| Class / Patent application number | Description | Number of patent applications / Date published |
|---|
| 060755000 | Having means to direct flow along inner surface of liner | 64 |
| 20080256956 | Methods and systems to facilitate reducing combustor pressure drops - A method for assembling a gas turbine combustor includes providing a combustor case having a first end, a second end, and a centerline extending there between, coupling an end cover to the case first end, coupling a combustor liner within the case such that the liner is substantially coaxially aligned with respect to the case, and providing a streamline flow conditioner including a body that includes a radially outer surface and a radially inner surface, a deflection plate that extends from the body, and coupling the streamline flow conditioner to the case second end such that the streamline flow conditioner is coupled radially between the case and the combustor liner such that the deflection plate is adjacent the case second end. The deflection plate inner surface at least one of extends radially outward with respect to the centerline and defines a plurality of openings within the plate inner surface. | 10-23-2008 |
| 20080271458 | Zero-Cross-Flow Impingement Via An Array of Differing Length, Extended Ports - A jet impingement array design and method are disclosed for efficiently cooling the liner of a gas turbine combustion chamber, while eliminating almost all effects of coolant gas crossflow. The design includes an array of extended jet ports for which the distance between the ends of the jet ports and the surface to be cooled progressively decreases from upstream to downstream. Spent air from upstream jets is directed away from downstream jets, thereby reducing the detrimental effects of crossflow, and optimizing heat transfer. | 11-06-2008 |
| 20090120096 | Gas Turbine Engine Systems Involving Cooling of Combustion Section Liners - Gas turbine engine systems involving cooling of combustion section liners are provided. In this regard, a representative liner includes: an outer side, an inner side, an upstream end and a downstream end, the outer side being configured to face away from a combustion reaction, the inner side being configured to face the combustion reaction; a cooling air channel, at least a portion of the cooling air channel being located in a vicinity of the downstream end; and cooling holes formed through the inner side of the liner, the cooling holes being in fluid communication with the cooling air channel such that cooling air provided to the cooling air channel is directed through the cooling holes and to the inner side of the liner such that at least a portion of the inner side of the liner receives cooling air despite a corresponding portion located on the outer side of the liner being obstructed from directly receiving cooling air. | 05-14-2009 |
| 20090145132 | METHODS AND SYSTEM FOR REDUCING PRESSURE LOSSES IN GAS TURBINE ENGINES - A method of assembling a combustor assembly is provided, wherein the method includes providing a combustor liner having a centerline axis and defining a combustion chamber therein, and coupling an annular flowsleeve radially outward from the combustor liner such that an annular flow path is defined substantially circumferentially between the flowsleeve and the combustor finer. The method also includes orienting the flowsleeve such that a plurality of inlets formed within the flowsleeve are positioned to inject cooling air in a substantially axial direction into the annular flow path to facilitate cooling the combustor finer. | 06-11-2009 |
| 20090249791 | TRANSITION PIECE IMPINGEMENT SLEEVE AND METHOD OF ASSEMBLY - Disclosed herein is an impingement sleeve for a gas turbine combustion transition piece. The impingement sleeve includes, a housing positionable proximate a first portion of a transition piece body having at least two flanges integrally formed thereon, and a cover positionable proximate a second portion of the transition piece body having at least two lips integrally formed thereon, the at least two lips are weldlessly attachable to the at least two flanges to position the impingement sleeve radially outwardly of the transition piece body. | 10-08-2009 |
