公开(公告)号：US20180045221A1

公开(公告)日：2018-02-15

申请号：US15236745

申请日：2016-08-15

Applicant: General Electric Company

Inventor： Daniel Elmer King ,  William Edward Cody

IPC: F04D29/54 ,  F04D29/52 ,  F04D29/56

CPC classification number: F04D29/542 ,  F01D9/02 ,  F01D9/047 ,  F04D29/522 ,  F04D29/563 ,  F05D2240/12 ,  F05D2240/128 ,  F05D2240/14 ,  F05D2250/00 ,  Y02T50/672

Abstract: A strut for a gas turbine engine includes a body defining a first side and an opposite second side. The first side is spaced from the second side along the circumferential direction of the gas turbine engine. Additionally, the body includes an inner section, a middle section, and an outer section. Each of the inner, middle, and outer sections are arranged in series order along a span of the strut and define a thickness between the first and second sides of the strut. The thickness of the middle section is greater than the thicknesses of the inner section and of the outer section to increase the strut's resistance to buckling in response to forces exerted thereon.

公开(公告)号：US09890657B2

公开(公告)日：2018-02-13

申请号：US13593814

申请日：2012-08-24

Applicant: Amr Ali

Inventor： Amr Ali

IPC: F01D25/00 ,  F02C7/24 ,  F02C7/045 ,  F02K1/82 ,  F02K3/06

CPC classification number: F01D25/00 ,  F02C7/045 ,  F02C7/24 ,  F02K1/827 ,  F02K3/06 ,  F05D2260/4031 ,  Y02T50/672

Abstract: An example acoustic treatment for a geared turbomachine establishes a portion of the core cowl aft the fan cowl.

公开(公告)号：US20180037515A1

公开(公告)日：2018-02-08

申请号：US15549658

申请日：2016-02-09

Applicant: Mitsubishi Heavy Industries Aero Engines, Ltd.

Inventor： Mineaki MATSUMOTO ,  Takayuki KURIMURA ,  Kosuke NISHIKAWA ,  Tadayuki HANADA

IPC: C04B41/89 ,  B32B18/00 ,  C04B41/87

CPC classification number: C04B41/89 ,  B32B18/00 ,  C04B41/009 ,  C04B41/52 ,  C04B41/87 ,  C04B2111/0025 ,  C23C4/02 ,  C23C4/11 ,  F01D5/282 ,  F01D5/284 ,  F01D5/288 ,  F05D2300/15 ,  F05D2300/211 ,  F05D2300/6033 ,  Y02T50/672 ,  C04B35/565 ,  C04B35/806 ,  C04B35/584 ,  C04B41/4527 ,  C04B41/5096 ,  C04B41/5024 ,  C04B2103/0021

Abstract: Provided are a coated member in which damage of a coating film can be suppressed in a high temperature environment and the coating may be performed at low cost, and a method of manufacturing the same. A coated member includes a bond coat and a top coat sequentially laminated on a substrate made of a Si-based ceramic or a SiC fiber-reinforced SiC matrix composite, wherein the top coat includes a layer composed of a mixed phase of a (Y1-aLn1a)2Si2O7 solid solution (here, Ln1 is any one of Nd, Sm, Eu, and Gd) and Y2SiO5 or a (Y1-bLn1′b)2SiO5 solid solution (here, Ln1′ is any one of Nd, Sm, Eu, and Gd), or a mixed phase of a (Y1-cLn2c)2Si2O7 solid solution (here, Ln2 is any one of Sc, Yb, and Lu) and Y2SiO5 or a (Y1-dLn2′d)2SiO5 solid solution (here, Ln2′ is any one of Sc, Yb, and Lu).

公开(公告)号：US09885244B2

公开(公告)日：2018-02-06

申请号：US14418534

申请日：2013-06-28

Applicant: GENERAL ELECTRIC COMPANY

Inventor： Nicholas Joseph Kray ,  Dong-Jin Shim ,  Tod Winton Davis

IPC: F01D5/28 ,  F01D5/26 ,  F01D5/14 ,  F04D29/32 ,  F04D29/02 ,  C22C19/05 ,  C22C38/02 ,  C22C38/04 ,  C22C38/42 ,  C22C38/48 ,  C22C38/54

CPC classification number: F01D5/26 ,  C22C19/057 ,  C22C38/02 ,  C22C38/04 ,  C22C38/42 ,  C22C38/48 ,  C22C38/54 ,  F01D5/147 ,  F01D5/282 ,  F01D5/288 ,  F04D29/023 ,  F04D29/324 ,  F05D2220/36 ,  F05D2240/303 ,  F05D2300/171 ,  F05D2300/175 ,  Y02T50/672 ,  Y10T29/49337

Abstract: A metallic leading edge protective strip adapted to provide impact protection for a leading edge of an airfoil of a turbomachine. The protective strip is formed of a stainless steel or a nickel-based alloy and is denser and provides increased strength and elasticity characteristics as compared to an identical protective strip formed of a titanium-based alloy. The protective strip is particularly suitable for use with composite blades and allows for thinner airfoils, thereby improving engine efficiency.

