WO2020187038A1 - 一种fod用弯曲疲劳试验件及其设计方法 - Google Patents

一种fod用弯曲疲劳试验件及其设计方法 Download PDF

Info

Publication number
WO2020187038A1
WO2020187038A1 PCT/CN2020/077925 CN2020077925W WO2020187038A1 WO 2020187038 A1 WO2020187038 A1 WO 2020187038A1 CN 2020077925 W CN2020077925 W CN 2020077925W WO 2020187038 A1 WO2020187038 A1 WO 2020187038A1
Authority
WO
WIPO (PCT)
Prior art keywords
blade
leading edge
width
test piece
simulated
Prior art date
Application number
PCT/CN2020/077925
Other languages
English (en)
French (fr)
Inventor
赵振华
王凌峰
陈伟
罗刚
刘璐璐
Original Assignee
南京航空航天大学
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by 南京航空航天大学 filed Critical 南京航空航天大学
Publication of WO2020187038A1 publication Critical patent/WO2020187038A1/zh

Links

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
    • G01M13/00—Testing of machine parts
    • 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
    • G01M7/08—Shock-testing

Definitions

  • the invention relates to the technical field of simulation test of aircraft engine blades.
  • the test piece is designed as a flat plate as a whole, and it is divided into two types without leading edge and with leading edge.
  • the former is used for materials.
  • the latter can show the influence of the front edge characteristics on the fatigue performance of the test piece, but due to the shape of the flat plate, the tensile testing machine is used for tensile and compression high-cycle fatigue tests.
  • High cycle fatigue test is an important research content to reveal the fatigue performance of FOD blades. Tension-compression high-cycle fatigue test can only study the effect of centrifugal load on the fatigue strength of engine blades, and cannot show the effect of vibration load.
  • Engine blades are subjected to the combined action of centrifugal load and vibration load in actual work. Therefore, it is necessary to carry out research on the bending fatigue performance of engine blades. It is necessary to invent test pieces that can meet the requirements of bending fatigue test, that is, simulation blades are required to perform The shape of the leading edge of the real blade feature position requires the simulated blade to avoid the problem of insufficient stress during bending vibration.
  • FOD test and high-cycle bending fatigue test using real engine blades are the most accurate research methods for this type of blade, but the foreign object damage test is a destructive test. After the test piece is damaged, it cannot be restored to the state before the test. A real blade test piece needs to be scrapped for one test, and the test cost is extremely high. This makes it impossible to carry out large-scale FOD tests and high-cycle bending fatigue tests, and the amount of test data obtained is limited, which cannot meet the needs of fatigue performance research.
  • the present invention provides a bending fatigue test piece for FOD, which can simulate the characteristics of the real blade leading edge and effectively avoid the problem of low bending vibration stress.
  • the invention also provides a design method of the above-mentioned bending fatigue test piece for FOD.
  • a bending fatigue test piece for FOD comprising a clamping end and a simulation blade, the rear end of the simulation blade is connected with the clamping end;
  • the simulation blade has an upper surface, a lower surface and two side surfaces; and an upper surface, a lower surface and There is a circular arc transition at the junction of the two sides and the clamping end;
  • the front end of the simulated blade is a wing leading edge structure, and the wing leading edge structure includes a flat plate section and two curved and warped ends of the plate section
  • the front edge, and the connection between the front edge and the flat plate section is a circular arc transition;
  • the two front edges are symmetrically arranged on both sides of the flat plate section.
  • the thickness of the clamping end is greater than six times the thickness of the flat plate section.
  • the length of the simulated blade is set to 1.5 times the width of the simulated blade.
  • the width of the slab section of the simulated blade is three times the width of the leading edge.
  • the bending fatigue test piece for FOD provided by the present invention can simulate the characteristics of the real blade leading edge and effectively avoid the problem of low bending vibration stress.
  • the clamping end can be effectively clamped in the FOD simulation test and bending fatigue test, so as to prevent the test piece from being loosened to affect the test results;
  • the simulated blade can simulate the shape of the leading edge of the real blade feature position, and the leading edge of the blade can be appropriately bent. Effectively avoid the problem of small bending stress caused by bending the neutral layer.