| 20100170258 | COOLING APPARATUS FOR COMBUSTOR TRANSITION PIECE - Disclosed is an apparatus for cooling a single-piece duct includes at least one metal wrapper disposed at the transition piece located outboard of the transition piece. At least one support boss is located between the metal wrapper and the transition piece. The support boss, the metal wrapper and transition piece define at least one cooling flow channel for directing flow for cooling the transition piece. A related method of cooling a single-piece duct extending between a forward end of a combustor and a first stage of a turbine includes providing a plurality of flow direction devices radially between the single-piece duct and a metal wrapper extending at least partially around the single-piece duct; flowing cooling air into a space between the single-piece duct and the metal wrapper such that the plurality of flow direction devices guide the cooling air about a surface of the single-piece duct enclosed by the metal wrapper. | 07-08-2010 |
| 20100170259 | METHOD AND APPARATUS TO ENHANCE TRANSITION DUCT COOLING IN A GAS TURBINE ENGINE - A method and apparatus are described that include a transition piece including a cooling sleeve. The cooling sleeve includes a first end and an opposite second end, the cooling sleeve is coupled to an inner wall of the transition piece, such that an annular passage is defined between the inner wall and the cooling sleeve. The first end defines an annular inlet and second end defines an annular outlet. | 07-08-2010 |
| 20100170260 | GAS TURBINE COMBUSTOR - A gas turbine combustor includes a fuel supplying section and a combustion tube. The fuel supplying section supplies fuel to a combustion zone inside the combustion tube. The combustion tube passes combustion gas to the turbine. The combustion tube is provided with a first region where an air passage for cooling air is formed and a second region where a steam passage for cooling steam is formed. The second region is located downstream of the first region in a direction of a mainstream flow of the combustion gas. | 07-08-2010 |
| 20100180601 | COOLING STRUCTURE OF GAS TURBINE COMBUSTOR - It is required for a gas turbine combustor to exhaust low NOx. The gas turbine combustor is provided with a combustion tube which has a cooling passage through which cooling air flows in a double wall structure. The cooling passage has a main cooling air supply opening opened to a side of a combustion zone. The cooling air supplied from the main cooling air supply opening is guided to a direction along an inner wall surface of the combustion tube by a guide. The cooling air flows through the cooling passage inside the combustion tube, and then is reused for film cooling along the inner wall surface. Thus, it is possible to save cooling air. Therefore, a more part of the air supplied from a compressor can be used as air for combustion and it becomes possible to exhaust low NOx. | 07-22-2010 |
| 20100186415 | TURBULATED AFT-END LINER ASSEMBLY AND RELATED COOLING METHOD - In a combustor for a turbine a cover sleeve is disposed between the aft end portion of the combustor liner and a resilient seal structure to define an air flow passage therebetween. The cover sleeve has at a forward end thereof a plurality of air inlet feed holes for directing cooling air into the air flow passage. A radially outer surface of the combustor liner aft end portion defining the air flow passage includes a plurality of turbulators projecting towards but spaced from the cover sleeve and a plurality of supports extending to and engaging the cover sleeve to space the cover sleeve from the turbulators to define the air flow passage. | 07-29-2010 |
| 20100186416 | FLOW CONDITIONER FOR USE IN GAS TURBINE COMPONENT IN WHICH COMBUSTION OCCURS - A gas turbine component in which combustion occurs. The gas turbine component includes a liner, including a first surface facing a first space and a second surface facing a second space, the liner being interposed between the first and second spaces and having a through-hole defined therein extending from the first to the second surface by which incoming flows proceed from the first space and to the second space. At least the first surface is formed to flow condition the incoming flows to resist separating from sidewalls of the through-hole. | 07-29-2010 |
| 20100251722 | Wall elements for gas turbine engine combustors - A combustor wall element arrangement for a gas turbine engine comprises an upstream wall element ( | 10-07-2010 |
