And Cooling Patents (Class 60/806)
  • Patent number: 10808537
    Abstract: In order to compensate for the increase in a leakage rate of ventilation air originating in a circuit via a labyrinth seal that gradually wears at the bottom of a ventilation chamber, and the reduction of a purge rate via an opening upstream from the chamber, a second injection opening is added to a main injection opening. The second injection opening communicates with the chamber via a second labyrinth seal that wears at the same time as the first, and thus gradually increases the air flow rate passing through the injector in order to compensate for the additional leakage rate through the first injector and maintain the ventilation at its initial level.
    Type: Grant
    Filed: January 28, 2019
    Date of Patent: October 20, 2020
    Assignee: SAFRAN AIRCRAFT ENGINES
    Inventors: Mathieu Jean Pierre Trohel, Paco Maurer
  • Patent number: 10731855
    Abstract: Combustors and panels for use in combustor sections of gas turbine engines, the panels having a panel body having a first wall with a first sidewall and a second sidewall extending each extending from the first wall in the same direction at edges of the first wall and a plurality of cavity walls arranged between the first sidewall and the second sidewall and extending from the panel in the direction of the first and second sidewalls, wherein a plurality of cooling cavities are defined by the cavity walls, the first wall, and at least one of the first sidewall and the second sidewall, wherein each cooling cavity extends from an inlet hole to an outlet hole, wherein the outlet hole is formed in the first wall.
    Type: Grant
    Filed: August 23, 2017
    Date of Patent: August 4, 2020
    Assignee: RAYTHEON TECHNOLOGIES CORPORATION
    Inventors: Mark F. Zelesky, Carey Clum
  • Patent number: 10724440
    Abstract: Cooling systems for high pressure compressor systems are provided. The cooling systems may comprise tangential on board injectors (“TOBIs”). The TOBIs may comprise one or more fluid channels configured to conduct cooling fluid flow to components of the compressor, including, for example, disk-hub portions of the compressor. In this regard, the TOBI may be configured to exhaust cooling air in a manner such that the exhausted air has a similar linear velocity of the disk-hub portion. The cooling air may also be exhausted in a manner that is substantially parallel to the disk-hub portion.
    Type: Grant
    Filed: January 24, 2019
    Date of Patent: July 28, 2020
    Assignee: RAYTHEON TECHNOLOGIES CORPORATION
    Inventor: Paul Hiester
  • Patent number: 10697312
    Abstract: A guide vane for a twin-spool aircraft turbomachine has an aerodynamic part that includes an internal lubricant cooling passage extending along a principal lubricant flow direction. The aerodynamic part is made in a single piece and also includes heat transfer fins arranged in the passage connecting the intrados and extrados walls and extending approximately parallel to the direction, these fins being distributed in successive rows along the principal direction and made such that for two rows of staggered directly consecutive fins, a first row includes fins forming a positive acute angle A1 with a dummy reference plane, while a second row includes fins forming a negative acute angle A2 with this plane.
    Type: Grant
    Filed: March 7, 2018
    Date of Patent: June 30, 2020
    Assignee: SAFRAN AIRCRAFT ENGINES
    Inventors: Mohamed-Lamine Boutaleb, Fabien Roger Gaston Caty, Sebastien Vincent Francois Dreano, Thierry Georges Paul Papin, Christophe Marcel Lucien Perdrigeon, Cedric Zaccardi
  • Patent number: 10669945
    Abstract: According to an aspect, a system includes a starter air valve in fluid communication with an air turbine starter to drive motoring of a gas turbine engine responsive to a compressed air flow from a compressed air source. The system also includes a variable-position electromechanical device operable to adjust positioning of the starter air valve and a discrete-position electromechanical device operable to adjust positioning of the starter air valve and limit a motoring speed of the gas turbine engine below a resonance speed of the gas turbine engine responsive to a pulse width modulation control based on a failure of the variable-position electromechanical device.
    Type: Grant
    Filed: February 6, 2017
    Date of Patent: June 2, 2020
    Assignee: RAYTHEON TECHNOLOGIES CORPORATION
    Inventors: Michael D. Greenberg, David Gelwan, Jesse W. Clauson, Robert Goodman, John M. Dehais, Myles R. Kelly
  • Patent number: 10563584
    Abstract: Combustor panels including panel bodies with first and second sides, a pin array extending from the first side, wherein each pin extends a first height, has a pin diameter, and is separated from adjacent pins by a pin array separation distance. A structural protrusion extends from the first side. No pins of the pin array are located within a flashing distance that is equal to a protrusion separation distance plus half of the pin diameter, wherein a location of the pin is measured from a center point of the pin to a closest point on the exterior surface of the structural protrusion. At least one pin array extension is integrally formed with the structural protrusion, the pin array extension extending along the first side to a position that replaces a pin of the pin array that would be within the flashing distance.