  • the width of the slab section of the simulated blade is three times the width of the leading edge
  • the bending angle of the leading edge of the simulated blade is set to 20°;
  • transition radius of the simulated blade flat section is equal to the width of the leading edge, and the corners are rounded during processing
  • the length of the simulated blade is set to 1.5 times the width of the simulated blade
  • the draft angle of the transition section is set to 5°, and the length of the transition section is set to one third of the length of the simulated blade;
  • the size of the clamping end is determined by the contour size of the simulated blade, and the thickness of the clamping end is greater than six times the thickness of the plate section.
  • Figure 1 is the front view of the test piece.
  • Figure 2 is the left side view of the test piece.
  • Fig. 3 is a cross-sectional view in the direction of A-A in Fig. 1.
  • the present invention discloses a bending fatigue test piece for FOD, including a clamping end 200 and a simulated blade 100, the rear end of the simulated blade 100 is connected with the clamping end 200; It has an upper surface, a lower surface and two side surfaces; and the connection between the upper surface, the lower surface and the two side surfaces and the clamping end has an arc transition; the front end of the simulated blade is the wing leading edge structure, and the front of the wing
  • the edge structure includes a flat plate section 101 and two front edges 102 bent and warped from both ends of the flat plate section 101, and the connection between the front edge 102 and the flat plate section 101 is a circular arc transition; the two front edges 102 are on both sides of the flat plate section Set symmetrically.
  • only one clamping end 200 is provided and has a rectangular parallelepiped structure, which is the same material as the simulated blade.
  • a circular arc structure is used to smoothly transition between the clamping end 200 and the simulated blade 100 to avoid stress concentration.
  • the simulated blade part adopts the design of the leading edge of the wing, and the size parameters of the leading edge are the same as the characteristic position of the real blade.
  • the front edge is bent as a whole, so that the simulated blade is not a simple flat plate, but a structural shape of a middle flat plate that is warped on both sides, which can effectively avoid the problem of small bending vibration stress due to the principle of bending neutral layer.
  • the test piece adopts the overall cutting processing method, and the surface is polished after processing to eliminate micro-cracks and burrs
  • the present invention also provides a design method of the above-mentioned bending fatigue test piece for FOD:
  • the width of the slab section of the simulated blade is three times the width of the leading edge
  • the bending angle 11 of the leading edge of the simulated blade is set to 20°, which can effectively improve the bending vibration stress of the leading edge;
  • transition radius 12 of the simulated blade flat section is equal to the width of the leading edge, and the corners are rounded during processing
  • the simulated blade length 8 is set to 1.5 times the simulated blade width
  • the draft angle 5 of the transition section is set to 5°, and the length of the transition section 4 is set to one-third of the length of the simulated blade 8.
  • the excessive fillet 6 has the effect of eliminating stress concentration and can be set as required;
  • the size of the clamping end is determined by the contour size of the simulated blade, and the clamping end thickness 3 is greater than six times the thickness 13 of the plate section to provide space for the fixture to be positioned.
  • the bending fatigue test piece for FOD can meet the needs of FOD test and bending fatigue test, broaden the research field of FOD blade fatigue performance, which is of great significance for engineering applications, and the research on the fatigue life prediction of real blades after FOD Has the value of promotion.
  • the size of the bending fatigue test piece for FOD is determined by the size of the dangerous section of the real blade, which is generally not very large, convenient for processing, short production cycle of the test piece, low production cost, and less materials;
  • the test of bending fatigue test piece for FOD requires a small site, low test cost and short test period.

Landscapes

  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Investigating Strength Of Materials By Application Of Mechanical Stress (AREA)

Abstract

本发明公开了一种FOD用弯曲疲劳试验件及其设计方法，该FOD用弯曲疲劳试验件包括夹持端及模拟叶片。其中，夹持端能满足在FOD模拟试验和弯曲疲劳试验时有效夹持，避免试验件因松动影响试验结果；模拟叶片能模拟真实叶片特征位置前缘的形状，并把叶片前缘适当弯曲，有效避免因弯曲中性层导致的弯曲应力小问题。