| 20100257863 | COMBINED CONVECTION/EFFUSION COOLED ONE-PIECE CAN COMBUSTOR - An industrial turbine engine comprises a combustion section, an air discharge section downstream of the combustion section, a transition region between the combustion and air discharge section, a combustion transition piece and a sleeve. The transition piece defines an interior space for combusted gas flow. The sleeve surrounds the combustor transition piece so as to form a flow annulus between the sleeve and the transition piece. The sleeve includes a first set of apertures for directing cooling air from compressor discharge air into the flow annulus. The transition piece includes an outer surface bounding the flow annulus and an inner surface bounding the interior surface, and includes a second set of apertures for directing cooling air in the flow annulus to the interior space. Each of the second set of apertures extends from an entry portion on the outer surface to an exit portion on the inner surface. | 10-14-2010 |
| 20100263384 | COMBUSTOR CAP WITH SHAPED EFFUSION COOLING HOLES - A combustor cap assembly for a gas turbine includes a plurality of effusion cooling apertures that allow air to pass through the cooling apertures to cool the combustor cap assembly. An inner diameter of the cooling apertures expands along at least a portion of the total length of the apertures so that cooling air passing through the cooling aperture will slow as it approaches the outlet. | 10-21-2010 |
| 20110011093 | GAS-TURBINE COMBUSTION CHAMBER WITH STARTER FILM FOR COOLING THE COMBUSTION CHAMBER WALL - A gas turbine combustion chamber has a starter film for cooling the combustion chamber wall, and a combustion chamber head, into which cooling air can be introduced and which is confined to the combustion chamber by a heat shield ( | 01-20-2011 |
| 20110088402 | ANNULAR COMBUSTION CHAMBER FOR A GAS TURBINE ENGINE - An annular combustion chamber for a gas turbine engine includes radially inner and outer walls connected together by a chamber end wall including openings, each of which receives a fuel injection system. Heat protection deflectors are fastened to the chamber end wall. Holes are formed through the chamber end wall to pass cooling air onto an upstream face of each deflector. The inner or outer edge of a deflector presents a sealing rim engaging the respective inner or outer wall of the chamber. | 04-21-2011 |
| 20110107766 | COMBUSTOR ASSEMBLY FOR A TURBINE ENGINE WITH ENHANCED COOLING - A combustor assembly for a turbine engine includes a combustor liner and a flow sleeve which surrounds the combustor liner. Compressed air flows through an annular space located between an outer surface of the combustor liner and an inner surface of the flow sleeve. A plurality of cooling holes are formed through the flow sleeve to allow compressed air to flow from a position outside the flow sleeve, through the cooling holes, and into the annular space. The height of the annular space may vary along the length of the combustor assembly. Thus, the flow sleeve may have reduced diameter portions which result in the height of the annular space being smaller in certain locations than at other locations along the length of the combustor assembly. | 05-12-2011 |
| 20110154826 | GAS TURBINE, METHOD OF CONTROLLING AIR SUPPLY AND COMPUTER PROGRAM PRODUCT FOR CONTROLLING AIR SUPPLY - The present invention provides a gas turbine capable of reducing energy consumption while suppressing a so-called cat back phenomenon. The gas turbine includes a combustor-accommodating chamber casing for accommodating therein a combustor which burns fuel and air compressed by a compressor to generate combustion gas and which injects the combustion gas to a turbine. The gas turbine also includes a first air supply passage and a second air supply passage on an upper portion of the combustor-accommodating chamber casing in the vertical direction. The first air supply passage discharges air toward the compressor in the combustor-accommodating chamber casing. The second air supply passage discharges air in a direction different from that of the first air supply passage. | 06-30-2011 |
| 20110185739 | GAS TURBINE COMBUSTORS WITH DUAL WALLED LINERS - A combustor for a turbine engine includes a hot wall and a cold wall forming a dual walled liner and a liner cavity with the hot wall. The cold wall defines a plurality of impingement cooling holes configured to deliver an impingement cooling flow. A first downstream end terminates the hot wall and is configured to receive the impingement cooling flow from the plurality of impingement cooling holes, and a second downstream end terminates the cold wall and is longer in a generally downstream direction than the first downstream end. A combustion chamber is formed with the dual walled liner and the liner and faces an opposite side of the hot wall relative to the combustion chamber. The combustion chamber has a longitudinal axis and is configured to receive an air-fuel mixture in the generally downstream direction along the longitudinal axis. | 08-04-2011 |