    Type: Grant
    Filed: October 27, 2017
    Date of Patent: February 18, 2020
    Assignee: UNITED TECHNOLOGIES CORPORATION
    Inventors: Jason Shenny, John J. Rup, Jr., Robert M. Sonntag, James B. Hoke, San Quach, Richard G. Ullrich
  • Patent number: 10458332
    Abstract: A high pressure compressor has a downstream most end. A housing surrounds the compressor section and a turbine section. A low pressure turbine has a downstream most end. A first tap selectively taps high pressure cooling air from a location downstream of the downstream most end in the high pressure compressor and passes the high pressure cooling air through a heat exchanger. A second tap taps compressed air from a location upstream of the downstream most end in the high pressure compressor, and passes air over the heat exchanger, cooling the high pressure cooling air. A chamber is defined between the core engine housing and a nacelle airflow wall, and the second tap air flows through the chamber. The second tap air moves from the chamber into a core engine flow at a location downstream of the downstream most end of the low pressure turbine.
    Type: Grant
    Filed: January 17, 2017
    Date of Patent: October 29, 2019
    Assignee: United Technologies Corporation
    Inventors: Frederick M. Schwarz, Nathan Snape
  • Patent number: 10443406
    Abstract: Components for gas turbine engines are provided. The components include an airfoil extending between a first end and a second end and from a leading edge to a trailing edge. The airfoil has a pressure side and a suction side. The airfoil has a non-leading edge stagnation line that transitions from a location along a leading edge surface of the airfoil at the first end to a location on the airfoil on the pressure side of the airfoil at the second end. A leading edge feed cavity is located in the leading edge of the airfoil and extending between the first end and the second end of the airfoil. A stagnation impingement cavity is formed extending along a surface of the airfoil that aligns with the non-leading edge stagnation line.
    Type: Grant
    Filed: January 31, 2018
    Date of Patent: October 15, 2019
    Assignee: UNITED TECHNOLOGIES CORPORATION
    Inventors: Tracy A. Propheter-Hinckley, Dominic J. Mongillo, Jr.
  • Patent number: 10371055
    Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. The cooling compressor is connected to be driven with at least one rotor in the main compressor section. A source of pressurized air is selectively sent to the cooling compressor to drive a rotor of the cooling compressor to rotate, and to in turn drive the at least one rotor of the main compressor section at start-up of the gas turbine engine. An intercooling system is also disclosed.
    Type: Grant
    Filed: August 27, 2015
    Date of Patent: August 6, 2019
    Assignee: United Technologies Corporation
    Inventors: Nathan Snape, Gabriel L. Suciu, Brian D. Merry
  • Patent number: 10352196
    Abstract: To remove water accumulated in a cooling air system while preventing performance loss of a gas turbine, the gas turbine is provided with a cooling air system that connects an intermediate stage or the exit of a compressor to a turbine to supply compressed air bled from the compressor to the turbine, a TCA cooler, i.e., a heat exchanger, that cools the compressed air on the route of the cooling air system, and a drain water discharge valve, i.e., a first drain water discharge valve and a second drain water discharge valve disposed downstream of the compressed air of the TCA cooler, wherein at least for a predetermined period of time after a rated speed of the gas turbine at start up has been reached, the drain water discharge valve is set to an open state, and thereafter, the drain water discharge valve is set to a closed state.
    Type: Grant
    Filed: January 21, 2015
    Date of Patent: July 16, 2019
    Assignee: MITSUBISHI HITACHI POWER SYSTEMS, LTD.
    Inventors: Shinichi Yamazaki, Hirotaka Nakasako
  • Patent number: 10202867
    Abstract: A flow transfer apparatus for transferring cooling flow from a primary gas flowpath to a turbine rotor. The apparatus includes a first supply plenum communicating with the primary gas flowpath and first inducers, the first inducers configured to accelerate a first fluid flow from the first supply plenum and discharge it toward the rotor with a tangential velocity; a second supply plenum communicating with the primary gas flowpath and second inducers, the second inducers configured to accelerate a second fluid flow from the second supply plenum towards the rotor with a tangential velocity; and a cooling modulation valve operable to selectively permit or block the second fluid flow from the primary gas flowpath to the second supply plenum. The valve includes a flow control structure disposed in the primary gas flowpath and an actuation structure extending to a location radially outside of a casing defining the primary gas flowpath.
    Type: Grant
    Filed: March 14, 2014
    Date of Patent: February 12, 2019
    Assignee: General Electric Company
    Inventors: John Joseph Rahaim, Daniel John Fusinato, Scott David Hunter, Joel Francis Kirk, Michael Robert Melzak, Curtis Walton Stover
  • Patent number: 10125686
    Abstract: A turbine engine assembly is provided. The assembly includes a booster compressor that discharges a flow of first compressed air at a first temperature, and a high-pressure compressor (HPC). The HPC receives a first portion of the flow of first compressed air and discharges the first portion at a second temperature higher than the first temperature such that a flow of second compressed air is formed. The assembly also includes a heat exchanger (HEX) that receives a second portion of the flow of first compressed air from the booster compressor and a first portion of the flow of second compressed air from the HPC such that heat is transferred therebetween. The HEX discharges the first portion of the flow of second compressed air at a third temperature lower than the second temperature and higher than the first temperature such that a flow of cooled bleed air is formed.
