With Cooling Passage Patents (Class 29/889.721)
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Publication number: 20120087803Abstract: A turbine bucket includes an airfoil portion at one end thereof; a root portion at an opposite end thereof; a platform portion between the airfoil portion and the root portion; at least one internal cavity within or radially inward of the platform portion having at least one film cooling hole extending between the at least one cavity and an external surface of the platform portion. The film cooling hole is curved along a length dimension of the film cooling hole.Type: ApplicationFiled: October 12, 2010Publication date: April 12, 2012Applicant: GENERAL ELECTRIC COMPANYInventors: Jesse Blair BUTLER, Benjamin LACY
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Publication number: 20120084981Abstract: In a method of working a film cooling hole which communicates with an internal cooling passage of a turbine blade from an outer surface, in the turbine blade which has a heat shield coating and the internal cooling passage, there is employed a method of working a cooling hole of a turbine including a step of executing a bond coat on a blade base material a step of piercing a cooling hole in accordance with an electric discharge machining, a step of executing a top coat, and a step of removing the top coat in accordance with a mechanical method with respect to a band-like region including a row of cooling holes, by using an abrasive blasting, a water jet method or the like. Accordingly, an occlusion of the cooling hole, and a damage of the TBC due to the piercing are hardly generated.Type: ApplicationFiled: October 5, 2011Publication date: April 12, 2012Inventors: Hideyuki Arikawa, Akira Mebata, Yoshitaka Kojima, Kunihiro Ichikawa, Takeshi Izumi
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Patent number: 8153923Abstract: There are a multiplicity of methods of making through-holes. In particular in the production of a multiplicity of film-cooling holes, as in gas turbine blades or combustion chamber elements, small time advantages are also important when making a hole. The method according to the invention, to make the hole close to the final contour in each case in sections in a top and a bottom region in order to then produce the final contour with other laser parameters, achieves time advantages.Type: GrantFiled: December 29, 2006Date of Patent: April 10, 2012Assignee: Siemens AktiengesellschaftInventors: Thomas Beck, Nigel-Philip Cox, Silke Settegast
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Publication number: 20120082564Abstract: A platform cooling arrangement in a turbine rotor blade having a platform that includes an interior cooling passage formed therein. The platform cooling arrangement may include: a main plenum residing just inboard of the planar topside and extending from an aft position to a forward position within one of the pressure side and the suction side of the platform, the main plenum having a longitudinal axis that is approximately parallel to the planar topside; a supply plenum that extends between the main plenum and the interior cooling passage; and a plurality of cooling apertures, each cooling aperture extending from one of the pressure side and the suction side slashface to a connection with the main plenum.Type: ApplicationFiled: September 30, 2010Publication date: April 5, 2012Inventors: Scott Edmond Ellis, Daniel Alan Hynum, John Wesley Harris, JR.
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Publication number: 20120082566Abstract: A platform cooling arrangement for a turbine rotor blade having a platform and an interior cooling passage and, in operation, a high-pressure coolant region and a low-pressure coolant region, wherein the platform includes a topside, which extends from the airfoil to a pressure side slashface, and an underside. The platform cooling arrangement may include: an airfoil manifold that resides near the junction of the pressure face of the airfoil and the platform; a slashface manifold that resides near the pressure side slashface; a high-pressure connector that connects the airfoil manifold to a high-pressure coolant region of the interior cooling passage; a low-pressure connector that connects the slashface manifold to a low-pressure coolant region of the interior cooling passage; cooling apertures that extend from a starting point along the pressure side slashface to a connection with the airfoil manifold, bisecting the slashface manifold therebetween; and a plurality of non-integral plugs.Type: ApplicationFiled: September 30, 2010Publication date: April 5, 2012Inventors: Scott Edmond Ellis, Daniel Alan Hynum
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Publication number: 20120082565Abstract: In a turbine rotor blade having a platform at an interface between an airfoil and a root, wherein the airfoil and the root include an interior cooling passage formed therein, wherein, in operation, the interior cooling passage comprises at least a high-pressure coolant region and a low-pressure coolant region, platform cooling arrangement that includes: a platform slot extending circumferentially from a mouth formed through the pressure side slashface; a high-pressure connector that connects the platform slot to the high-pressure coolant region of the interior cooling passage; a low-pressure connector that connects the platform slot to the low-pressure coolant region of the interior cooling passage; and a platform cooling cartridge removably engaged within the platform slot, the platform cooling cartridge comprising one or more cartridge cooling channels.Type: ApplicationFiled: September 30, 2010Publication date: April 5, 2012Inventors: Scott Edmond Ellis, Xiaoyong Fu