| 20110185740 | COMBUSTOR LINER SEGMENT SEAL MEMBER - A combustor for a gas turbine engine is provided that includes a support shell, a forward liner segment, an aft liner segment, and a seal member. The support shell has an interior surface, and exterior surface, and a plurality of impingement apertures disposed within the shell. The forward liner segment and aft liner segment are attached to the inner surface of the shell. The forward liner segment has an edge surface extending between a face surface and a back surface, and a seal shoulder. The aft liner segment has an edge surface extending between a face surface and a back surface, and a seal shoulder. The forward liner segment and the aft liner segment are separated from one another by a gap. The seal member is disposed within the gap. At least some of the plurality of impingement apertures disposed within the shell are aligned with the seal member and oriented to direct cooling air to impinge on the seal member. | 08-04-2011 |
| 20110214429 | ANGLED VANES IN COMBUSTOR FLOW SLEEVE - A turbine combustor liner assembly includes a combustor liner having upstream and downstream ends; a transition duct attached to the downstream end of the combustor liner; a first flow sleeve surrounding the combustor liner, with a first radial flow passage therebetween; and a first annular inlet at an aft end of the flow sleeve, the inlet provided with a plurality of circumferentially spaced, angled flow vanes arranged to swirl air entering the first radial flow passage via the annular inlet. | 09-08-2011 |
| 20110265484 | COMBUSTION DEVICE FOR A GAS TURBINE - A combustion device ( | 11-03-2011 |
| 20120055165 | COMBUSTOR LINER ASSEMBLY WITH ENHANCED COOLING SYSTEM - An enhanced cooling system for the downstream end of a combustor liner assembly is provided. The downstream end region of the combustor liner assembly is formed by a subassembly comprising an outer tubular member, an inner tubular member and a middle band. The middle band includes a plurality of undulations having alternating peaks and valleys. The middle band is operatively positioned between the inner and outer tubular members to form a plurality of elongated cooling passages. A turbulence generator is provided along each of the cooling passages and extends radially inward from a portion of the peaks such that the turbulence generator and the valleys contact an outer peripheral surface of the inner tubular member. The surface area of the middle band that is in contact with the inner tubular member is greater than the surface area of the middle band that is in contact with the outer tubular member. | 03-08-2012 |
| 20120102960 | GAS TURBINE ENGINE SYSTEMS INVOLVING COOLING OF COMBUSTION SECTION LINERS - Gas turbine engine systems involving cooling of combustion section liners are provided. A representative liner includes: an outer side, an inner side, an upstream end, and a downstream end, the outer side being configured to face away from a combustion reaction, the inner side being configured to face the combustion reaction; a cooling air channel, a portion of the cooling air channel being located proximate the downstream end; and cooling holes formed through the inner side of the liner, the cooling holes being in fluid communication with the cooling air channel such that cooling air provided to the cooling air channel is directed through the cooling holes and to the inner side of the liner such that at least a portion of the inner side of the liner receives cooling air despite a corresponding portion located on the outer side of the liner being obstructed from receiving cooling air. | 05-03-2012 |
| 20120117973 | GAS-TURBINE COMBUSTION CHAMBER WITH A COOLING-AIR SUPPLY DEVICE - A gas-turbine combustion chamber has a cooling-air supply device delivering cooling air to a combustion chamber wall | 05-17-2012 |
| 20120125006 | GAS TURBINE COMBUSTOR AND GAS TURBINE - In a gas turbine combustor and a gas turbine, an air passage ( | 05-24-2012 |