    Type: Grant
    Filed: December 3, 2015
    Date of Patent: November 13, 2018
    Assignee: GENERAL ELECTRIC COMPANY
    Inventors: Carlos E. Diaz, John Joseph Rahaim, Jorge de Luis, Noel Istvan Macsotai
  • Patent number: 9850819
    Abstract: A gas turbine engine comprises a main compressor section having a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses ng air downstream of the heat exchanger, and delivers air into the high pressure turbine. The heat exchanger has at least two passes, with one of the passes passing air radially outwardly, and a second of the passes returning the air radially inwardly to the compressor. An intercooling system for a gas turbine engine is also disclosed.
    Type: Grant
    Filed: April 24, 2015
    Date of Patent: December 26, 2017
    Assignee: United Technologies Corporation
    Inventors: Gabriel L. Suciu, Jesse M. Chandler, Joseph Brent Staubach, Brian D. Merry, Wesley K. Lord
  • Patent number: 9719372
    Abstract: A gas turbomachine includes a casing assembly surrounding a portion of the gas turbomachine and a counter-flow cooling system arranged within the casing. The counter-flow cooling system is configured and disposed to guide cooling fluid through the casing assembly in a first axial direction and return cooling fluid through the casing assembly in a second axial direction that is opposite the first axial direction.
    Type: Grant
    Filed: May 1, 2012
    Date of Patent: August 1, 2017
    Assignee: General Electric Company
    Inventors: Henry Grady Ballard, Jr., Kenneth Damon Black, John David Memmer
  • Patent number: 9556794
    Abstract: The chamber (4) of the turbine engine (1) comprises a continuous detonation wave engine (6) provided with an annular detonation chamber (7) and associated means (8, 9) that can be used to generate a continuous production of hot gases from a detonation mixture of fuel and air. The continuous detonation wave engine (6) is arranged such as to form, from a flow of incoming air (E), a first flow (F1) which enters the detonation chamber (7) and which is used by the engine (6) and a second flow (F2) which bypasses the chamber. The turbine engine (1) also includes auxiliary means (10) for mixing the hot gases (F3) leaving the detonation chamber (7) with the second flow of air (F2) before directing same towards the turbine (5). A plurality of detonation chambers (7) are arranged concentrically to one another relative to the axis of the turbine engine.
    Type: Grant
    Filed: May 9, 2012
    Date of Patent: January 31, 2017
    Assignee: MBDA FRANCE
    Inventors: François Falempin, Bruno Le Naour
  • Patent number: 9488057
    Abstract: A micro jet gas film generation apparatus aims to eject a gas to a work object which is desired for cooling or insulating from heat. The micro jet gas film generation apparatus has a spout formed at a diameter of 5-100 ?m to generate a gas film on the work object. As the diameter of the spout of the micro jet gas film generation apparatus is small, the micro jet gas film generated from the spout cannot produce a large eddy due to the lack of sufficient energy, hence can maintain a thin film after a long distance ejection to improve cooling and heat insulation performance. Moreover, due to the small diameter of the spout, it also consumes less gas and can reduce the amount of gas required.
    Type: Grant
    Filed: June 10, 2014
    Date of Patent: November 8, 2016
    Assignee: AERONAUTICAL SYSTEMS RESEARCH DIVISION NATIONAL CHUNG SHAN INSTITUTE OF SCIENCE AND TECHNOLOGY
    Inventors: Dun-Zen Jeng, Chao-Ching Wang, Ting-Hua Chien, Chi Kao
  • Patent number: 9388739
    Abstract: A buffer air cooler system for gas turbine engines disposed in a bypass duct of the engine, includes a housing for containing the buffer air cooler therein and an inlet portion attached to the housing. In one embodiment, the inlet portion has a double-skin configuration in at least one region of a top, bottom and sides of the inlet portion.
    Type: Grant
    Filed: May 2, 2012
    Date of Patent: July 12, 2016
    Assignee: PRATT & WHITNEY CANADA CORP.
    Inventors: Remo Marini, Patrick Germain, Ljubisa Vrljes
  • Patent number: 9341119
    Abstract: In one aspect, a turbine engine is provided. The engine includes a fan, a core engine case configured to receive a core airflow, and an inner fan case. An inner air bypass is defined between the core engine case and the inner fan case, and the engine further includes an outer fan case, an outer air bypass defined between the inner fan case and the outer fan case, a first heat exchanger arranged in the core engine case, the first heat exchanger providing heat exchange between the liquid and the core airflow to cool the core airflow, and a second heat exchanger arranged in the outer air bypass. The second heat exchanger is fluidly coupled to the first heat exchanger to receive the cooled core airflow. The second heat exchanger provides heat exchange between cooled core airflow and the second portion of airflow to further cool the cooled core airflow.