Description

一种FOD用弯曲疲劳试验件及其设计方法 技术领域
本发明涉及飞行器发动机叶片模拟试验技术领域。
背景技术
飞机在跑道或航母甲板起降过程中，高转速运行的航空发动机常会吸入石子、砂砾、螺栓、金属碎片等硬物，这些异物进入发动机时与压气机叶片的相对速度可达100m/s-350m/s，会对前几级压气机叶片尤其是叶片的前缘造成严重的冲击损伤损伤。现有技术中，将金属、砂石等坚硬物体撞击发动机带来的损伤称作“外物损伤”(简称FOD，Foreign Object Damage)。
现有技术中，用于FOD试验与高周疲劳试验的平板模拟叶片试验件几乎均存在如下问题：试验件整体为平板设计，分为不带前缘与带前缘两类，前者用于材料的疲劳性能试验，后者可表现前缘特征对试验件疲劳性能的影响，但由于其平板的形状，都是使用拉伸试验机进行拉压高周疲劳试验。高周疲劳试验是揭示FOD叶片疲劳性能的重要研究内容。拉压高周疲劳试验只能研究离心载荷对发动机叶片疲劳强度的影响，无法表现振动载荷的影响。发动机叶片在实际工作中受到离心载荷与振动载荷的复合作用，因此需要开展对发动机叶片弯曲疲劳性能的研究，这就必须要发明能够满足弯曲疲劳试验要求的试验件，即要求模拟叶片既能表现真实叶片特征位置前缘的形状，又要求模拟叶片能避免弯曲振动时应力不足的问题。
使用真实发动机叶片进行FOD试验及高周弯曲疲劳试验是针对该型号叶片最为准确的研究方法，但外物损伤试验是破坏性试验，试验件受损后将无法恢复到试验前状态，因此每进行一次试验就需要报废一件真实叶片试验件，试验成本极其高昂。这使得无法进行大规模FOD试验及高周弯曲疲劳试验，获得的试验数据量有限，无法满足疲劳性能研究所需。
发明内容
发明目的：本发明提供了一种FOD用弯曲疲劳试验件，能够模拟真实叶片前缘特征并有效避免弯曲振动应力低问题。
本发明同时提供了上述FOD用弯曲疲劳试验件的设计方法。
技术方案：为实现上述发明目的，本发明采用如下技术方案。
一种FOD用弯曲疲劳试验件，包括夹持端及模拟叶片，模拟叶片的后端与夹持端连接；所述模拟叶片具有上表面、下表面及两个侧面；且上表面、下表面及两个侧面与夹持端连接处具有圆弧过渡；所述模拟叶片的前端为机翼前缘结构，且该机翼前缘结构包括平板段及自平板段两端弯折翘曲的两个前缘，且前缘与平板段的连接处为圆弧过渡；两个前缘在平板段两侧呈对称设置。
进一步的，所述夹持端仅设置一个。
进一步的，夹持端的厚度大于六倍平板段厚度。
进一步的，模拟叶片长度置为模拟叶片宽度的1.5倍。
进一步的，模拟叶片平板段宽度为前缘宽度的三倍。
有益效果：本发明提供的FOD用弯曲疲劳试验件能模拟真实叶片前缘特征并有效避免弯曲振动应力低问题。其中，夹持端能满足在FOD模拟试验和弯曲疲劳试验时有效夹持，避免试验件因松动影响试验结果；模拟叶片能模拟真实叶片特征位置前缘的形状，并把叶片前缘适当弯曲，有效避免因弯曲中性层导致的弯曲应力小问题。
本发明提供的上述FOD用弯曲疲劳试验件的设计方法，可采用以下技术方案：
根据真实叶片的三维模型进行有限元仿真，得到前缘的危险位置，从模型中获得危险位置的截面尺寸，从中提取前缘半径和前缘夹角，叶身厚度稳定时的尺寸作为模拟叶片平板段厚度，以及从前缘到叶身厚度稳定位置的长度，作为前缘背面宽度和前缘正面宽度；
模拟叶片平板段宽度为前缘宽度的三倍；
模拟叶片前缘弯曲角设置为20°；
模拟叶片平板段过度半径与前缘宽度数值相等，加工时倒圆角；