| 20120144835 | COMBUSTION CHAMBER - A combustion chamber has an outer wall supporting a number of wall elements or tiles. The tiles are located on the wall by bosses so that a cooling space is provided between the wall and the tiles. Air holes are provided through the wall and the tiles and a flow passage is provided adjacent the air holes. A flow passage is defined by changing the profiles of the air holes in the outer wall and/or the location features to provide a localised gap through which cooling air is directed to cool regions subject to overheating and extend service life. | 06-14-2012 |
| 20120198854 | RESONATOR SYSTEM WITH ENHANCED COMBUSTOR LINER COOLING - A combustor liner has a plurality of resonators formed thereon. Each resonator has a radially outer wall and at least one side wall. The outer wall can be free of holes. Each resonator has an inner cavity defined between the outer wall, the at least one side wall and the outer peripheral surface of the liner. The at least one side wall of each resonator surrounds a subset of a plurality of holes that extend substantially radially through the liner. A plurality of cooling passages extends generally longitudinally within the liner. Each cooling passage has an inlet in fluid communication with the exterior of the liner and an outlet in fluid communication with the inner cavity of a respective one of the resonators. A coolant, such as compressor air, can enter and flow along the cooling passages to thereby cool the liner and purge the inner cavity of the resonator. | 08-09-2012 |
| 20120297778 | COMBUSTORS WITH QUENCH INSERTS - A combustor is provided for a turbine engine. The combustor includes a first liner having a first hot side and a first cold side; a second liner having a second hot side and a second cold side, the second hot side and the first hot side forming a combustion chamber therebetween. The combustion chamber is configured to receive an air-fuel mixture for combustion therein. The combustor further includes an insert having a body portion extending through the first liner and terminating at a tip, the body portion configured to direct air flow into the combustion chamber. The insert further includes a cooling hole defined in the body portion and configured to direct a first portion of the air flow toward the tip as cooling air. | 11-29-2012 |
| 20130000310 | FLOW SLEEVE IMPINGEMENT COOLING BAFFLES - A combustor assembly for a turbine engine includes a combustor liner, a flow sleeve and a baffle ring. The flow sleeve surrounds the combustor liner. An annulus is formed between the flow sleeve and the combustor liner. A plurality of row of cooling holes are formed in the flow sleeve. The baffle ring radially surrounds the combustor liner and is located in the annulus. | 01-03-2013 |
| 20130086915 | FILM COOLED COMBUSTION LINER ASSEMBLY - A combustion liner assembly is provided and includes a combustion liner and a transition piece. A portion of the transition piece is circumferentially disposed around a portion of the combustion liner. A seal is attached to the transition piece, and the seal is configured to apply a compressive force to an aft end of the combustion liner. | 04-11-2013 |
| 20140096528 | Cooling for Combustor Liners with Accelerating Channels - A combustor liner which reduces cooling flow to a combustion chamber and augments pressure drop split between impingement holes and effusion holes is disclosed. The combustor liner may further include accelerating channels, trip strips, pedestals, and cone-shaped effusion holes to provide further cooling of the liner. The combustor liner may reduce NOx production and the temperature of the combustion chamber of a gas turbine engine or the like. | 04-10-2014 |
| 20140116060 | COMBUSTOR AND A METHOD FOR COOLING THE COMBUSTOR - A combustor includes a first shroud extending circumferentially inside the combustor and at least partially defining an inlet passage. A second shroud extends circumferentially inside the combustor. The second shroud defines an outlet passage. A first plate extends radially inside the second shroud downstream from the inlet passage of the first shroud and upstream from the outlet passage of the second shroud. The first plate generally defines an inlet port and an outlet port. A second plate extends radially around the first plate downstream from the inlet port and upstream from the outlet port of the first plate. A first fluid flow path extends from the inlet passage to the inlet port. A second fluid flow path extends from the outlet port to the outlet passage. A baffle extends from the first shroud to the first plate. The baffle separates the first and second fluid flow paths. | 05-01-2014 |