    Type: Grant
    Filed: July 3, 2014
    Date of Patent: May 17, 2016
    Assignee: HAMILTON SUNDSTRAND CORPORATION
    Inventor: William E. Rhoden
  • Patent number: 9267407
    Abstract: An engine exhaust system attachment is provided that includes a housing a fan, a filter, and a conduit. The housing is configured to enclose an exhaust turbine, an exhaust manifold, and at least a portion of an exhaust pipe. The exhaust turbine, the exhaust manifold, and the exhaust pipe are connected to receive exhaust gas from an engine and are mounted to an engine frame of a device. The fan is configured for mounting to the device to move air. The filter is configured for mounting to the fan to receive the moved air and to provide filtered air. The conduit is configured for connecting the filter to the housing to provide the filtered air to the housing.
    Type: Grant
    Filed: July 8, 2013
    Date of Patent: February 23, 2016
    Assignee: South Dakota Board of Regents
    Inventors: Daniel S. Humburg, Kevin J. Dalsted, ZhengRong Gu, Joseph P. Polin
  • Patent number: 9255492
    Abstract: A gas turbine engine has, in flow series, a compressor section, a combustor, and a turbine section. The gas turbine engine further has a system (i) for cooling the turbine section and (ii) for providing tip clearance control between turbine blades of the turbine section and a plurality of circumferentially distributed segments which form an annular shroud surrounding the outer tips of the turbine blades. The system includes a turbine section cooling sub-system which diverts a first cooling air flow received from the compressor section to a heat exchanger and then to the turbine section to cool components thereof. The first cooling air flow by-passes the combustor and is cooled in the heat exchanger. The turbine section cooling subsystem has a first valve arrangement which regulates the first cooling air flow. The system further includes a tip clearance control sub-system which supplies a second cooling air flow to an engine case to which the segments are mounted.
    Type: Grant
    Filed: November 20, 2012
    Date of Patent: February 9, 2016
    Assignee: ROLLS-ROYCE plc
    Inventor: Marko Bacic
  • Patent number: 9206699
    Abstract: A cooling system is provided for a transition (420) of a gas turbine engine (410). The cooling system includes a cowling (460) configured to receive an air flow (111) from an outlet of a compressor section of the gas turbine engine (410). The cowling (460) is positioned adjacent to a region of the transition (420) to cool the transition region upon circulation of the air flow within the cowling (460). The cooling system further includes a manifold (121) to directly couple the air flow (111) from the compressor section outlet to an inlet (462) of the cowling (460). The cowling (460) is configured to circulate the air flow (111) within an interior space (426) of the cowling (460) that extends radially outward from an inner diameter (423) of the cowling to an outer diameter (424) of the cowling at an outer surface.
    Type: Grant
    Filed: February 29, 2012
    Date of Patent: December 8, 2015
    Assignee: Siemens Energy, Inc.
    Inventors: David J. Wiebe, Jose L. Rodriguez
  • Patent number: 9188016
    Abstract: An impingement plate for a turbine vane with an integrated cooling flow metering device is disclosed. The impingement plate—which covers the outer end of the vane and allows some cooling air to pass through to the vane's top surface—is re-designed to incorporate an orifice plate for metering the amount of cooling air flow which enters a cooling passage in the vane. The multi-hole orifice pattern in the metering device is designed to optimize the downstream airflow pattern, thus improving heat transfer from the vane to the cooling air. The reduced cooling air flow through the vane results in increased turbine engine efficiency, and the re-designed impingement plate can be used with the existing vane design.
    Type: Grant
    Filed: December 10, 2013
    Date of Patent: November 17, 2015
    Assignee: Siemens Energy, Inc.
    Inventors: Gilles Carrier, Kerri Santucci
  • Patent number: 9181816
    Abstract: A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a stationary vane assembly and a rotating blade assembly axially upstream from the vane assembly. A platform of the blade assembly has a radially outwardly facing first surface, an axially downstream facing second surface defining an aft plane, and a plurality of grooves extending into the second surface such that the grooves are recessed from the aft plane The grooves are arranged such that a circumferential space is defined between adjacent grooves During operation of the engine, the grooves impart a circumferential velocity component to purge air flowing out of a disc cavity through the grooves to guide the purge air toward a hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.
    Type: Grant
    Filed: February 25, 2014
    Date of Patent: November 10, 2015
    Assignee: SIEMENS AKTIENGESELLSCHAFT
    Inventors: Ching-Pang Lee, Kok-Mun Tham, Eric Schroeder, Erik Johnson, Dustin Muller, Steven Coppess, Manjit Shivanand, Kahwai G. Muriithi
  • Patent number: 9115587
    Abstract: Cooling air is provided from a source of cooling air through a cooling air circuit in a turbine section of a gas turbine engine. A first portion of cooling air is provided from the source along a first path of the circuit to a plurality of blades associated with a stage of the turbine section. A second portion of cooling air is provided from the source along a second path of the circuit. The second path includes a turbine disc bore where the cooling air provides cooling to a radially innermost portion of at least one turbine disc that forms a part of a rotor of the engine. The second path is independent from the first path such that the second portion of cooling air bypasses the stage and is not mixed with the first portion of cooling air in the circuit after leaving the source.