模拟叶片长度设置为模拟叶片宽度的1.5倍；
过渡段拔模角设置为5°，过渡段长度设置为模拟叶片长度的三分之一；
夹持端尺寸由模拟叶片轮廓尺寸决定，夹持端厚度大于六倍平板段厚度。
附图说明
图1为试验件主视图。
图2为试验件左视图。
图3为图1中A-A方向剖视图。
其中标号的说明如下：
1—夹持端宽度；2—夹持端长度；3—夹持端厚度；4—过渡段长度；5—过渡段拔模角；6—过度圆角；7—模拟叶片宽度；8—模拟叶片长度；9—模拟叶片 前缘半径；10—模拟叶片前缘夹角；11—模拟叶片前缘弯曲角；12—模拟叶片平板段过度半径；13—模拟叶片平板段厚度；14—模拟叶片平板段背面宽度；15—模拟叶片平板段正面宽度；16—模拟叶片前缘背面宽度；17—模拟叶片前缘正面宽度；100—模拟叶片；200—夹持端；101—平板段；102—前缘。
具体实施方式
下面结合附图对本发明作进一步详细说明。
请参阅图1至图3所示，本发明公开了一种FOD用弯曲疲劳试验件，包括夹持端200及模拟叶片100，模拟叶片100的后端与夹持端200连接；所述模拟叶片具有上表面、下表面及两个侧面；且上表面、下表面及两个侧面与夹持端连接处具有圆弧过渡；所述模拟叶片的前端为机翼前缘结构，且该机翼前缘结构包括平板段101及自平板段101两端弯折翘曲的两个前缘102，且前缘102与平板段101的连接处为圆弧过渡；两个前缘102在平板段两侧呈对称设置。
本实施方式中，夹持端200仅设置一个，且为长方体结构，与模拟叶片材料相同，夹持端200与模拟叶片100之间用圆弧结构平滑过渡，避免应力集中。模拟叶片部分采用机翼前缘设计，前缘尺寸参数与真实叶片特征位置处相同。对前缘部分整体作弯曲处理，使得模拟叶片不是简单的平板型，而是两边翘曲中间平板的结构形状，能够有效避免由于弯曲中性层原理而导致弯曲振动应力小的问题。试验件采用整体切削加工方式，加工完成后对表面进行抛光处理，消除微裂纹与毛刺
而本发明同时提供上述FOD用弯曲疲劳试验件的设计方法：
根据真实叶片的三维模型进行有限元仿真，得到前缘的危险位置，从模型中获得危险位置的截面尺寸，从中提取前缘半径9和前缘夹角10，叶身厚度稳定时的尺寸作为模拟叶片平板段厚度13，以及从前缘到叶身厚度稳定位置的长度，作为前缘背面宽度16和前缘正面宽度17；
模拟叶片平板段宽度为前缘宽度的三倍；
模拟叶片前缘弯曲角11设置为20°，可以有效提高前缘弯曲振动应力；
模拟叶片平板段过度半径12与前缘宽度数值相等，加工时倒圆角；
模拟叶片长度8设置为模拟叶片宽度的1.5倍；
过渡段拔模角5设置为5°，过渡段长度4设置为模拟叶片长度8的三分之一，过度圆角6起消除应力集中作用，按需设置即可；
夹持端尺寸由模拟叶片轮廓尺寸决定，夹持端厚度3大于六倍平板段厚度13，以提供夹具定位的空间。
本发明的FOD用弯曲疲劳试验件设计方法的优点和积极效果主要体现在：
1、FOD用弯曲疲劳试验件可以满足FOD试验与弯曲疲劳试验的需求，拓宽了FOD叶片疲劳性能的研究领域，对于工程上的应用具有重要意义，对于真实叶片受FOD后的疲劳寿命预测的研究具有推动价值。
2、FOD用弯曲疲劳试验件尺寸由真实叶片危险截面的尺寸决定，一般不会很大，加工方便，试验件生产周期短、生产费用低、用料少；
3、FOD用弯曲疲劳试验件试验需要场地小，试验费用低，试验周期短。
以上所述仅是本发明的优选实施方式，应当指出在不脱离本发明的构思前提下，还可以做出若干推演或替代，这些推演或替代都应视为本发明的保护范围。