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Publication number: 20120082563Abstract: A micro gas turbine engine in which the turbine rotor blades are formed as an integral bladed rotor with cooling air passages formed within the blades and the rotor disk by an EDM process. an adjacent stator vane includes an air riding seal with an air cushion supplied through the vanes to provide cooling, and where the air cushion is then passed into the turbine blades and rotor disk to provide cooling for the turbine blades. With cooling of the turbine blades, higher turbine inlet temperatures for micro gas turbine engines can be produced.Type: ApplicationFiled: September 30, 2010Publication date: April 5, 2012Applicant: FLORIDA TURBINE TECHNOLOGIES, INC.Inventors: Jack W. Wilson, JR., John E. Ryznic, James P. Downs
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Publication number: 20120070307Abstract: A turbine blade includes a first side wall including a first tip edge, a second side wall opposite the first side wall and including a second tip edge, a tip wall between the first and second side walls, the tip wall recessed from the first tip edge of the first side wall and the second tip edge of the second side wall forming a coolant cavity, a tip recess cavity, a first parapet wall on the first side wall, and a second parapet wall on the second side wall, the coolant cavity defined by the tip wall, and the tip recess cavity defined by the tip wall, and the first and second parapet walls, a step formed between the first tip edge and the tip wall, a cooling hole through the first parapet wall, the step, and the tip wall, the cooling hole including an open and a closed channel section.Type: ApplicationFiled: September 22, 2010Publication date: March 22, 2012Applicant: HONEYWELL INTERNATIONAL INC.Inventors: Kin Poon, Malak Fouad Malak, Ardeshir Riahi, David Chou
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Patent number: 8109712Abstract: Disclosed is a turbine or compressor component with an integrated cooling channel, in particular a turbine blade, and a method for producing the same. The cooling channel of the component is subjected to internal pressure during a pressure impingement phase, the internal pressure being at a level sufficiently high that it causes the at least semiplastic deformation of the wall regions delimiting the cooling channel.Type: GrantFiled: January 24, 2007Date of Patent: February 7, 2012Assignee: Siemens AktiengesellschaftInventors: Fathi Ahmad, Michael Dankert
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Publication number: 20120027619Abstract: An airfoil for a turbine engine includes a structure having a cooling passage that has a generally radially extending cooling passageway arranged interiorly relative to an exterior surface of the structure. The cooling passageway includes multiple cooling slots extending there from toward the exterior surface and interconnected by a radially extending trench. The trench breaks the exterior surface, and the exterior surface provides the lateral walls of the trench. The airfoil is manufactured by providing a core having multiple generally axially extending tabs and a generally radially extending ligament interconnecting the tabs. The structure is formed about the core to provide the airfoil with its exterior surface. The ligament breaks the exterior surface to form the radially extending trench in the exterior surface of the structure.Type: ApplicationFiled: June 14, 2011Publication date: February 2, 2012Inventors: Jason Edward Albert, Atul Kohli, Eric L. Couch
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Patent number: 8092159Abstract: A cooling arrangement for a first stage nozzle of a turbine includes a slot formed in a forward face of the first stage nozzle, the slot opening in a direction facing a combustor transition piece and adapted to receive a flange portion of a seal extending between the first stage nozzle and the transition piece. The slot has a closed end formed with at least one cooling cavity provided with at least one cooling passageway extending between the cavity and an external surface of the first stage nozzle.Type: GrantFiled: March 31, 2009Date of Patent: January 10, 2012Assignee: General Electric CompanyInventor: Jaime Maldonado
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Publication number: 20110299999Abstract: Multi-component vane segment and method for forming the same. Assembly includes: positioning a pre-formed airfoil component (12) and a preformed shroud heat resistant material (18) in a mold, wherein the airfoil component (12) and the shroud heat resistant material (18) each comprises an interlocking feature (24); preheating the mold; introducing molten structural material (46) into the mold; and solidifying the molten structural material such that it interlocks the pre-formed airfoil component (12) with respect to the preformed shroud heat resistant material (18) and is effective to provide structural support for the shroud heat resistant material (18). Surfaces between the airfoil component (12) and the structural material (46), between the airfoil component (12) and the shroud heat resistant material (18), and between the shroud heat resistant material (18) and the structural material (46) are free of metallurgical bonds.Type: ApplicationFiled: June 7, 2010Publication date: December 8, 2011Inventor: Allister W. James