    Type: Grant
    Filed: August 22, 2012
    Date of Patent: August 25, 2015
    Assignee: Siemens Energy, Inc.
    Inventors: Jiping Zhang, Yan Yin
  • Patent number: 9080449
    Abstract: A seal assembly for a gas turbine engine includes an annular body and a flow-through tube that extends through the annular body. The flow-through tube includes an upstream orifice, a downstream orifice and a tube body that extends between the upstream orifice and the downstream orifice. The tube body establishes a gradually increasing cross-sectional area between the downstream orifice and the upstream orifice.
    Type: Grant
    Filed: August 16, 2011
    Date of Patent: July 14, 2015
    Assignee: United Technologies Corporation
    Inventors: Joseph W. Bridges, David F. Cloud, David P. Houston, Eric W. Malmborg
  • Patent number: 9068513
    Abstract: A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a rotating blade assembly having a plurality of blades that rotate with a turbine rotor during operation of the engine, and a stationary vane assembly having a plurality of vanes and an inner shroud. The inner shroud includes a radially outwardly facing first surface, a radially inwardly facing second surface, and a plurality of grooves extending into the second surface. The grooves are arranged such that a space having a component in a circumferential direction is defined between adjacent grooves. During operation of the engine, the grooves guide purge air out of the disc cavity toward the hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.
    Type: Grant
    Filed: January 23, 2013
    Date of Patent: June 30, 2015
    Assignee: Siemens Aktiengesellschaft
    Inventor: Ching-Pang Lee
  • Patent number: 9062557
    Abstract: A gas turbine having rotor discs (9), a disc cavity (13) and a stator stage (25) extending to the disc cavity (13). Seal housing flanges (43, 44) extend from a seal housing (29) of the stator stage (25). Rotor flanges (41i, 41o) extend from a rotor disk (9-1). An inner rotor flange (41i) and first seal housing flange (43) are inward from a second seal housing flange (44). One rotor flange (41o) is outward from the second seal housing flange (44). The inner rotor flange (41i) and first seal housing flange (43) extend toward one another to limit movement of main gas flow (17). An inlet (47) injects air (50) between the outward rotor flange (41o) and second seal housing flange (44).
    Type: Grant
    Filed: September 7, 2011
    Date of Patent: June 23, 2015
    Assignee: Siemens Aktiengesellschaft
    Inventors: Kok-Mun Tham, Ching-Pang Lee, Abdullatif M. Chehab, Gm Salam Azad, Shantanu P. Mhetras, Manjit Shivanand, Vincent P. Laurello, Christopher Rawlings
  • Patent number: 9038398
    Abstract: A gas turbine engine includes a heat exchanger, a bearing compartment, and a nozzle assembly in fluid communication with the bearing compartment. The heat exchanger exchanges heat with a bleed airflow to provide a conditioned airflow. The bearing compartment is in fluid communication with the heat exchanger. A first passageway communicates the conditioned airflow from the heat exchanger to the bearing compartment. A second passageway communicates the conditioned airflow from the bearing compartment to the nozzle assembly.
    Type: Grant
    Filed: February 27, 2012
    Date of Patent: May 26, 2015
    Assignee: United Technologies Corporation
    Inventors: Gabriel L. Suciu, Ioannis Alvanos
  • Patent number: 9038396
    Abstract: Disclosed is an apparatus for cooling a transition piece of a combustor includes at least one wrapper disposed at the transition piece located outboard of the transition piece. At least one support boss is located between the at least one wrapper and the transition piece. The at least one support boss, the at least one wrapper, and transition piece define at least one cooling flow channel for directing flow for cooling the transition piece. A method of cooling a transition piece of a combustor includes flowing cooling flow into at least one cooling flow channel located at the transition piece, the at least one cooling flow channel defined by the transition piece, at least one wrapper located at the transition piece and located outboard of the transition piece, and at least one support boss located between the at least one wrapper and the transition piece.
    Type: Grant
    Filed: April 8, 2008
    Date of Patent: May 26, 2015
    Assignee: GENERAL ELECTRIC COMPANY
    Inventors: Ronald James Chila, David Purnell
  • Publication number: 20150135715
    Abstract: The invention provides a gas turbine cooling system that efficiently cools a first-stage turbine wheel and the attachment portions of first-stage turbine blades. A gas turbine comprises: a compressor 1 for compressing air; a combustor 2 for combusting the air compressed by the compressor 1 with fuel; and a turbine 3 driven by the combustion gas generated by the combustor 2, the turbine 3 including at least one turbine wheel 12a having turbine blades 11a on an outer circumferential section thereof. A gas turbine cooling system comprises a group of impingement cooling holes 32, provided in a partition wall 16a that separates the exit space 42 of the compressor 1 and the wheel space 43a located upstream of the turbine wheel 12a, for ejecting the compressed air 101 in the exit space 42 toward the turbine wheel 12a and the attachment portions of the turbine blades 11a.