公开(公告)号：US20180023408A1

公开(公告)日：2018-01-25

申请号：US15217305

申请日：2016-07-22

Applicant: Rolls-Royce North American Technologies, Inc.

Inventor： Edward C. Rice ,  Daniel K. Vetters

IPC: F01D11/08 ,  F01D25/24 ,  F01D25/00

CPC classification number: F01D11/08 ,  F01D25/005 ,  F01D25/24 ,  F05D2220/32 ,  F05D2230/60 ,  F05D2230/642 ,  F05D2240/11 ,  F05D2240/12 ,  F05D2240/14 ,  F05D2260/36 ,  F05D2300/6033 ,  Y02T50/672

Abstract: A gas turbine engine includes a turbine shroud assembly arranged around a turbine wheel assembly. The turbine shroud assembly includes an annular turbine case, a blade track, and a forward case. The turbine case is arranged around a central axis of the turbine assembly. The blade track is located radially between the turbine case and the central axis. The forward case is coupled to the turbine case and is arranged to block axial movement of the blade track relative to the turbine case.

公开(公告)号：US20180023407A1

公开(公告)日：2018-01-25

申请号：US15708603

申请日：2017-09-19

Applicant: United Technologies Corporation

Inventor： Brandon W. Spangler ,  Ryan Alan Waite ,  Steven Bruce Gautschi ,  Benjamin F. Hagan

IPC: F01D9/06 ,  F01D25/14 ,  F01D25/12

CPC classification number: F01D9/065 ,  F01D5/186 ,  F01D5/187 ,  F01D25/12 ,  F01D25/14 ,  F05D2240/126 ,  F05D2260/20 ,  F05D2260/202 ,  F05D2260/204 ,  Y02T50/672 ,  Y02T50/673 ,  Y02T50/676

Abstract: An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges. The pressure and suction side walls surround an airfoil cavity and provide an exterior airfoil surface. A baffle is arranged in the airfoil cavity and includes a supply hole. Ribs extend between at least one of the pressure and suction side walls into the airfoil cavity to support the baffle relative to the at least one of the pressure and suction side walls. The ribs are configured to provide a serpentine cooling passage between the baffle and at least one of the pressure and suction side walls. The serpentine cooling passage has first and second passes joined by a bend. The ribs form a film cooling cavity between the first and second passes. The supply hole fluidly connects the baffle to the film cooling cavity. Film cooling holes extend through the at least one of the pressure and suction side walls. The film cooling holes are in fluid communication with the film cooling cavity.

公开(公告)号：US09874103B2

公开(公告)日：2018-01-23

申请号：US14402182

申请日：2013-05-31

Applicant: SNECMA

Inventor： Gilbert Michel Marin Leconte ,  Gilles Charles Casimir Klein ,  Jean-Michel Patrick Maurice Franchet ,  Dominique Magnaudeix

IPC: F01D5/14 ,  F01D5/28 ,  B21K23/00 ,  B21D53/78 ,  B21J5/12 ,  B23P15/04 ,  B21K3/04 ,  F04D29/02 ,  F04D29/32 ,  B21D26/033 ,  B21D26/053 ,  B21D35/00 ,  B23K9/167 ,  B23K20/02 ,  B23K20/233 ,  B23K20/24 ,  B23K101/00 ,  B23K103/14

CPC classification number: F01D5/282 ,  B21D26/033 ,  B21D26/053 ,  B21D35/00 ,  B21D35/003 ,  B21D53/78 ,  B21J5/12 ,  B21K3/04 ,  B21K23/00 ,  B23K9/167 ,  B23K20/021 ,  B23K20/233 ,  B23K20/24 ,  B23K2101/001 ,  B23K2103/14 ,  B23P15/04 ,  F04D29/023 ,  F04D29/324 ,  F05D2240/303 ,  F05D2300/174 ,  F05D2300/603 ,  F05D2300/702 ,  Y02T50/672 ,  Y10T29/49337

Abstract: A method of making a metal reinforcing member that is to be mounted on a leading edge or a trailing edge of a composite blade of a turbine engine, the method including: shaping two metal sheets, positioning them on either side of a core including at least one recess that is to form a mold for a spacer for positioning the reinforcing member, assembling them together under a vacuum, conforming them against the core by hot isostatic compression, and cutting them to separate the reinforcing member and release the core.

