公开(公告)号：WO2014025440A3

公开(公告)日：2014-04-17

申请号：PCT/US2013041776

申请日：2013-05-20

Applicant: UNITED TECHNOLOGIES CORP

Inventor： BEATTIE JEFFREY S ,  JACQUES JEFFREY MICHAEL ,  PIGGUSH JUSTIN D

IPC: F01D5/18 ,  F01D25/12 ,  F02C7/12

CPC classification number: F01D5/187 ,  B22C9/04 ,  B22C9/103 ,  F05D2230/211 ,  F05D2260/204 ,  Y02T50/676 ,  Y10T29/49316

Abstract: A method of fabricating an airfoil includes the steps of fabricating a first core including a first plurality of ribs defining a first plurality of passages of a completed airfoil, and fabricating as second core including a second plurality of ribs defining a second plurality of passages of the completed airfoil. The second plurality of ribs includes a plurality of standoffs. The plurality of standoffs set a spacing between the first plurality of ribs and the second plurality of ribs to define a spacing between the first plurality of channels and the second plurality of channels of the completed airfoil. The airfoil is then molded about the core assembly. Once completed, the core assembly is removed to provide a completed airfoil incorporating multiple microcircuits with a desired stability and structural integrity.

Abstract translation: 一种制造翼型件的方法包括以下步骤：制造包括限定完整翼型的第一多个通道的第一多个肋的第一芯，以及制造包括限定第二多个通道的第二多个肋的第二芯， 完成机翼。 第二多个肋包括多个支架。 多个支座在第一组多个肋和第二组多个肋之间设置间距，以限定完成的翼型的第一多个通道和第二多个通道之间的间距。 然后翼型围绕芯组件成型。 一旦完成，核心组件被移除以提供包含具有期望的稳定性和结构完整性的多个微电路的完整翼型。

公开(公告)号：WO2014186005A3

公开(公告)日：2015-02-26

申请号：PCT/US2014015194

申请日：2014-02-07

Applicant: UNITED TECHNOLOGIES CORP

Inventor： LEWIS SCOTT D ,  RAPP BRANDON M ,  BEATTIE JEFFREY S ,  HOUGH MATTHEW ANDREW ,  TELLER BRET M ,  JACQUES JEFFREY MICHAEL ,  ASTERLIN MAX

IPC: F01D5/18 ,  F01D5/08 ,  F01D25/12 ,  F02C7/18 ,  F04D29/58

CPC classification number: F01D5/186 ,  F01D9/023 ,  F05D2240/81 ,  F05D2260/202 ,  F23R3/02

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge, circumferentially extends between a first mate face and a second mate face, and includes a gas path surface and a non-gas path surface. The component defines at least one cavity that extends at least partially inside of the component. A first plurality of cooling holes extends from the at least one cavity to at least one of the first mate face and the second mate face and a second plurality of cooling holes extends from either the at least one cavity or the non-gas path surface to the gas path surface.

Abstract translation: 根据本公开的示例性方面的用于燃气涡轮发动机的部件尤其包括在前缘和后缘之间轴向延伸的平台，其周向地在第一配合面和第二配合面之间延伸，以及 包括气体通道表面和非气体通道表面。 该部件限定至少一个至少部分地在部件内部延伸的空腔。 第一多个冷却孔从所述至少一个空腔延伸到所述第一配合面和所述第二配合面中的至少一个，并且第二多个冷却孔从所述至少一个空腔或所述非气体通道表面延伸到 气路面。

公开(公告)号：WO2014099082A3

公开(公告)日：2014-08-28

申请号：PCT/US2013061920

申请日：2013-09-26

Applicant: UNITED TECHNOLOGIES CORP

Inventor： BEATTIE JEFFREY S ,  JACQUES JEFFREY MICHAEL

IPC: F01D5/14 ,  F01D5/20 ,  F02C3/14 ,  F02C7/20

CPC classification number: F01D5/30 ,  F01D5/18 ,  F01D5/3007

Abstract: A turbine blade for a gas turbine engine includes an airfoil that extends in a first radial direction from a platform. A root extends from the platform in a second radial direction and has opposing lateral sides that provide a firtree-shaped contour. The contour includes first, second and third lobes on each of the lateral sides and that tapers relative to the radial direction away from the platform. The first, second and third lobes each provide contact surfaces arranged at about 45° relative to the radial direction. A contact plane on each lateral side at an angle of about 11° relative to the radial direction defining a contact point on each of the contact surfaces. The first, second and third lobes each include first, second and third grooves that are substantially aligned with one another along an offset plane spaced a uniform offset distance from the contact plane.

Abstract translation: 用于燃气涡轮发动机的涡轮机叶片包括从平台沿第一径向方向延伸的翼型件。 根部沿第二径向方向从平台延伸，并具有提供冷杉状轮廓的相对侧面。 轮廓包括在每个横向侧上的第一，第二和第三凸角，并且所述轮廓相对于远离平台的径向方向逐渐变细。 第一，第二和第三凸角各自提供相对于径向方向布置成约45°的接触表面。 在每个侧面上相对于径向方向成约11°角的接触面​​限定了每个接触面上的接触点。 第一，第二和第三瓣各自包括第一，第二和第三槽，所述第一，第二和第三槽沿着与接触平面间隔均匀偏移距离的偏移平面彼此基本对齐。

公开(公告)号：SG11201407921VA

公开(公告)日：2015-01-29

申请号：SG11201407921V

申请日：2013-05-20

Applicant: UNITED TECHNOLOGIES CORP

Inventor： BEATTIE JEFFREY S ,  JACQUES JEFFREY MICHAEL ,  PIGGUSH JUSTIN D

IPC: F01D5/18

Abstract: A method of fabricating an airfoil includes the steps of fabricating a first core including a first plurality of ribs defining a first plurality of passages of a completed airfoil, and fabricating as second core including a second plurality of ribs defining a second plurality of passages of the completed airfoil. The second plurality of ribs includes a plurality of standoffs. The plurality of standoffs set a spacing between the first plurality of ribs and the second plurality of ribs to define a spacing between the first plurality of channels and the second plurality of channels of the completed airfoil. The airfoil is then molded about the core assembly. Once completed, the core assembly is removed to provide a completed airfoil incorporating multiple microcircuits with a desired stability and structural integrity.

公开(公告)号：EP2867501A4

公开(公告)日：2015-07-08

申请号：EP13810147

申请日：2013-05-30

Applicant: UNITED TECHNOLOGIES CORP

Inventor： ZELESKY MARK F ,  TRINDADE RICARDO ,  TELLER BRET M ,  LEWIS SCOTT D ,  BEATTIE JEFFREY S ,  JACQUES JEFFREY MICHAEL ,  RAPP BRANDON M

IPC: F02C7/12 ,  F01D5/08 ,  F01D5/18 ,  F02K3/00

CPC classification number: F01D5/085 ,  F01D5/081 ,  F01D5/18 ,  F05D2240/81 ,  Y02T50/676

Abstract: A turbine blade for a gas turbine engine includes an airfoil including leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface extending in a radial direction. The trailing edge is arranged on an aft side of the turbine blade. A root supports a platform from which the airfoil extends and a cooling passage extends within the root in the radial direction to the airfoil. A lower wing is arranged beneath the platform on the aft side and extends in an axial direction to provide a U-shaped channel with the platform that extends in a circumferential direction. An impingement hole extends from the U-channel to the cooling passage.

公开(公告)号：EP2900924A4

公开(公告)日：2016-06-08

申请号：EP13864512

申请日：2013-09-26

Applicant: UNITED TECHNOLOGIES CORP

Inventor： BEATTIE JEFFREY S ,  JACQUES JEFFREY MICHAEL

IPC: F01D5/14 ,  F01D5/20 ,  F02C3/14 ,  F02C7/20

CPC classification number: F01D5/30 ,  F01D5/18 ,  F01D5/3007

Abstract: A turbine blade for a gas turbine engine includes an airfoil that extends in a first radial direction from a platform. A root extends from the platform in a second radial direction and has opposing lateral sides that provide a firtree-shaped contour. The contour includes first, second and third lobes on each of the lateral sides and that tapers relative to the radial direction away from the platform. The first, second and third lobes each provide contact surfaces arranged at about 45° relative to the radial direction. A contact plane on each lateral side at an angle of about 11° relative to the radial direction defining a contact point on each of the contact surfaces. The first, second and third lobes each include first, second and third grooves that are substantially aligned with one another along an offset plane spaced a uniform offset distance from the contact plane.

公开(公告)号：EP2956627A4

公开(公告)日：2016-09-28

申请号：EP14797049

申请日：2014-02-07

Applicant: UNITED TECHNOLOGIES CORP

Inventor： LEWIS SCOTT D ,  RAPP BRANDON M ,  BEATTIE JEFFREY S ,  HOUGH MATTHEW ANDREW ,  TELLER BRET M ,  JACQUES JEFFREY MICHAEL ,  ASTERLIN MAX

IPC: F01D5/18 ,  F01D5/08 ,  F01D11/24 ,  F01D25/12 ,  F02C7/18 ,  F04D29/58

CPC classification number: F01D5/186 ,  F01D9/023 ,  F05D2240/81 ,  F05D2260/202 ,  F23R3/02

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge, circumferentially extends between a first mate face and a second mate face, and includes a gas path surface and a non-gas path surface. The component defines at least one cavity that extends at least partially inside of the component. A first plurality of cooling holes extends from the at least one cavity to at least one of the first mate face and the second mate face and a second plurality of cooling holes extends from either the at least one cavity or the non-gas path surface to the gas path surface.

Abstract translation: 根据本公开的示例性方面的用于燃气涡轮发动机的部件尤其包括在前缘和后缘之间轴向延伸的平台，其周向地在第一配合面和第二配合面之间延伸，以及 包括气体路径表面和非气体通道表面。 该部件限定至少一个至少部分地在部件内部延伸的空腔。 第一多个冷却孔从所述至少一个空腔延伸到所述第一配合面和所述第二配合面中的至少一个，并且第二多个冷却孔从所述至少一个空腔或非气体通道表面延伸到 气路面。

公开(公告)号：EP2855854A4

公开(公告)日：2016-05-11

申请号：EP13827123

申请日：2013-05-20

Applicant: UNITED TECHNOLOGIES CORP

Inventor： BEATTIE JEFFREY S ,  JACQUES JEFFREY MICHAEL ,  PIGGUSH JUSTIN D

IPC: F01D5/18 ,  F01D25/12 ,  F02C7/12

CPC classification number: F01D5/187 ,  B22C9/04 ,  B22C9/103 ,  F05D2230/211 ,  F05D2260/204 ,  Y02T50/676 ,  Y10T29/49316

Abstract: A method of fabricating an airfoil includes the steps of fabricating a first core including a first plurality of ribs defining a first plurality of passages of a completed airfoil, and fabricating as second core including a second plurality of ribs defining a second plurality of passages of the completed airfoil. The second plurality of ribs includes a plurality of standoffs. The plurality of standoffs set a spacing between the first plurality of ribs and the second plurality of ribs to define a spacing between the first plurality of channels and the second plurality of channels of the completed airfoil. The airfoil is then molded about the core assembly. Once completed, the core assembly is removed to provide a completed airfoil incorporating multiple microcircuits with a desired stability and structural integrity.