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Publication number: 20110274559Abstract: A turbine engine component includes a platform and one or more microcircuit cooling passages embedded within one or more walls of an airfoil portion of the component. Each microcircuit cooling passage terminates within the thickness of the platform so as to provide cooling to the initial 10% span of the airfoil portion. Each microcircuit cooling passage has an inlet for receiving cooling fluid, which inlet is also embedded within the platform.Type: ApplicationFiled: May 6, 2010Publication date: November 10, 2011Applicant: UNITED TECHNOLOGIES CORPORATIONInventors: Douglas C. Jenne, Matthew S. Gleiner, Matthew A. Devore
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Publication number: 20110274537Abstract: An impeller shroud is configured to receive an impeller. The impeller shroud establishes a plurality of air-bleed holes configured to communicate air with an impeller. The air-bleed holes are circumferentially distributed about the impeller shroud. The circumferential spacing between some adjacent air-bleed holes within the plurality of air-bleed holes is different than the circumferential spacing between other adjacent air-bleed holes within the plurality of air-bleed holes. A diffuser vane is configured to direct pressurized air to the combustor, the diffuser vanes are circumferentially distributed about the blade exducer. The circumferential spacing between some adjacent diffuser vanes within the plurality of diffuser vanes is different than the circumferential spacing between other adjacent diffuser vanes within the plurality of diffuser vanes.Type: ApplicationFiled: May 9, 2010Publication date: November 10, 2011Inventors: Loc Quang Duong, Xiaolan Hu
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Publication number: 20110268583Abstract: A method for machining an airfoil including a plurality of internal cooling channels is provided. The method includes selectively removing a pressure side section proximate to a trailing edge of the airfoil to expose a portion of the plurality of internal cooling channels proximate to the trailing edge of the airfoil. The method also includes machining the exposed portion of the plurality of internal cooling channels to a predefined shape.Type: ApplicationFiled: April 30, 2010Publication date: November 3, 2011Applicant: GENERAL ELECTRIC COMPANYInventor: Ronald Scott Bunker
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Publication number: 20110243755Abstract: A blade for a gas turbine includes an airfoil extending in radial direction of the turbine or longitudinal direction of the blade, respectively, between a platform and a blade tip. The airfoil is bordered across the airfoil by a leading edge and a trailing edge and has a suction side and a pressure side. At the trailing edge a first cooling passage runs parallel to the trailing edge from the platform to the blade tip in the interior of the airfoil. The cooling passage is supplied with a cooling air flow from the platform side, and from which cooling air is discharged through a plurality of cooling holes arranged all over the blade. For such a blade the cooling is optimized by providing a first cooling passage, the passage area of which is tapered in radial direction by between 35% and 59%.Type: ApplicationFiled: April 27, 2011Publication date: October 6, 2011Applicant: ALSTOM Technology Ltd.Inventors: Shailendra NAIK, Gaurav Pathak
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Patent number: 8020296Abstract: A method for the production of secondary fluid ducts for a turbomachine, especially for fluid supply or fluid removal to/from a flow-wetted surface of the turbomachine, with the secondary fluid ducts being produced by an electrochemical machining process.Type: GrantFiled: December 18, 2006Date of Patent: September 20, 2011Assignee: Rolls-Royce Deutschland Ltd & Co KGInventors: Volker Guemmer, Andreas Scholz
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Patent number: 8015705Abstract: A turbine rotor blade with a spar and shell construction, where the shell has an airfoil shape and is formed of two shell segments with an upper shell half and a lower shell half. The upper shell half is radially supported by a tip of the spar while the lower shell half is radially loaded by an attachment so that its load is not carried by the upper shell half and the tip of the spar in order to reduce overall stress levels.Type: GrantFiled: July 27, 2010Date of Patent: September 13, 2011Assignee: Florida Turbine Technologies, Inc.Inventors: Jack W Wilson, Jr., Wesley D Brown
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Publication number: 20110217181Abstract: Provided are gas turbine blades in which it is possible to simplify the formation of cooling channels provided inside the turbine blades while simultaneously avoiding loss of turbine blade strength and rigidity due to forming of the cooling channels. In a gas turbine blade (41), cooling channels (410) provided in the interior thereof include a plurality of straight channel-like base-side elongated holes (401a) that extend in a longitudinal direction at a base side of the turbine blade (41), a plurality of straight channel-like tip-side elongated holes (410b) that extend in a longitudinal direction at a tip side of the turbine blade (41), and a plurality of communicating hollow portions (410c) that are interposed at connection portions between the two types of elongated holes (410a, 410b) to individually allow the two types of elongated holes (410a, 410b) to communicate with each other and that have larger cross-sectional areas than the channel cross-sectional areas of both elongated holes (410a, 410b).Type: ApplicationFiled: April 25, 2011Publication date: September 8, 2011Applicant: MITSUBISHI HEAVY INDUSTRIES, LTD.Inventors: Satoshi Hada, Takahiko Imada, Tomofumi Shintani, Katsutoshi Ooe, Norifumi Hirata, Hiroshi Asano
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Publication number: 20110217180Abstract: Provided are gas turbine blades in which it is possible to simplify the formation of cooling channels provided inside the turbine blades while simultaneously avoiding loss of turbine blade strength and rigidity due to forming of the cooling channels. In a gas turbine blade, cooling channels provided in the interior thereof include a plurality of straight channel-like base-side elongated holes that extend in a longitudinal direction at a base side of the turbine blade, a plurality of straight channel-like tip-side elongated holes that extend in a longitudinal direction at a tip side of the turbine blade, and a plurality of communicating hollow portions that are interposed at connection portions between the two types of elongated holes to individually allow the two types of elongated holes to communicate with each other and that have larger cross-sectional areas than the channel cross-sectional areas of both elongated holes.Type: ApplicationFiled: December 6, 2010Publication date: September 8, 2011Applicant: MITSUBISHI HEAVY INDUSTRIES, LTD.Inventors: Satoshi Hada, Takahiko Imada, Tomofumi Shintani, Katsutoshi Ooe, Norifumi Hirata, Hiroshi Asano
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Publication number: 20110206536Abstract: A turbine blade for a turbine engine includes main coolant passageways which extend through the turbine blade to cool the blade. A tip coolant passageway conveys coolant from a location adjacent the base of the blade directly to the tip of the blade to provide cooling fluid directly to the tip of the blade. This ensures that the coolant arriving at the tip of the blade is at a relatively low temperature and can therefore provide effective cooling of the material located at the tip of the blade.Type: ApplicationFiled: February 25, 2010Publication date: August 25, 2011Inventor: Dipankar PAL
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Publication number: 20110192024Abstract: Fabricating a turbine component (50) by casting a core structure (30), forming an array of pits (24) in an outer surface (32) of the core structure, depositing a transient liquid phase (TLP) material (40) on the outer surface of the core structure, the TLP containing a melting-point depressant, depositing a skin (42) on the outer surface of the core structure over the TLP material, and heating the assembly, thus forming both a diffusion bond and a mechanical interlock between the skin and the core structure. The heating diffuses the melting-point depressant away from the interface. Subsurface cooling channels (35) may be formed by forming grooves (34) in the outer surface of the core structure, filling the grooves with a fugitive filler (36), depositing and bonding the skin (42), then removing the fugitive material.Type: ApplicationFiled: February 5, 2010Publication date: August 11, 2011Inventor: David B. Allen
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Patent number: 7992272Abstract: A method is provided for milling a thermal barrier coated metal part. This method includes selectively removing a portion of a ceramic coating using a mechanical cutting tool, thereby forming a counterbore, and machining the metal part through the counterbore. A drilling head for drilling thermal barrier coated metal parts is also provided. The drilling head comprises a mechanical cutting tool, which is operable to mill through ceramic, and an electrode for electrical discharge machining. The electrode may be used to mill the metal part, and may be interchangeable with the mechanical cutting tool.Type: GrantFiled: May 29, 2007Date of Patent: August 9, 2011Assignee: Metem CorporationInventors: Joseph W. Janssen, John Malek
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Publication number: 20110150666Abstract: A turbine blade having an internal skeleton having a plurality of internal ribs that form a plurality of open cooling channels; an internal environmental coating applied to the internal skeleton; an outer wall applied about the open cooling channels of the internal skeleton to form a near wall circuit of cooling channels; and an external environmental coating applied to the outer wall wherein the internal environmental coating is different from the external environmental coating.Type: ApplicationFiled: April 30, 2010Publication date: June 23, 2011Inventors: BRIAN THOMAS HAZEL, DOUGLAS GERARD KONITZER, MICHAEL HOWARD RUCKER, JOHN DOULGAS EVANS, SR.
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Patent number: 7950147Abstract: A method for producing gas turbine components, in particular blades, blade sections or rotors with integral blades for a jet engine is provided. The method comprises at least the following steps: preparation of at least one metallic powder and at least one expanding agent; mixing of the metallic powder or each metallic powder with the expanding agent or each expanding agent; compression of the resultant mixture to form at least one semi-finished product; expansion of the semi-finished product or each semi-finished product by heating to achieve a defined degree of expansion; termination of the expansion process by cooling, once the defined degree of expansion has been reached.Type: GrantFiled: November 24, 2004Date of Patent: May 31, 2011Assignee: MTU Aero Engines GmbHInventors: Reinhold Meier, Erich Steinhardt
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Publication number: 20110038708Abstract: An airfoil is provided and includes an airfoil body having a pressure surface extendable between radial ends and a fluid path in an airfoil interior defined therein. The pressure surface is formed to further define a passage by which coolant is deliverable from the fluid path in the airfoil interior, in a perimetric direction from the pressure surface for the purpose of cooling a portion on the surface of the radial end.Type: ApplicationFiled: August 11, 2009Publication date: February 17, 2011Applicant: GENERAL ELECTRIC COMPANYInventor: Jeffrey John Butkiewicz
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Publication number: 20110016884Abstract: To provide a cover of a cooling passage that forms a cooling passage for supplying cooling air to a turbine rotor blade at the last stage via inside of a disk of a turbine, and the cover comprises: a cylindrical cover portion that covers a cavity provided in a annular pattern in an outer circumference of the disk in a mode where a first passage opened from inside of the disk to the cavity and a second passage opened from a cooling passage of the turbine rotor blade at the last stage to the cavity are connected to each other; and a flexible portion that is formed integrally with the cover portion and allows flexure in an axial direction of the turbine.Type: ApplicationFiled: January 15, 2009Publication date: January 27, 2011Applicant: MITSUBISHI HEAVY INDUSTRIES, LTD.Inventors: Shinya Hashimoto, Kenichi Arase
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Publication number: 20100329888Abstract: A method is disclosed that includes providing a turbomachinery blade having an airfoil connected to a platform in a root region of the turbomachinery blade. The airfoil has a trailing edge extending from the root region to a tip distal from the root region. The method further includes forming a blind relief hole in the platform proximate the trailing edge of the airfoil, and forming a plurality of cooling holes in the platform.Type: ApplicationFiled: April 20, 2010Publication date: December 30, 2010Inventors: Gregory M. Nadvit, Andrew D. Williams, Leone J. Tessarini, Michel P. Amal
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Publication number: 20100329835Abstract: An apparatus for a gas turbine engine includes an airfoil, a metering opening for metering a cooling fluid, a cutback slot configured to deliver the cooling fluid from the metering opening, and a cooling hole. The airfoil defines a trailing edge, opposite first and second faces, and a mean camber line. The cutback slot is defined along the first face of the airfoil adjacent to the trailing edge and offset from the mean camber line of the airfoil. The cooling hole has an outlet that is located at the trailing edge and substantially aligned with the mean camber line of the airfoil. The cooling hole delivers a portion of the cooling fluid from the metering opening.Type: ApplicationFiled: June 26, 2009Publication date: December 30, 2010Applicant: UNITED TECHNOLOGIES CORPORATIONInventors: Brandon W. Spangler, Sam Draper
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Publication number: 20100322783Abstract: A rotor blade or a stator blade for rotary machinery is disclosed where the rotor/stator blade comprises a number of thin blade plates. A plurality of the rotor/stator blades are provided with at least one hole forming at least one supply duct when the blade plates are stacked on top of each other to form the rotor/stator blade. The blade plates are joined together by means of sintering such that a solid rotor/stator blade with an outer surface is formed.Type: ApplicationFiled: June 17, 2010Publication date: December 23, 2010Applicant: NEBB TECHNOLOGY ASInventor: Inge Tronstad
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Patent number: 7841083Abstract: A method of producing cooling fluid discharge orifices in the wall of a part manufactured by the technique of lost wax casting in which a pattern of the part is produced in a wax mold, the orifices each having a first portion emerging at the external surface of the wall is disclosed. The method includes making cavities in the wax pattern that correspond to the first portions of said orifices of the part. Thus, it is possible to produce cooling air discharge orifices without sharp edges.Type: GrantFiled: January 29, 2007Date of Patent: November 30, 2010Assignee: SNECMAInventors: Thierry Henri Raymond Alaux, Patrick Emilien Paul Emile Huchin, Patrice Jean-Marc Rosset, Boris Soulalioux
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Patent number: 7805822Abstract: A process which uses an air jet containing non-abrasive particulate media at a low pressure which selectively removes thermal barrier coatings from components without damaging the metallic substrate. This process selectively removes thermal barrier coatings from the cooling holes of components.Type: GrantFiled: December 15, 2003Date of Patent: October 5, 2010Assignee: Turbocombustor Technology, Inc.Inventor: Gary Lynn Hanley
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Publication number: 20100247329Abstract: A turbine blade assembly for a gas turbine includes a spar with raised ribs, a spacer with a plurality of protrusions mounted around the spar, and an outer shell mounted around the spacer. The protruding portions on the spacer surround the raised ribs on the spar. The protruding portions of the spacer act to space the interior surfaces of the outer shell away from the spar to provide a thermal insulation layer of cooling air.Type: ApplicationFiled: March 30, 2009Publication date: September 30, 2010Inventor: Victor MORGAN
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Publication number: 20100239409Abstract: A method of using a film-cooling insert for a turbine airfoil is disclosed. The method includes forming an airfoil sidewall having a film-cooling hole that extends between an airfoil cooling circuit and an airfoil surface. The method also includes forming a film-cooling insert and disposing the film-cooling insert in the film-cooling hole. A method of reconstructing a film-cooling insert for a turbine airfoil is also disclosed. The method includes removing a remnant of a film-cooling insert from a film-cooling hole of a turbine airfoil. The method also includes disposing a second film-cooling insert in the film-cooling hole.Type: ApplicationFiled: March 18, 2009Publication date: September 23, 2010Applicant: GENERAL ELECTRIC COMPANYInventor: Samuel David Draper
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Publication number: 20100229388Abstract: A wall in which there is formed at least one cooling channel, said wall being cooled by cool air flowing in the channel, the channel comprising a hole and a diffusion portion, the hole opening out at one end into the inside surface of the wall, and at its other end into the diffusion portion where it forms an orifice, the diffusion portion flaring around said orifice and opening out into the outside surface of the wall, the diffusion portion having a bottom whose front end is substantially plane, sloping, and extending in front of the orifice, and also having a margin extending behind, round the sides, and in front of the orifice, said margin joining the sides of the front end. A method and an electrode for making such a cooling channel. A turbomachine blade presenting such a wall.Type: ApplicationFiled: May 26, 2010Publication date: September 16, 2010Applicant: SNECMAInventors: Emmanuel Pierre CAMHI, Laurent CROUILBOIS, Mareix Jean Pierre, Didier PASQUIET
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Patent number: 7784183Abstract: A manufacturing system and manufacturing method for adjusting the performance of manufacturing operations or steps in manufacturing components having three-dimensional external structural characteristics. An embodiment of the system broadly comprises: (a) a plurality of manufacturing operations for processing a component having three-dimensional external structural characteristics; (b) at least one analytical device for analyzing at least one characteristic of the component after the performance of one or more manufacturing operations to generate a component data set; (c) at least one data storage device for storing the generated component data sets and for providing at least a relevant portion of accumulated component data; and (d) a communication mechanism for transmitting at least a relevant portion of accumulated component data to one or more manufacturing operations so that the performance thereof can be adjusted in response to the transmitted portion of accumulated component data.Type: GrantFiled: June 9, 2005Date of Patent: August 31, 2010Assignee: General Electric CompanyInventors: Todd Jay Rockstroh, James Joseph Hoffman
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Patent number: 7772314Abstract: Process of masking cooling holes of a gas turbine component with an external surface, comprising a cavity and a plurality of cooling holes before coating the gas turbine component, comprising the steps of first applying a mask material to the cooling holes so that the cooling holes are filled at least closest to the external surface, whereby the mask material contains a substance which fluoresces under ultraviolet light and a filler material. Then the mask material within the cooling holes is thickening. An inspection using ultraviolet light to locate any unwanted residual mask material on the external surface is carried out and unwanted residual mask material is removed before the coating is applied to the external surface of the component and the masked cooling holes. In the end the mask material is removed from the cooling holes.Type: GrantFiled: August 10, 2006Date of Patent: August 10, 2010Assignee: Alstom Technology LtdInventors: John Fernihough, Andreas Boegli, Alexander Stankowski