| 20140190171 | COMBUSTORS WITH HYBRID WALLED LINERS - A combustor for a turbine engine includes a first liner and a second liner forming a combustion chamber with the first liner. The combustion chamber is configured to receive an air-fuel mixture for combustion therein and has a longitudinal axis that defines axial, radial and circumferential directions. The first liner is a first dual walled liner having a first hot wall facing the combustion chamber and a first cold wall that forms a first liner cavity with the first hot wall. The combustor further includes a primary air admission hole defined in the first hot wall and a first fixed liner seal between the first hot wall and the first cold wall proximate to the primary air admission hole. | 07-10-2014 |
| 20140216043 | COMBUSTOR LINER FOR A CAN-ANNULAR GAS TURBINE ENGINE AND A METHOD FOR CONSTRUCTING SUCH A LINER - A combustor liner ( | 08-07-2014 |
| 20140318139 | PREMIXER ASSEMBLY FOR GAS TURBINE COMBUSTOR - A novel and improved combustion liner for use in a gas turbine engine is provided in which the premixer assembly is removably fastened to the combustion liner. Apparatus and method for removably securing the premixer assembly to the combustion liner is provided. A combustor dome plate, swirler assembly and inlet ring basket are coupled together by a plurality of removable fasteners in order to increase accessibility to portions of the combustion liner for inspection and repair processes. | 10-30-2014 |
| 20140318140 | PREMIXER ASSEMBLY AND MECHANISM FOR ALTERING NATURAL FREQUENCY OF A GAS TURBINE COMBUSTOR - A system and method are provided for altering the natural frequency of a dome plate portion of a premixer assembly of a gas turbine combustor. The plate assembly has a dome plate with a central pilot mixer and a plurality of extension tabs extending radially outward from the pilot mixer. A plurality of radially extending struts are secured to the extension tabs in order to alter the natural frequency of the dome plate. | 10-30-2014 |
| 20140331681 | WAKE MANIPULATNIG STRUCTURE FOR A TURBINE SYSTEM - A wake manipulating structure for a turbine system includes a combustor liner defining a combustor chamber. Also included is an airflow path located along an outer surface of the combustor liner. Further included is a wake generating component disposed in the airflow path and proximate the combustor liner, wherein the wake generating component generates a wake region located downstream of the wake generating component. Yet further included is a venturi structure or section disposed in the airflow path configured to reduce the wake region. | 11-13-2014 |
| 20150068212 | COMBUSTOR LINER STOP - A combustion liner includes a substantially cylindrical body having forward and aft ends, the forward end provided with plural, circumferentially-shaped stops adapted for connection to a flow sleeve surrounding the substantially cylindrical body. Each stop includes a radially-projecting strut having an end remote from the substantially-cylindrical body provided with a tab having a shape different than the strut. Each strut is received in a slot formed in the flow sleeve, with the tab on an exterior surface of the flow sleeve. | 03-12-2015 |
| 20150101336 | Cooling Structure for Gas Turbine Combustor Liner - A gas turbine combustor is provided in which product reliability and heat transfer promotion are compatible while suppressing an increase in pressure loss. | 04-16-2015 |
| 20150345298 | GAS TURBINE ENGINE COMPONENT HAVING VASCULAR ENGINEERED LATTICE STRUCTURE - A component according to an exemplary aspect of the present disclosure includes, among other things, a wall and a vascular engineered lattice structure formed inside of the wall. The vascular engineered lattice structure includes at least one of a hollow vascular structure and a solid vascular structure configured to communicate fluid through the vascular engineered lattice structure. | 12-03-2015 |
| 20150375297 | GAS TURBINE ENGINE COMPONENT COOLING PASSAGE AND SPACE EATING CORE - A method of manufacturing a component that includes providing a core structure, casting a component about the core structure, removing a first portion of the core structure from the cast component, and leaving a second portion of the core structure in the cast component to provide a reduced cross-section in the cast component. | 12-31-2015 |
