an impact angle of a simulated foreign object damage test which accords with the assessed damage severity index in the step (3) according to the assessed damage severity index, and then launching the foreign object by using an air cannon to impact a blade front edge simulation test piece to damage the blade front edge simulation test piece;
(5) carrying out high-cycle fatigue test on the damaged simulated blade by high-cycle fatigue test equipment; the static stress sigma under the working load of the dangerous position of the leading edge calculated in the step (1) is usedmVibration stress sigma under working load at dangerous position as static load of high cycle fatigue testaPerforming a high cycle fatigue test of a design life N cycle on the damaged simulation blade as a high cycle fatigue test dynamic load; n is a positive integer;
(6) and evaluating the FOD resistance of the blade: if a plurality of simulation blades with similar damage degrees are subjected to fatigue fracture in N cycles, the fatigue strength of the damaged blades in the corresponding service life cannot meet the safety requirement of the working load of the engine; if no fatigue fracture occurs in the N cycles, the test is performed in each step by a step method for N cycles with the dynamic load σ at that timeaFor initial dynamic load, the dynamic load of each step is higher than that of the last stepAnd increasing until fatigue fracture occurs to obtain the fatigue strength of the damaged test piece under the specified service life.
2. The method for evaluating the FOD resistance of an aircraft engine blade according to claim 1, wherein: the method also comprises the following steps of (7) and true leaf test verification: carrying out a foreign object damage test on the measuring point positions of the front edges of the real blades to obtain impact damage; and carrying out high-cycle fatigue test on the damaged real blade to obtain the fatigue strength of the damaged real blade so as to verify the conformity of the simulated blade fatigue strength result and the real blade result.
3. The method for evaluating the FOD resistance of an aircraft engine blade according to claim 1, wherein: in step (5), N is not less than 3X 107。
4. The method for evaluating the FOD resistance of an aircraft engine blade according to claim 1, wherein: in the step (4), the damage severity index includes damage type, size, and stress concentration degree.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811464884.0A CN109815521B (en) | 2018-12-03 | 2018-12-03 | A method for evaluating the anti-FOD capability of aero-engine blades |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811464884.0A CN109815521B (en) | 2018-12-03 | 2018-12-03 | A method for evaluating the anti-FOD capability of aero-engine blades |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN109815521A CN109815521A (en) | 2019-05-28 |
| CN109815521B true CN109815521B (en) | 2021-04-16 |
Family
ID=66601944
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201811464884.0A Active CN109815521B (en) | 2018-12-03 | 2018-12-03 | A method for evaluating the anti-FOD capability of aero-engine blades |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN109815521B (en) |
Families Citing this family (21)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110595709B (en) * | 2019-08-14 | 2020-07-17 | 南京航空航天大学 | A Method for Determining Allowable Amplitude of Turbine Engine Blades |
| 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 |
| CN110987389B (en) * | 2019-11-27 | 2021-07-27 | 南京航空航天大学 | A foreign object damage notch calibration method with a stress concentration factor of not less than 3 |
| 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 |
| CN110889172B (en) * | 2019-12-04 | 2023-03-28 | 中国直升机设计研究所 | Method for prefabricating impact damage of helicopter rotor system |
| 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 |
| CN112881203B (en) * | 2021-01-15 | 2022-03-15 | 北京航空航天大学 | A preparation device and method for a foreign object damage notch in a blade leading edge simulation part |
| CN113075066B (en) * | 2021-03-22 | 2023-10-31 | 中国人民解放军空军工程大学 | Laser enhanced blade edge performance testing 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 |
| CN113962021B (en) * | 2021-09-28 | 2025-12-12 | 中国人民解放军空军工程大学 | A test specimen for simulating the stress gradient of compressor blades and its optimization method |
| 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 |
| CN114492113B (en) * | 2022-01-05 | 2024-06-11 | 南京航空航天大学 | A numerical simulation optimization method for impact damage based on laser mapping solid mesh |
| 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 |
| CN116735219A (en) * | 2022-03-04 | 2023-09-12 | 中国航发商用航空发动机有限责任公司 | Test method for burn-through strength of aeroengine combustion chamber casing |