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Patent number: 7731481Abstract: A turbine engine component has an airfoil portion with a pressure side wall and a suction side wall and a cooling system. The cooling system has at least one cooling circuit disposed longitudinally along the airfoil portion. Each cooling circuit has a plurality of staggered internal pedestals for increasing heat pick-up.Type: GrantFiled: December 18, 2006Date of Patent: June 8, 2010Assignee: United Technologies CorporationInventors: Francisco J. Cunha, Edward F. Pietraszkiewicz
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Publication number: 20100126960Abstract: A cooling microcircuit for use in a turbine engine component is provided. The cooling microcircuit has at least one leg through which a cooling fluid flows. A plurality of cast vortex generators are positioned within the at least one leg to improve the cooling effectiveness of the cooling microcircuit.Type: ApplicationFiled: January 28, 2010Publication date: May 27, 2010Applicant: UNITED TECHNOLOGIES CORPORATIONInventor: Francisco J. Cunha
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Publication number: 20100111671Abstract: A turbine cooling component comprising a circumferential leading edge, a circumferential trailing edge, a pair of spaced and opposed side panels connected to the leading and trailing edges, an arcuate base connected to the trailing and leading edges having a fore portion, a midsection portion, an aft portion, opposed side portions, an outer surface partially defining a cavity operative to receive pressurized air, and an arcuate inner surface in contact with a gas flow path of a turbine engine, a first side cooling air passage in the base extending along the first side portion from the fore portion to the aft portion, and a fore cooling air passage in the fore portion of the base communicative with the side cooling air passage and the cavity, operative to receive the pressurized air from the cavity.Type: ApplicationFiled: November 5, 2008Publication date: May 6, 2010Applicant: GENERAL ELECTRIC COMPANYInventors: Kevin Thomas McGovern, Bryan David Lewis, Zachary James Taylor
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Publication number: 20100098526Abstract: A turbine engine airfoil includes an airfoil structure having a side with an exterior surface. The structure includes a cooling passage extending a length within the structure and providing a convection surface facing the side. The convection surface is twisted along the length, which varies a heat transfer rate between the exterior surface and the convection surface along the length. In one example, the cooling passage is provided by a refractory metal core that is used during the airfoil casting process. The core includes multiple legs joined by a connecting portion. At least one of the legs is twisted along its length. The legs are deformed toward one another opposite the connecting portion to provide a desired core shape that corresponds to the shape of the cooling passage. Accordingly, the cooling passage provides desired cooling of the airfoil.Type: ApplicationFiled: October 16, 2008Publication date: April 22, 2010Inventor: Justin D. Piggush
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Publication number: 20100054955Abstract: A rotary blade, such as a turbine blade for a gas turbine engine, has an aerofoil portion with a tip partly shrouded by winglets. A gutter extends across the radially outer face of the tip to leave upstands. Cooling air feed galleries are drilled into each upstand, from the trailing edge, toward the upper end of a cooling air feed void, which is spaced from the trailing edge. Cooling passages are drilled from the winglet edges to the gallery. Cooling air supplied through the void passes along the gallery, through the passages and leaves the blade at the cooling holes. This allows cooling to be provided near the trailing edge of the tip without requiring the geometry around the trailing edge to be thickened to accommodate a cooling air void.Type: ApplicationFiled: July 15, 2009Publication date: March 4, 2010Applicant: ROLLS-ROYCE, PLCInventors: Caner H. Helvaci, Roderick M. Townes, Stephen Diamond
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Publication number: 20100024216Abstract: A method of fabricating a rotor blade is provided. The method includes forming at least one passageway within the rotor blade, wherein the passageway extends substantially radially from a root of the rotor blade to a tip of the rotor blade, and coupling a shroud to the tip of the rotor blade. The shroud includes at least one substantially radially-outward extending wall that at least partially defines an outer plenum that is radially outward from at least the shroud, wherein the outer plenum is in flow communication with the passageway.Type: ApplicationFiled: July 29, 2008Publication date: February 4, 2010Inventors: Donald Brett DeSander, James Earl Kopriva, Robert Francis Manning, Robert Edward Athans, Mark Douglas Gledhill
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Patent number: 7640661Abstract: A process for manufacturing reinforcing parts (14) for leading edges (10) and/or trailing edges (12) of fan blades (1) is described. In particular, this invention relates to the use of the SPF/DB (<<Super Plastic Forming/Diffusion Bonding >>) process for this type of non-hollow part.Type: GrantFiled: March 4, 2005Date of Patent: January 5, 2010Assignee: SnecmaInventors: Jean-Louis Despreaux, Jean-Michel Franchet, Philippe Joffroy, Gilles Klein, Stéphane Leveque, Daniel Lhomme, Alain Lorieux