    Type: Application
    Filed: November 18, 2014
    Publication date: May 21, 2015
    Inventors: Tomoyuki MATSUI, Hayato MAEKAWA, Ryo KAWAI
  • Publication number: 20150121898
    Abstract: A turbine and so on capable of enabling high reliability are provided. In the turbine of an embodiment, a turbine rotor is accommodated in a turbine casing, and is rotated by a working medium which is introduced after flowing in an inlet pipe of a combustor. A sleeve is provided at the turbine casing, and accommodates the inlet pipe therein. Here, the sleeve is thicker than the inlet pipe, and a cooling fluid whose temperature is lower than the working fluid flows between the inlet pipe and the sleeve.
    Type: Application
    Filed: January 13, 2015
    Publication date: May 7, 2015
    Applicant: KABUSHIKI KAISHA TOSHIBA
    Inventors: Tsuguhisa TASHIMA, Shogo IWAI, Masao ITO, Shunsuke TAKAE
  • Patent number: 9021816
    Abstract: A turbine vane for a gas turbine engine includes inner and outer platforms joined by a radially extending airfoil. The airfoil includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. The airfoil includes an airfoil cooling passage. A platform cooling passage is arranged within at least one of the inner and outer platforms. The platform cooling passage includes multiple cooling regions with one of the cooling regions arranged beneath the airfoil cooling passage.
    Type: Grant
    Filed: July 2, 2012
    Date of Patent: May 5, 2015
    Assignee: United Technologies Corporation
    Inventors: Russell J. Bergman, Leonard A. Bach, Brandon W. Spangler
  • Publication number: 20150107267
    Abstract: Effusion cooling holes formed through a transition component provided in a combustion section of a gas turbine engine. The transition component directs a hot working gas from a combustion basket to a first row of vanes in a turbine section of the engine. The effusion cooling holes are formed through an outer wall of the transition component in a direction so that the flow of air through the effusion holes is in a direction substantially opposite to the bulk flow direction of the working gas through the transition component.
    Type: Application
    Filed: October 21, 2013
    Publication date: April 23, 2015
    Inventors: Blake R. Cotten, Richard L. Thackway, Charalambos Polyzopoulos
  • Patent number: 9010127
    Abstract: A transition piece aft frame assembly is provided, and includes a transition piece aft frame and a heat shield. The transition piece aft frame has an aft face. At least a portion of the aft face is exposed to an exhaust gas stream. The heat shield is connected to the transition piece aft frame. The heat shield is oriented to generally deflect the exhaust gas stream away from the aft face of the transition piece aft frame.
    Type: Grant
    Filed: March 2, 2012
    Date of Patent: April 21, 2015
    Assignee: General Electric Company
    Inventors: Christopher Paul Willis, William Lawrence Byrne, David William Cihlar, Donald Timothy Lemon, Patrick Benedict Melton
  • Patent number: 9003807
    Abstract: The present invention comprises a gas turbine engine and a process for operating a gas turbine engine. A fluid structure receives compressed air from a compressor and extends toward a stationary blade ring in a turbine to discharge the compressed air directly against a surface of the blade ring such that the compressed air impinges on the blade ring surface. The compressed air then passes through at least one opening in the stationary blade ring and into cooling passages of a corresponding row of vanes.
    Type: Grant
    Filed: November 8, 2011
    Date of Patent: April 14, 2015
    Assignee: Siemens Aktiengesellschaft
    Inventors: Abdullatif M. Chehab, David A. Little
  • Publication number: 20150096306
    Abstract: A turbine nozzle vane segment includes one or more nozzle vanes extending between radially inner and outer side walls, each nozzle vane having a peripheral edge wall extending between a leading edge and a trailing edge of the vane. In one exemplary embodiment, at least one substantially radially-oriented cooling channel is formed in the peripheral edge wall at the leading edge, with openings at opposite ends of the cooling channel. The location and length of the cooling channels may vary about the peripheral edge wall, and the inner cavity of the vane may be provided with ribs extending along and adjacent the one or more cooling channels to reinforce the wall and to also provide additional cooling surface areas in the inner cavity.
    Type: Application
    Filed: October 8, 2013
    Publication date: April 9, 2015
    Applicant: General Electric Company
    Inventors: Sandeep Munshi SARANGAPANI, Ajay Gangadhar PATIL, Poorna Chandra RAO
  • Patent number: 8997500
    Abstract: A turbine engine includes a shaft, a fan, at least one bearing mounted on the shaft and rotationally supporting the fan, a fan drive gear system coupled to drive the fan, a bearing compartment around the at least one bearing and a source of pressurized air in communication with a region outside of the bearing compartment.