| CN115146393A (en) * | 2022-07-26 | 2022-10-04 | 中国航发沈阳发动机研究所 | A kind of three-dimensional machining foreign object damage notch morphology and its size determination method |
| CN115266119B (en) * | 2022-07-27 | 2025-04-04 | 中国航发沈阳发动机研究所 | A method for determining the test procedure of foreign object damage assessment for aircraft engines |
| CN115292925B (en) * | 2022-07-29 | 2023-07-07 | 中国航发沈阳发动机研究所 | Method for evaluating working blade of single-crystal high-pressure turbine |
| CN118278113B (en) * | 2024-06-04 | 2024-09-27 | 南京航空航天大学 | A CART-based fatigue state detection method for aircraft structural components |
| CN119804178A (en) * | 2024-12-11 | 2025-04-11 | 南京航空航天大学 | A blade damage simulation method based on coupling hard object damage and corrosion test |
| CN119808467B (en) * | 2024-12-11 | 2025-12-05 | 南京航空航天大学 | A fatigue failure analysis method for aero-engine blades |
| CN119808465B (en) * | 2024-12-11 | 2025-12-05 | 南京航空航天大学 | A method, system, device, and storage medium for classifying vulnerable areas of aero-engine blades. |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH0861086A (en) * | 1994-08-22 | 1996-03-05 | Toshiba Corp | Evaluation method for foreign matter collision damage on gas turbine blades |
-
2018
- 2018-12-03 CN CN201811464884.0A patent/CN109815521B/en active Active
Non-Patent Citations (4)
| Title |
|---|
| 刘超.叶片外物损伤模拟及其疲劳强度预测技术研究.《中国优秀硕士学位论文全文数据库工程科技II辑》.2015,(第02期),第C031-245页. * |
| 孙护国 等.前缘半径对钛合金叶片抗外物损伤能力影响的数值分析.《航空发动机》.2016,第42卷(第02期),第1-6页. * |
| 张雪强.钛合金叶片外物损伤试验与数值模拟研究.《中国优秀硕士学位论文全文数据库工程科技II辑》.2015,(第02期),第C031-431页. * |
| 葛宁.发动机叶片抗外物损伤能力评估技术研究.《中国优秀硕士学位论文全文数据库工程科技II辑》.2015,(第02期),第C031-418页. * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN109815521A (en) | 2019-05-28 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN109815521B (en) | A method for evaluating the anti-FOD capability of aero-engine blades | |
| CN112733398B (en) | A method for determining the repair-free limit of impact damage of pit-shaped hard objects | |
| CN111046490B (en) | Grid size inversion method in foreign object damage notch analysis | |
| CN110134990B (en) | Detection method for dangerous part of front edge of blade of aircraft engine | |
| Witek | Numerical stress and crack initiation analysis of the compressor blades after foreign object damage subjected to high-cycle fatigue | |
| CN110990948A (en) | Method for predicting damage fatigue strength of foreign object of blade of aircraft engine | |
| CN107084844A (en) | A kind of test method of blade construction simulating piece | |
| Poursaeidi et al. | Failure analysis of an axial compressor first row rotating blades | |
| CN109583147B (en) | Design method of centrifugal impeller pre-rotation disk center gradient simulation piece | |
| Witek | Crack growth simulation in the compressor blade subjected to vibration using boundary element method | |
| CN110987388B (en) | A Method for Equivalent Machining Notches Based on Notched Fatigue Strength | |
| Jia et al. | Foreign object damage characteristics and their effects on high cycle fatigue property of Ni‐based superalloy GH4169 | |
| Hanschke et al. | The effect of foreign object damage on compressor blade high cycle fatigue strength | |
| CN111859729B (en) | Method for calculating service life of wheel disc by considering shot blasting model with multiple shot randomly distributed | |
| Pedram et al. | Total life estimation of a compressor blade with corrosion pitting, SCC and fatigue cracking | |
| Schrade et al. | Experimental and numerical investigation of erosive change of shape for high-pressure compressors | |
| CN116296419A (en) | Method for verifying bird-swallowing navigable compliance of aviation turbofan engine | |
| CN111238804B (en) | Test piece configuration method for over-rotation rupture test of aircraft engine rotor | |
| CN115270467A (en) | Method for determining bird swallowing capacity of aero-engine | |
| CN115391929B (en) | Method for evaluating damage resistance of aero-engine fan or compressor blade to foreign objects | |
| CN112541230B (en) | A method for analyzing the containment of casing with blade offset loss | |
| CN119808465B (en) | A method, system, device, and storage medium for classifying vulnerable areas of aero-engine blades. | |
| Zhang et al. | Effect of bird strike on residual strength and vibration of fan blades for aero-engines | |
| CN120217762A (en) | A method for verifying the repair-free limit of leading and trailing edge notch-type foreign object damage based on step-by-step loading | |
| CN111850442B (en) | Strengthening method for preventing high-order vibration type induced blade tip block dropping of titanium alloy blisk blade |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |