公开(公告)号：US20220178260A1

公开(公告)日：2022-06-09

申请号：US17113166

申请日：2020-12-07

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： San Quach ,  Tyler G. Vincent ,  Cheng Gao ,  Howard J. Liles

IPC: F01D5/18 ,  F01D5/14 ,  F01D5/28

Abstract: A vane arc segment includes an airfoil piece that defines first and second platforms and a hollow airfoil section that has an internal cavity and extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the flange. There is a conformal thermal insulation blanket disposed on the flange.

公开(公告)号：US11220917B1

公开(公告)日：2022-01-11

申请号：US17010904

申请日：2020-09-03

Applicant: Raytheon Technologies Corporation

Inventor： San Quach ,  Raymond Surace

IPC: F01D5/18 ,  F01D9/06 ,  F01D25/12

Abstract: A gas turbine engine component according to an example of the present disclosure includes, among other things, an external wall extending in a thickness direction between first and second wall surfaces. The first wall surface bounds an internal cavity, and establishes at least one surface depression along an external surface contour. The external wall includes at least one cooling passage having an outlet port established along the at least one surface depression. A method of fabricating a gas turbine engine component is also disclosed.

公开(公告)号：US11215059B1

公开(公告)日：2022-01-04

申请号：US17011171

申请日：2020-09-03

Applicant: Raytheon Technologies Corporation

Inventor： San Quach ,  Bryan P. Dube ,  Raymond Surace ,  Alex J. Schneider ,  Tyler G. Vincent

IPC: F01D5/18 ,  F01D25/12 ,  F02C3/04

Abstract: An airfoil includes a ceramic airfoil section that defines leading and trailing edges, pressure and suction sides, and radially inner and outer ends. The span of the airfoil section has first, second, and third radial span zones. There is a row of cooling holes in an aft 25% of the axial span. The row of cooling holes extends though the first, second, and third radial span zones. The cooling holes in the first radial span zone define a first pitch P1, the cooling holes in the second radial span zone define a second pitch P2, and the cooling holes in the third radial span zone define a third pitch P3, wherein P2

公开(公告)号：US12044130B2

公开(公告)日：2024-07-23

申请号：US17978315

申请日：2022-11-01

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： San Quach ,  Robert A. White, III ,  Tyler G. Vincent

IPC: F01D11/00 ,  F01D5/28 ,  F01D9/04 ,  F01D11/14 ,  F01D25/12

CPC classification number: F01D11/005 ,  F01D5/284 ,  F01D9/041 ,  F01D11/14 ,  F01D25/12 ,  F01D9/042 ,  F05D2220/32 ,  F05D2240/55 ,  F05D2240/81 ,  F05D2260/201 ,  F05D2300/20 ,  F05D2300/6033

Abstract: A gas turbine engine includes a ceramic wall for bounding an engine core gas path. The ceramic wall has a ceramic wall first side that faces the engine core gas path and a ceramic wall second side that faces away from the engine core gas path. There is a metallic wall adjacent the ceramic wall second side. The metallic wall has a metallic wall first side that faces the ceramic wall and a metallic wall second side that faces away from the ceramic wall. The metallic wall and the ceramic wall are spaced apart such that there is a channel there between. There is a seal on the ceramic wall second side, and the metallic wall has at least one cooling hole adjacent the seal for emitting cooling air to cool the seal.

公开(公告)号：US11536145B2

公开(公告)日：2022-12-27

申请号：US17227084

申请日：2021-04-09

Applicant: Raytheon Technologies Corporation

Inventor： San Quach ,  Alex J. Schneider

IPC: F01D5/14 ,  F01D5/28 ,  F01D9/02

Abstract: An airfoil vane assembly according to an exemplary embodiment of this disclosure, among other possible things includes a vane piece having a first vane platform, a second vane platform, and a hollow airfoil section joining the first vane platform and the second vane platform, a spar piece having a spar platform and a spar extending from the spar platform into the hollow airfoil section, and at least one seal arranged at a sealing surface of the spar platform and sealing between the spar platform and the first vane platform. The airfoil vane assembly also includes a thermal barrier coating disposed on the spar piece. The sealing surface is free from the thermal barrier coating. A gas turbine engine and a method of making a spar piece for an airfoil vane assembly are also disclosed.