    Type: Grant
    Filed: December 30, 2011
    Date of Patent: April 7, 2015
    Assignee: United Technologies Corporation
    Inventors: Jorn A. Glahn, Frederick M. Schwarz
  • Publication number: 20150082808
    Abstract: An airfoil for a gas turbine engine is provided. The airfoil includes a body with a leading edge, a trailing edge, a first side wall extending between the leading edge and the trailing edge, and a second side wall extending between the leading edge and the trailing edge. The body defines an interior cavity. The airfoil includes an interior wall disposed within the interior cavity of the body and extending between the first wall and the second wall to define a supply chamber and a leading edge chamber. The interior wall defines a cooling hole with a base portion and a locally extended portion to direct cooling air from the supply chamber to the leading edge chamber such that the cooling air impinges upon the leading edge.
    Type: Application
    Filed: April 2, 2013
    Publication date: March 26, 2015
    Applicant: HONEYWELL INTERNATIONAL INC.
    Inventor: HONEYWELL INTERNATIONAL INC.
  • Patent number: 8984858
    Abstract: One embodiment of the present invention is a unique gas turbine engine. Another embodiment of the present invention is a unique gas turbine engine bearing system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine engine bearing systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.
    Type: Grant
    Filed: December 23, 2011
    Date of Patent: March 24, 2015
    Assignee: Rolls-Royce Corporation
    Inventor: Matthew Michael Miller
  • Patent number: 8984895
    Abstract: A method and apparatus are disclosed for a gas turbine spool design combining metallic and ceramic components in a way that controls clearances between critical components over a range of engine operating temperatures and pressures. In a first embodiment, a ceramic turbine rotor rotates just inside a ceramic shroud and separated by a small clearance gap. The ceramic rotor is connected to a metallic volute. In order to accommodate the differential rates of thermal expansion between the ceramic rotor and metallic volute, an active clearance control system is used to maintain the desired axial clearance between ceramic rotor and the ceramic shroud over the range of engine operating temperatures. In a second embodiment, a ceramic turbine rotor rotates just inside a ceramic shroud which is part of a single piece ceramic volute/shroud assembly.
    Type: Grant
    Filed: July 11, 2011
    Date of Patent: March 24, 2015
    Assignee: ICR Turbine Engine Corporation
    Inventors: James B. Kesseli, Matthew Stephen Baldwin
  • Publication number: 20150075180
    Abstract: A turbine bucket is disclosed herein. The turbine bucket may include a platform and a shank portion extending radially inward from the platform. The shank portion may include a slash face, a radial seal pin groove formed in the slash face, and at least one cooling hole disposed in the slash face about the radial seal pin groove.
    Type: Application
    Filed: September 18, 2013
    Publication date: March 19, 2015
    Applicant: General Electric Company
    Inventors: Xiuzhang James Zhang, James W. Vehr
  • Patent number: 8978390
    Abstract: A wall of a component of a gas turbine engine includes first and second wall surfaces, an inlet located at the first wall surface, an outlet located at the second surface, a metering section commencing at the inlet and extending downstream from the inlet, and a diffusing section extending from the metering section and terminating at the outlet. The diffusing section includes a leading edge formed at an upstream end of the outlet, a trailing edge formed at a downstream end of the outlet, a body region upstream of the trailing edge, and a plurality of crenellation features located on the body region.
    Type: Grant
    Filed: March 11, 2014
    Date of Patent: March 17, 2015
    Assignee: United Technologies Corporation
    Inventors: Glenn Levasseur, Edward F. Pietraszkiewicz
  • Patent number: 8978387
    Abstract: The flow through the core of a hybrid pulse detonation combustion system is passed through a compressor and then separated into a primary flow, that passes directly to the combustor, and a bypass flow, which is routed to a portion of the system to be used to cool components of the system. The bypass flow is routed to a nozzle of the pulse detonation combustor. The flow is then passed back into the primary flow through the core downstream of where it was extracted.
    Type: Grant
    Filed: December 21, 2010
    Date of Patent: March 17, 2015
    Assignee: General Electric Company
    Inventors: Venkat Eswarlu Tangirala, Narendra Digamber Joshi, Adam Rasheed, Brian Gene Brzek, Douglas Carl Hofer, Thomas Michael Lavertu, Fuhua Ma
  • Patent number: 8973371
    Abstract: A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section cooling system. The turbine section cooling system including a first compressed air bleed arrangement and a second compressed air bleed arrangement. The first compressed air bleed arrangement bleeds a first flow of compressed air from a high pressure stage of the compressor section. The first flow of compressed air bypasses the combustor and arrives at the turbine section to form a sealing and/or cooling flow at a row of stator vanes upstream of an adjacent rotor disc. The second compressed air bleed arrangement bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section.