| 20150377032 | GAS TURBINE ENGINE COMPONENT WITH COMBINED MATE FACE AND PLATFORM COOLING - A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge, circumferentially extends between a first mate face and a second mate face, and includes a gas path surface and a non-gas path surface. The component defines at least one cavity that extends at least partially inside of the component. A first plurality of cooling holes extends from the at least one cavity to at least one of the first mate face and the second mate face and a second plurality of cooling holes extends from either the at least one cavity or the non-gas path surface to the gas path surface. | 12-31-2015 |
| 20150377033 | COOLING HOLE FOR A GAS TURBINE ENGINE COMPONENT - A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a wall having an internal surface, an outer skin and a cooling hole having an inlet extending from the internal surface and merging into a metering section, and a diffusion section downstream of the metering section that extends to an outlet located at the outer skin. At least two lobes are embedded within the diffusion section of the cooling hole. At least one surface of each of the at least two lobes is at least partially cylindrical. | 12-31-2015 |
| 20160109130 | PRODUCTION OF TURBINE COMPONENTS WITH HEAT-EXTRACTING FEATURES USING ADDITIVE MANUFACTURING - Turbine engine components having heat-extracting features and methods for manufacturing such components using additive manufacturing are disclosed. An exemplary method comprises providing a base portion of the engine component manufactured by a first manufacturing process and adding one or more heat-extracting features to the engine component using a second manufacturing process different from the first manufacturing process where the second manufacturing process comprises an additive manufacturing process. | 04-21-2016 |
| 20160116167 | D5/D5A DF-42 DOUBLE WALLED EXIT CONE AND SPLASH PLATE - A combustor basket assembly for a gas turbine engine that includes a combustor basket having a basket liner including an input end and an output end. A double-wall exit cone is mounted to the output end of the basket liner, where the exit cone includes an inner wall and an outer wall defining an exit cone channel therebetween. A splash plate is mounted to the outer wall to define a splash plate channel between the splash plate and the basket liner. A series of pairs of cooling feed holes are provided through the basket liner, where one of the feed holes in each pair provides cooling air to the cone channel and the other feed hole provides cooling air to the splash plate channel. The outer surface of the outer wall and the inner surface of the inner wall are coated with a thermal barrier coating. | 04-28-2016 |
| 20160123591 | METHOD FOR DIVERTING FLOW AROUND AN OBSTRUCTION IN AN INTERNAL COOLING CIRCUIT - A rotary machine including: a casing providing an annular chamber for rotating components of the machine; a cooling passage extending through the casing or mounted to a surface of casing; a plug assembly connected to the cooling passage and in the casing or mounted to the casing, wherein the plug assembly includes a collar and a conduit aligned with an axis of the collar, and the collar includes a cooling air by-pass passage in fluid communication with the cooling passage such that cooling air from the cooling passage flows through the by-pass passage and returns to the cooling passage, and another cooling passage or a port extending through the conduit of the plug assembly. | 05-05-2016 |
| 20160123593 | CAN COMBUSTION CHAMBER - The can combustion chamber includes a casing housing a plurality of cans. Each can includes a wall and a perforated cooling liner around the wall. Cooling liners of adjacent cans have staggered perforations. | 05-05-2016 |
| 20160123594 | LOW LUMP MASS COMBUSTOR WALL WITH QUENCH APERTURE(S) - An assembly is provided for a turbine engine. This turbine engine assembly includes a combustor wall including a first layer vertically connected with a second layer. A first portion of the first layer overlaps and is vertically spaced from the second layer by a cavity. A second portion of the first layer is substantially vertically inline with an adjacent portion of the second layer. The second portion of the first layer at least partially forms a quench aperture vertically through the combustor wall. | 05-05-2016 |
