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公开(公告)号：EP4269751A1
公开(公告)日：2023-11-01
申请号：EP23170157.4
申请日：2023-04-26
发明人： BALIKE, Krishna Prasad , WONG, Edwin , FUDGE, Daniel , THERATIL, Ignatius , STONE, Paul , MASON, Charles , ANAND, Karan , DORAN, Pascal
摘要： A method of manufacturing a mistuned rotor includes: obtaining a rotor 12 having a hub 22 and a plurality of blades 24 protruding from the hub 22, the plurality of blades 24 including first blades 28 and second blades 30 disposed in alternation around a central axis 11 of the rotor 12, natural vibration frequencies of the first blades 28 different from natural vibration frequencies of the second blades 30; determining that a difference between a first natural vibration frequency of a first blade 28 of the first blades and a second natural vibration frequency of a second blade 30 of the second blades is below a threshold L; and modifying a shape of the first blade 28 until the difference between the first natural vibration frequency and the second natural vibration frequency is at or above the threshold L.
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公开(公告)号：EP2990604A1
公开(公告)日：2016-03-02
申请号：EP15182551.0
申请日：2015-08-26
CPC分类号： F01D5/141 , F01D5/145 , F04D29/324 , F05D2240/303 , F05D2250/71
摘要： A rotary airfoil (20) in a gas turbine engine is provided. The airfoil (20) includes opposed pressure and suction sides (32, 34) joined together at chordally opposite leading and trailing edges (36, 38). The pressure and suction sides (32, 34) extend in a span direction from a root (25) to a tip (27). An axial component of a center of gravity of a cross-section taken chordally toward the tip (27) of the airfoil (20) is being upstream relative to an axial component of a center of gravity of a cross-section taken chordally toward the root (25) of the airfoil (20). A method for forming such blade is also presented.
摘要翻译： 在燃气涡轮发动机的旋转翼型件（20）设置。 该翼型件（20）包括相对的压力侧和吸力侧（32，34）在弦相对的前缘和后缘（36，38）连接在一起。 压力侧和吸力侧（32，34）从根（25）至尖端（27）在一个方向上翼展延伸。 的横截面的重心的弦朝向所述翼型件的尖端（27）（20）截取的中心的轴向分量是在弦朝向根截取的横截面的重心的轴向部件能够相对于上游侧 （25）的翼型件（20）。 因此，一种用于形成检查刀片方法。
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公开(公告)号：EP2990604B1
公开(公告)日：2019-07-17
申请号：EP15182551.0
申请日：2015-08-26
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公开(公告)号：EP2990603B1
公开(公告)日：2020-03-18
申请号：EP15182547.8
申请日：2015-08-26
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公开(公告)号：EP3608547A1
公开(公告)日：2020-02-12
申请号：EP19189674.5
申请日：2019-08-01
摘要： A diffuser pipe assembly (20) for a compressor (14b) of a gas turbine engine (10) includes diffuser pipes (22) circumferentially distributed around a central axis (A) and configured for distributing a flow of compressed air from the compressor (14b) to the combustor (16). Each of the diffuser pipes (22) curves between an inlet end (22e) and an outlet end (22f). A first subset of the diffuser pipes (221) has a natural vibration frequency different than a natural vibration frequency of at least a second subset of the diffuser pipes (222). A method of operating a compressor (14b) including the diffuser pipe assembly (20) is also disclosed.
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公开(公告)号：EP3026214B1
公开(公告)日：2019-04-10
申请号：EP15186827.0
申请日：2015-09-25
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7.发明公开
公开(公告)号：EP4290050A1
公开(公告)日：2023-12-13
申请号：EP23173772.7
申请日：2023-05-16
发明人： AZMI, Ali , ANAND, Karan , CAPRON, Alexandre , YAM, Tammy , CLAUDE, Sylvain , BALIKE, Krishna Prasad , BRUNI, Anthony Robert , GONZALEZ, Juan Ignacio Ruiz-Gopegui , MERINO, Cesar Valbuena , JARNE, Ignacio Javier Ucin
摘要： An aircraft engine, has: an upstream stator (30) having upstream stator vanes (31) circumferentially distributed about a central axis; and a downstream stator (40) having downstream stator vanes (41; 42, 43) circumferentially distributed about the central axis, the downstream stator (40) located downstream of the upstream stator (30) relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes (31) being different than a number of the downstream stator vanes (41; 42, 43), the downstream stator vanes (41; 42, 43) including: a first vane (42) made of a first material, a major portion of a leading edge (41A) of the first vane (42) circumferentially overlapped by one of the upstream stator vanes (31), and a second vane (43) made of a second material having a greater stiffness, strength, and/or ductility than that of the first material, a major portion of a leading edge of the second vane (43) exposed via a spacing (32) defined between two of the upstream stator vanes (31).
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公开(公告)号：EP4286649A1
公开(公告)日：2023-12-06
申请号：EP23176107.3
申请日：2023-05-30
发明人： CAPRON, Alexandre , ANAND, Karan , YAM, Tammy , AZMI, Ali , CLAUDE, Sylvain , BALIKE, Krishna Prasad , MITTAL, Prakul
IPC分类号： F01D5/14
摘要： An aircraft engine (10), has: an upstream stator (30) having upstream stator vanes (31) distributed about a central axis (11); and a downstream stator (40) having downstream stator vanes (41) distributed about the central axis (11), the downstream stator (40) located downstream of the upstream stator (30), a number of the upstream stator vanes (31) different than a number of the downstream stator vanes (41), the downstream stator vanes (41) including: a first vane (42), a major portion of a leading edge (41A) of the first vane (42) circumferentially overlapped by one of the upstream stator vanes (31); and a second vane (43) differing from the first vane (42) by a geometric parameter, the geometric parameter causing the second vane (43) to have one or more of: a stiffness greater than that of the first vane (42), and a major portion of a leading edge (41A) of the second vane (43) circumferentially overlapped by another one of the upstream stator vanes (31).
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公开(公告)号：EP3026214A1
公开(公告)日：2016-06-01
申请号：EP15186827.0
申请日：2015-09-25
CPC分类号： F01D5/141 , F01D9/041 , F04D29/324 , F05D2220/32 , F05D2240/12 , F05D2240/30 , F05D2240/301 , F05D2250/292
摘要： An airfoil (26) in a gas turbine engine includes opposed pressure and suction sides (32, 34) joined together at chordally opposite leading and trailing edges (36, 38). The pressure and suction sides (32, 34) extend spanwise from a root (25) to a tip (27) of the airfoil (26). The airfoil (26) has a spanwise distribution of maximum thicknesses of chordwise cross-sections of the airfoil (26). The spanwise distribution of maximum thicknesses decreases from the root (25) to the tip (27). The spanwise distribution is stepped between a first portion (P1) extending from the root (25) and a second portion (P2) extending to the tip (27).
摘要翻译： 燃气涡轮发动机中的翼型件（26）包括在弦向相对的前缘和后缘（36,38）处连接在一起的相对的压力侧和吸力侧（32,34）。 压力侧和吸力侧（32,34）从翼型件（26）的根部（25）向尖端（27）展向延伸。 翼型件（26）具有翼型件（26）的弦向横截面的最大厚度的翼展方向分布。 最大厚度的展向分布从根部（25）到尖端（27）减小。 翼展方向分布在从根部（25）延伸的第一部分（P1）和延伸至尖端（27）的第二部分（P2）之间分级。
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10.发明公开COMPRESSOR ROTOR AIRFOIL AND METHOD OF FORMING A BLADE 审中-公开VERDICHTERROTORSCHAUFEL UND VERFAHREN ZUR HERSTELLUNG EINER SCHAUFEL
公开(公告)号：EP2990603A1
公开(公告)日：2016-03-02
申请号：EP15182547.8
申请日：2015-08-26
摘要： A compressor rotor airfoil (20) in a gas turbine engine is presented. Opposed pressure and suction sides (32,34) are joined together at chordally opposite leading and trailing edges (36,38). The pressure and suction sides (32,34) extend in a span direction from a root (25) to a tip (27). A leading edge dihedral angle (β) is defined at a point on the leading edge (36) between a tangent (T) to the airfoil (20) and a vertical (V). The leading edge dihedral angle (β) has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.
摘要翻译： 提出了一种在燃气涡轮发动机中的压缩机转子翼型件（20）。 相反的压力和吸力侧（32,34）在弦向相反的前缘和后缘（36,38）处连接在一起。 压力和吸力侧（32,34）在跨度方向上从根部（25）延伸到尖端（27）。 前缘二面角（²）被限定在前缘（36）上与翼型件（20）的切线（T）和垂直（V）之间的点处。 前缘二面角（²）具有跨度分布。 分布至少有一个拐点。 还提出了一种减小压缩机转子叶片和围绕叶片的壳体之间的摩擦角的方法。
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