    Type: Grant
    Filed: September 2, 2011
    Date of Patent: March 10, 2015
    Assignee: Rolls-Royce PLC
    Inventors: Jonathan M King, Crispin D. Bolgar, Guy D. Snowsill, Michael J. Sheath, Geoffrey M Dailey
  • Patent number: 8973373
    Abstract: A method to provide clearance control for a gas turbine having a multi-stage compressor and a turbine having turbine buckets rotating within a turbine shell, the method includes: selecting a first compressor stage from which to extract compressed air; ducting the compressed air from the first compressor stage to the turbine shell; passing the compressed air from the first compressor stage to thermally contract the turbine shell; selecting a second compressor stage from which to extract compressed air and deselecting the first compressor stage; ducting the compressed air from the second compressor stage to the turbine shell, and passing the compressed air from the second compressor stage to thermal expand the turbine shell.
    Type: Grant
    Filed: October 31, 2011
    Date of Patent: March 10, 2015
    Assignee: General Electric Company
    Inventor: Malath Ibrahim Arar
  • Publication number: 20150059357
    Abstract: A system for providing cooling for a turbine component that includes an outer surface exposed to combustion gases is provided. A component base includes at least one fluid supply passage coupleable to a source of cooling fluid. At least one feed passage communicates with the at least one fluid supply passage. At least one delivery channel communicates with the at least one feed passage. At least one cover layer covers the at least one feed passage and the at least one delivery channel, defining at least in part the component outer surface. At least one discharge passage extends to the outer surface. A diffuser section is defined in at least one of the at least one delivery channel and the at least one discharge passage, such that a fluid channeled through the system is diffused prior to discharge adjacent the outer surface.
    Type: Application
    Filed: September 3, 2013
    Publication date: March 5, 2015
    Applicant: General Electric Company
    Inventors: Victor John Morgan, Benjamin Paul Lacy
  • Publication number: 20150040582
    Abstract: A cooling circuit for a turbine bucket having an airfoil portion includes a trailing edge cooling circuit portion provided with a first radially outwardly directed inlet passage intermediate leading and trailing edges of the airfoil portion of the bucket, extending from a platform portion of the bucket to a location adjacent a radially outer tip of the bucket, and connecting to a second radially inwardly directed passage extending from a location adjacent the radially outer tip to a location adjacent the platform portion. The second radially inwardly directed passage connects to a third trailing edge region passage, and a plurality of crossover passages connect a radially outer half of the second radially inwardly directed passage to a radially outer half of the third trailing edge region passage.
    Type: Application
    Filed: August 7, 2013
    Publication date: February 12, 2015
    Applicant: General Electric Company
    Inventors: Zhirui Dong, Xiuhang James Zhang, Melbourne James Myers, Camilo Andres Sampayo
  • Patent number: 8950192
    Abstract: A gas turbine includes a turbine section; an annular combustor disposed upstream of the turbine section and configured to discharge a hot gas flow on an outlet side to the turbine section; an outer shell delimiting the combustor and splittable at a parting plane; a plenum enclosing the outer shell; a rotor; a turbine vane carrier encompassing the rotor; a plurality of stator vanes disposed on the vane carrier, and at least two sealing segments forming a ring, each of the at least two sealing segments having an inner edge and a head and a foot section and being movably mounted on the inner edge by the foot section to the outer shell and by the head section to the turbine vane carrier so as to mechanically connect the combustor to the turbine vane carrier.
    Type: Grant
    Filed: August 16, 2010
    Date of Patent: February 10, 2015
    Assignee: Alstom Technology Ltd.
    Inventors: Remigi Tschuor, Russell Bond Jones, Luis J. Rodriguez, Hartmut Haehnle, Marion Duggan (-Oneil)
  • Publication number: 20150033761
    Abstract: A heat transfer assembly for controlling heat transfer of a turbine engine is provided. The turbine engine includes a housing and includes a compressor, a combustor and a turbine located within the housing. The heat transfer assembly includes a flow control device having a sidewall coupled to the turbine, the sidewall is in flow communication with a compressor vane. The sidewall is configured to define a first flow path from the compressor vane to a turbine vane and a second flow path from the compressor vane to a turbine blade. A heat exchanger is coupled to the housing and located between the compressor and the turbine, wherein the heat exchanger is in flow communication with at least one of the first flow path and the second flow path. A fluid supply device is coupled to the housing and in flow communication with the heat exchanger.
    Type: Application
    Filed: July 31, 2013
    Publication date: February 5, 2015
    Applicant: General Electric Company
    Inventors: Corey Bourassa, William Dwight Gerstler
  • Patent number: 8943827
    Abstract: A gas turbine engine with a fuel air heat exchanger located in the high pressure plenum. The heat exchanger includes at least one air conduit and at least one fuel conduit in heat exchange relationship with one another, with a fuel flow communication between a fuel source and fuel distribution members of the combustor being provided at least partly through the at least one fuel conduit, and the at least one air conduit defining a fluid flow communication between the high pressure plenum and an engine component to be cooled by the compressed air.
    Type: Grant
    Filed: May 31, 2011
    Date of Patent: February 3, 2015
    Assignee: Pratt & Whitney Canada Corp.
    Inventors: Lev Alexander Prociw, Eduardo Hawie