公开(公告)号：US11499443B2

公开(公告)日：2022-11-15

申请号：US17128398

申请日：2020-12-21

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： San Quach ,  Robert A. White, III ,  Tyler G. Vincent

IPC: F01D11/00 ,  F01D11/14 ,  F01D25/12

Abstract: A gas turbine engine includes a ceramic wall for bounding an engine core gas path. The ceramic wall has a ceramic wall first side that faces the engine core gas path and a ceramic wall second side that faces away from the engine core gas path. There is a metallic wall adjacent the ceramic wall second side. The metallic wall has a metallic wall first side that faces the ceramic wall and a metallic wall second side that faces away from the ceramic wall. The metallic wall and the ceramic wall are spaced apart such that there is a channel there between. There is a seal on the ceramic wall second side, and the metallic wall has at least one cooling hole adjacent the seal for emitting cooling air to cool the seal.

公开(公告)号：US20220325630A1

公开(公告)日：2022-10-13

申请号：US17227084

申请日：2021-04-09

Applicant: Raytheon Technologies Corporation

Inventor： San Quach ,  Alex J. Schneider

IPC: F01D5/28

Abstract: An airfoil vane assembly according to an exemplary embodiment of this disclosure, among other possible things includes a vane piece having a first vane platform, a second vane platform, and a hollow airfoil section joining the first vane platform and the second vane platform, a spar piece having a spar platform and a spar extending from the spar platform into the hollow airfoil section, and at least one seal arranged at a sealing surface of the spar platform and sealing between the spar platform and the first vane platform. The airfoil vane assembly also includes a thermal barrier coating disposed on the spar piece. The sealing surface is free from the thermal barrier coating. A gas turbine engine and a method of making a spar piece for an airfoil vane assembly are also disclosed.

公开(公告)号：US20220299206A1

公开(公告)日：2022-09-22

申请号：US17207193

申请日：2021-03-19

Applicant: Raytheon Technologies Corporation

Inventor： San Quach ,  Thomas N. Slavens

IPC: F23R3/42 ,  F02C3/00

Abstract: In one exemplary embodiment, a combustor liner includes a first portion extending in a substantially axial direction and a second portion that extending in the substantially axial direction. A step connects the first portion and the second portion. The step is arranged at an angle to the second portion that is less than 90°. The step has a step height defined as a distance between the first portion and the second portion. The first portion, second portion, and step are formed as a unitary ceramic component. A slot extends through the step, and a ratio of a height of the slot to a height of the step is greater than 0.66.

公开(公告)号：US20220195879A1

公开(公告)日：2022-06-23

申请号：US17128398

申请日：2020-12-21

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： San Quach ,  Robert A. White, III ,  Tyler G. Vincent

IPC: F01D11/00 ,  F01D11/14 ,  F01D25/12

Abstract: A gas turbine engine includes a ceramic wall for bounding an engine core gas path. The ceramic wall has a ceramic wall first side that faces the engine core gas path and a ceramic wall second side that faces away from the engine core gas path. There is a metallic wall adjacent the ceramic wall second side. The metallic wall has a metallic wall first side that faces the ceramic wall and a metallic wall second side that faces away from the ceramic wall. The metallic wall and the ceramic wall are spaced apart such that there is a channel there between. There is a seal on the ceramic wall second side, and the metallic wall has at least one cooling hole adjacent the seal for emitting cooling air to cool the seal.

公开(公告)号：US11339667B2

公开(公告)日：2022-05-24

申请号：US16990420

申请日：2020-08-11

Applicant: Raytheon Technologies Corporation

Inventor： San Quach ,  Bryan P. Dube ,  Tracy A. Propheter-Hinckley ,  Allan N. Arisi ,  Adam P. Generale ,  Lucas Dvorozniak ,  Howard J. Liles

IPC: F01D5/18 ,  F01D5/28

Abstract: A gas turbine engine component according to an example of the present disclosure includes a wall extending in a thickness direction between first and second wall surfaces. The first wall surface bounds an internal cavity. The wall includes a plurality of cooling passages. Each of the cooling passages extend in a first direction between an inlet port and an outlet port coupled to a respective diffuser, and the inlet port coupled to the internal cavity along the first wall surface. Sidewalls of adjacent diffusers are conjoined to establish a common diffuser region interconnecting the diffusers and a common outlet along the second wall surface. A method of cooling a gas turbine engine component is also disclosed.