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Publication number: 20090324421Abstract: A turbine blade is provided. The turbine blade includes a support structure and a shell which surrounds the support structure and which is connected to and at a distance from the support structure by at least one spacing element. For example the spacing element can be a solder globule in order to form a space through which a cooling medium can flow between the support structure and the shell. A method for the production of a turbine blade having a support structure and a shell which surrounds the support structure and which is connected to and at a distance from the support structure is also provided. The shell is soldered to the support structure in at least at one place of the support structure in order to connect the shell to and at a distance from the support structure.Type: ApplicationFiled: January 14, 2008Publication date: December 31, 2009Inventors: Fathi Ahmad, Scarlett Fajardo-Reina, Markus Gill, Stefan Werner Kiliani, Silvio-Ulrich Martin, Ralf Müsgen, Oliver Schneider
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Publication number: 20090297361Abstract: Contaminant build-up and cooling airflow looses are reduced in a turbine airfoil by joining root and airfoil cooling air passages thereof with a transition passage.Type: ApplicationFiled: January 22, 2008Publication date: December 3, 2009Inventors: Matthew T. Dahmer, Alexander V. Staroselsky
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Patent number: 7625178Abstract: An air-cooled turbine blade is provided having an airfoil shape defined by a convex suction side wall, a concave pressure side wall, a leading edge, a trailing edge, a root and a tip, the walls and the tip each including an interior surface that defines an interior with the root, the trailing edge including a plurality of slots formed thereon. The blade includes a suction side flow circuit formed therein comprising a forward and an aft flow circuit. The blade also includes a tip flow circuit extending along the tip interior surface to at least one of the trailing edge slots and including a first and a second opening, the first opening in flow communication with the suction side forward flow circuit outlet, and the second opening in flow communication with a cross-over hole of the suction side aft flow circuit. Methods of manufacturing the blade are also provided.Type: GrantFiled: August 30, 2006Date of Patent: December 1, 2009Assignee: Honeywell International Inc.Inventors: Mark C. Morris, George W. Wolfmeyer, Luis A. Tapia, Vivek Agarwal, Kin Poon, Jason C. Smoke, William C. Baker, Henry M. Armstrong
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Publication number: 20090235525Abstract: A method for making a gas turbine component (100). A central core (20) is positioned to occupy a space that will define a central channel (42), and an outer channel core (30) is positioned spaced apart from the central core (20). A mold (35) is formed around the central core (20) and the outer channel core (30), so that an exterior wall (32) contacts the mold (35). A substrate material, such as a metal alloy (247) in liquid form, is added to the mold (35) to form an internal volume (41) of the component (100). The central core (20) and the outer channel core (30) are removed, and interconnect channels (44) are formed between the thus-formed central channel (42) and the inner portion (49) of the outer channel (62) thus far formed. A preform (55) is placed into the inner portion (49) and may have a desired outer surface (57) shape. An overlay material is applied to form an outer layer (60), thus defining the remainder of the outer channel (62), which is obtained upon removal of the preform (55).Type: ApplicationFiled: March 21, 2008Publication date: September 24, 2009Applicant: Siemens Power Generation, Inc.Inventors: Douglas J. Arrell, Allister W. James, Anand A. Kulkarni
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Patent number: 7584538Abstract: A method of constructing a turbine blade for a gas turbine engine is provided. The method includes casting the blade, forming a plurality of spaced-apart notches in the airfoil proximate the tip on the pressure sidewall and forming at least one hole in each tip shelf communicating with the interior void of the airfoil for channeling cooling air from the interior void of the airfoil to thereby form a squealer tip.Type: GrantFiled: June 21, 2007Date of Patent: September 8, 2009Assignee: General Electric CompanyInventor: Ching-Pang Lee
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Publication number: 20090193657Abstract: The present invention is a vane for us in a gas turbine engine, in which the vane is made of an exotic, high temperature material that is difficult to machine or cast. The vane includes a shell made from either Molybdenum, Niobium, alloys of Molybdenum or Niobium (Columbium), Oxide Ceramic Matrix Composite (CMC), or SiC-SiC ceramic matrix composite, and is formed from a wire electric discharge process. The shell is positioned in grooves between the outer and inner shrouds, and includes a central passageway within the spar, and forms a cooling fluid passageway between the spar and the shell. Both the spar and the shell include cooling holes to carry cooling fluid from the central passageway to an outer surface of the vane for cooling. This cooling path eliminates a serpentine pathway, and therefore requires less pressure and less amounts of cooling fluid to cool the vane.Type: ApplicationFiled: January 16, 2009Publication date: August 6, 2009Applicant: Florida Turbine Technologies, Inc.Inventors: Jack W. Wilson, JR., Wesley D. Brown