公开(公告)号：US09869183B2

公开(公告)日：2018-01-16

申请号：US14799994

申请日：2015-07-15

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Brian D. Merry ,  James D. Hill ,  Wesley K. Lord

IPC: F01D5/08 ,  F01D5/28

CPC classification number: F01D5/082 ,  F01D5/084 ,  F01D5/085 ,  F01D5/288 ,  Y02T50/672 ,  Y02T50/676

Abstract: A rotor for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a rotor disk rotatable about an axis and a gas path wall coupled to and radially outward of the rotor disk. The gas path wall bounds a radially inward portion of a gas path. A plurality of rotor spokes are radially intermediate the rotor disk and the gas path wall. The plurality of rotor spokes is circumferentially spaced to define a plurality of cooling channels intermediate the rotor disk and the gas path wall. A thermal barrier coating is disposed on a surface of at least one of the plurality of cooling channels. A method of cooling a rotor assembly is also disclosed.

公开(公告)号：US20180010518A1

公开(公告)日：2018-01-11

申请号：US15642987

申请日：2017-07-06

Applicant: Rolls-Royce Deutschland Ltd & Co KG

Inventor： Jens ORTMANNS

IPC: F02C7/04 ,  B64C7/02 ,  B64D29/00 ,  B64C21/02 ,  B64D33/02

CPC classification number: F02C7/04 ,  B64C7/02 ,  B64C21/02 ,  B64C2230/22 ,  B64D29/00 ,  B64D33/02 ,  B64D2033/0226 ,  Y02T50/672

Abstract: The invention relates to an engine nacelle, including: a nacelle wall that has an inner side and an outer side; an inlet lip that is embodied at that end of the engine nacelle that is formed upstream; and an engine intake that takes in the air required for the respective engine and that is formed by the inner side of the nacelle wall. It is provided that the nacelle wall includes an air-permeable structure that extends from the outer side to the inner side of the nacelle wall, and that is configured for passing air that flows against the outer side from the outer side to the inner side. The invention further relates to a method for influencing the flows inside an engine nacelle.

公开(公告)号：US20180010469A1

公开(公告)日：2018-01-11

申请号：US15547655

申请日：2015-12-08

Applicant: Siemens Aktiengesellschaft

Inventor： Kai Kadau ,  Jose Antonio Pascual-Gutierrez ,  Atin Sharma

IPC: F01D5/28 ,  F01D5/14 ,  F01D11/12

CPC classification number: C04B41/81 ,  C04B41/91 ,  C23C4/134 ,  C23C16/045 ,  F01D5/14 ,  F01D5/147 ,  F01D5/18 ,  F01D5/186 ,  F01D5/28 ,  F01D5/282 ,  F01D5/288 ,  F01D9/023 ,  F01D11/08 ,  F01D11/12 ,  F01D11/122 ,  F04D29/324 ,  F04D29/542 ,  F04D29/5853 ,  F05D2220/32 ,  F05D2230/10 ,  F05D2230/312 ,  F05D2230/313 ,  F05D2230/90 ,  F05D2240/35 ,  F05D2250/132 ,  F05D2250/294 ,  F05D2250/60 ,  F05D2300/502 ,  F05D2300/5023 ,  F05D2300/6033 ,  Y02T50/672 ,  Y02T50/673 ,  Y02T50/676 ,  Y02T50/6765

Abstract: Turbine engine (80) components, such as blades (92), vanes (104, 106), ring segment 110 abradable surfaces 120, or transitions (85), have furcated engineered groove features (EGFs) (403, 404, 418, 509, 511, 512) that cut into the outer surface of the component's thermal barrier coating (TBC). In some embodiments, the EGF planform pattern defines adjoining outer hexagons (560, 640, 670, 690, 710). In some embodiments, the EGF pattern further defines within each outer hexagon (560, 640, 670, 690, 710) a planform pattern of adjoining inner polygons (570, 580, 590, 600, 610, 680, 682, 700, 702, 704, 705, 720). At least three respective groove segments (509, 511, 512) within the EGF pattern (506, 507, 508) converge at each respective outer hexagonal vertex (510, 564) or inner polygonal vertex (574, 564, 604, 614) in a multifurcated pattern, so that crack-inducing stresses are attenuated in cascading fashion, as the stress (σA) is furcated (σB, σC) at each successive vertex juncture.