Claims (6)

  1. 一种FOD用弯曲疲劳试验件，其特征在于：包括夹持端及模拟叶片，模拟叶片的后端与夹持端连接；所述模拟叶片具有上表面、下表面及两个侧面；且上表面、下表面及两个侧面与夹持端连接处具有圆弧过渡；所述模拟叶片的前端为机翼前缘结构，且该机翼前缘结构包括平板段及自平板段两端弯折翘曲的两个前缘，且前缘与平板段的连接处为圆弧过渡；两个前缘在平板段两侧呈对称设置。
  2. 根据权利要求1所述的FOD用弯曲疲劳试验件，其特征在于：所述夹持端仅设置一个。
  3. 根据权利要求2所述的FOD用弯曲疲劳试验件，其特征在于：夹持端的厚度大于六倍平板段厚度。
  4. 根据权利要求1所述的FOD用弯曲疲劳试验件，其特征在于：模拟叶片长度置为模拟叶片宽度的1.5倍。
  5. 根据权利要求1所述的FOD用弯曲疲劳试验件，其特征在于：模拟叶片平板段宽度为前缘宽度的三倍。
  6. 一种对权利要求1所述的FOD用弯曲疲劳试验件的设计方法，其特征在于：
    根据真实叶片的三维模型进行有限元仿真，得到前缘的危险位置，从模型中获得危险位置的截面尺寸，从中提取前缘半径(9)和前缘夹角(10)，叶身厚度稳定时的尺寸作为模拟叶片平板段厚度(13)，以及从前缘到叶身厚度稳定位置的长度，作为前缘背面宽度(16)和前缘正面宽度(17)；
    模拟叶片平板段宽度为前缘宽度的三倍；
    模拟叶片前缘弯曲角(11)设置为20°；
    模拟叶片平板段过度半径(12)与前缘宽度数值相等，加工时倒圆角；
    模拟叶片长度(8)设置为模拟叶片宽度的1.5倍；
    过渡段拔模角(5)设置为5°，过渡段长度(4)设置为模拟叶片长度(8)的三分之一；
    夹持端尺寸由模拟叶片轮廓尺寸决定，夹持端厚度(3)大于六倍平板段厚度(13)。
PCT/CN2020/077925 2019-03-15 2020-03-05 一种fod用弯曲疲劳试验件及其设计方法 WO2020187038A1 (zh)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CN201910197262.4 2019-03-15
CN201910197262.4A CN110006636A (zh) 2019-03-15 2019-03-15 一种fod用弯曲疲劳试验件及其设计方法

Publications (1)

Publication Number Publication Date
WO2020187038A1 true WO2020187038A1 (zh) 2020-09-24

Family

ID=67167156

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/CN2020/077925 WO2020187038A1 (zh) 2019-03-15 2020-03-05 一种fod用弯曲疲劳试验件及其设计方法

Country Status (2)

Country Link
CN (1) CN110006636A (zh)
WO (1) WO2020187038A1 (zh)

Families Citing this family (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110006636A (zh) * 2019-03-15 2019-07-12 南京航空航天大学 一种fod用弯曲疲劳试验件及其设计方法
CN113495002B (zh) * 2020-04-03 2023-01-06 中国航发商用航空发动机有限责任公司 航空发动机非全环风扇叶片的旋转冲击试验装置
CN111695286B (zh) * 2020-06-22 2023-10-20 中国航发沈阳发动机研究所 一种航空发动机涡轮叶片取样夹具设计方法
CN112197922B (zh) * 2020-08-25 2022-03-25 中国航发湖南动力机械研究所 一种涡轮叶片振动疲劳模拟件及其设计方法
CN112748008A (zh) * 2020-12-29 2021-05-04 中国航空工业集团公司西安飞机设计研究所 一种工字梁拉伸疲劳试验件及拉伸疲劳试验方法
CN112881197B (zh) * 2021-01-15 2022-05-03 北京航空航天大学 一种叶片前缘高周疲劳模拟件及设计方法

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20110243748A1 (en) * 2010-03-30 2011-10-06 Panagiota Tsifourdaris High pressure turbine blade airfoil profile
CN105067791A (zh) * 2015-08-06 2015-11-18 中国航空工业集团公司北京航空材料研究院 一种模拟高温合金超高周疲劳损伤的方法
CN110006636A (zh) * 2019-03-15 2019-07-12 南京航空航天大学 一种fod用弯曲疲劳试验件及其设计方法

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8051565B2 (en) * 2006-12-30 2011-11-08 General Electric Company Method for increasing fatigue notch capability of airfoils
CN104062104B (zh) * 2013-03-19 2016-12-28 徐可君 航空发动机压气机叶片疲劳循环试验装置
CN104062105B (zh) * 2013-03-19 2016-08-17 徐可君 航空发动机高压压气机一级盘疲劳循环试验装置
CN107084844B (zh) * 2017-04-19 2019-02-05 西北工业大学 一种叶片结构模拟件的试验方法
CN207019856U (zh) * 2017-06-20 2018-02-16 青岛苏试海测检测技术有限公司 一种用于发动机叶片疲劳试验装置
CN108491570B (zh) * 2018-02-08 2020-08-07 南京航空航天大学 一种spfdb三层板空心夹芯结构弯曲疲劳寿命预测方法
CN108693030B (zh) * 2018-04-10 2019-11-29 南京航空航天大学 Fod缺口型损伤的疲劳性能预测方法
CN108765565A (zh) * 2018-04-10 2018-11-06 南京航空航天大学 一种fod损伤的三维实体建模方法

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20110243748A1 (en) * 2010-03-30 2011-10-06 Panagiota Tsifourdaris High pressure turbine blade airfoil profile
CN105067791A (zh) * 2015-08-06 2015-11-18 中国航空工业集团公司北京航空材料研究院 一种模拟高温合金超高周疲劳损伤的方法
CN110006636A (zh) * 2019-03-15 2019-07-12 南京航空航天大学 一种fod用弯曲疲劳试验件及其设计方法

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
WANG, LINGFENG: "Research on Influence of Foreign Object Damage on Stainless Steel Blade High Cycle Fatigue Property", CHINESE MASTER'S THESES FULL-TEXT DATABASE, 1 March 2019 (2019-03-01), pages 1 - 85, XP055734453, ISSN: 1674-0246 *

Also Published As

Publication number Publication date
CN110006636A (zh) 2019-07-12

Similar Documents

Publication Publication Date Title
WO2020187038A1 (zh) 一种fod用弯曲疲劳试验件及其设计方法
CN108491570B (zh) 一种spfdb三层板空心夹芯结构弯曲疲劳寿命预测方法
JP6080174B2 (ja) ファンケーシングで観測される衝撃痕の検査方法
CN101446208A (zh) 涡轮机叶片
Kuroki et al. Application of linear friction welding technique to aircraft engine parts
CN109145335A (zh) 一种通过预旋转提高轮盘低循环疲劳寿命的方法
Witek Crack growth simulation in the compressor blade subjected to vibration using boundary element method
CN109374449B (zh) 一种考虑高、低周疲劳的叶片前后缘裂纹型硬物损伤可用极限确定方法
Giuliese et al. A cohesive zone model for three-dimensional fatigue debonding/delamination
Wang et al. Investigation on experimental load spectrum for high and low cycle combined fatigue test
CN110987388B (zh) 一种基于缺口疲劳强度的等效机械加工缺口的方法
CN113962021A (zh) 一种压气机模拟叶片应力梯度试验件及优化方法
US10401268B2 (en) Rig
CN103558072A (zh) 一种复杂结构预制裂纹加工方法
CN114139276A (zh) 一种盘轴一体式整体叶盘结构疲劳寿命分析方法
CN112733398A (zh) 一种凹坑型硬物冲击损伤免修极限确定方法
CN208395275U (zh) 一种用于压气机叶片激光修复过程中夹持固定的冷却装置
CN208778045U (zh) 一种自带冠叶片带状阻尼拉筋条减振结构
Xie et al. Foreign object damage simulation of aero-engine blade
CN214749371U (zh) 一种低速冲击碳纤维复合材料梁夹具
Troha et al. Composite Inlays Increase Flutter Resistance of Turbine Engine Fan Blades
Zuo Research on foreign object damage and grinding repair schemes of compressor blades
CN112903426A (zh) 一种低速冲击碳纤维复合材料梁夹具
CN102825565A (zh) 关小涡轮导向器喉部面积的掰尾夹具及具有其的掰尾工具
CN115408777A (zh) 涡轮叶片叶根尾缘的特征模拟件设计方法及特征模拟件

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 20772902

Country of ref document: EP

Kind code of ref document: A1

NENP Non-entry into the national phase

Ref country code: DE

122 Ep: pct application non-entry in european phase

Ref document number: 20772902

Country of ref document: EP

Kind code of ref document: A1