| 20160131363 | COMBUSTOR WALL APERTURE BODY WITH COOLING CIRCUIT - An assembly for a turbine engine is provided that includes a combustor wall, which includes an aperture body between a shell and a heat shield. The aperture body at least partially forms a cavity and an aperture that extends through the combustor wall. An inlet passage extends in the combustor wall to the cavity. An outlet passage extends in the combustor wall from the cavity to the aperture. | 05-12-2016 |
| 20160131365 | IMPINGEMENT FILM-COOLED FLOATWALL WITH BACKSIDE FEATURE - Aspects of the disclosure are directed to a liner associated with an engine of an aircraft. The liner includes a panel and an array of projections configured to enhance a cooling of the panel and distributed on at least part of a first side of the panel that corresponds to a cold side of the panel. | 05-12-2016 |
| 20160177740 | Gas Turbine Engine Component With Conformal Fillet Cooling Path | 06-23-2016 |
| 20160186994 | BOSS FOR COMBUSTOR PANEL - A combustor for use in a gas turbine engine has a combustor outer shell. A panel has an inner face which will face hot products of combustion, and a boss surrounding a feature, with the boss extending to an outer end. A spacing surface is spaced from the boss, and is at an outer position that is inward of the outer end of the boss. The spacing surface spaces the panel from the outer shell. A trough is intermediate the boss and the spacing surface. The trough extends to an outer end which is inward of the outer position of the spacing surface. A gas turbine engine is also disclosed. | 06-30-2016 |
| 20160201559 | TUBULAR COMBUSTION CHAMBER HAVING A FLAME TUBE END REGION AND GAS TURBINE | 07-14-2016 |
| 20160201912 | WEDGE-SHAPED CERAMIC HEAT SHIELD OF A GAS TURBINE COMBUSTION CHAMBER | 07-14-2016 |
| 20160201913 | HYBRID THROUGH HOLES AND ANGLED HOLES FOR COMBUSTOR GROMMET COOLING | 07-14-2016 |
| 060756000 | Air directed to flow along inner surface of liner dome | 2 |
| 20090255267 | COMUBSTOR SEAL HAVING MULTIPLE COOLING FLUID PATHWAYS - A combustor for a gas turbine includes a first combustor component and a second combustor component. The second combustor component is at least partially insertable into the first combustor component, and the first combustor component and second combustor component define a combustion fluid pathway. A combustor seal is located between the first combustor component and the second combustor component. The combustor seal defines at least one inner cooling pathway between the combustor seal and the second combustor component and at least one outer cooling pathway between the combustor seal and the first combustor component for cooling the first combustor component and second combustor component. A method for cooling a first combustor component and a second combustor component is also disclosed. | 10-15-2009 |
| 20140090385 | SYSTEM AND METHOD FOR SWIRL FLOW GENERATION - A system includes a turbine combustor containing a first wall portion disposed about a combustion zone, a second wall portion disposed about the first wall portion, and a plurality of vanes disposed within an annulus between the first and second wall portions. The plurality of vanes is configured to swirl air flowing within the annulus. | 04-03-2014 |
| 060757000 | In an axial direction | 6 |
| 20090077977 | COMBUSTION CHAMBER OF A TURBOMACHINE - Annular combustion chamber of a turbomachine, comprising two walls of revolution each having an annular groove extending around the longitudinal axis of the chamber and opening out inside the chamber, this groove having in cross section a substantially U or V shape that is splayed towards the downstream end and having lateral annular surfaces that are inclined with respect to the area of the wall at which the groove is situated. | 03-26-2009 |
| 20090255268 | Divergent cooling thimbles for combustor liners and related method - A cooling arrangement for a turbine combustor liner includes a combustor liner; a flow sleeve surrounding at least a portion of the combustor liner with a flow annulus therebetween, the flow sleeve having a plurality of rows of cooling holes formed about a circumference thereof for directing cooling air into the flow annulus and toward the combustor liner. One or more of the cooling holes is fitted with a thimble extending in a radial direction toward the combustor liner, the thimble having a peripheral wall diverging in a direction of flow of the cooling air. | 10-15-2009 |
| 20090282833 | Method and apparatus for cooling and dilution tuning a gas turbine combustor liner and transition piece interface - A combustor liner includes a forward end and an aft end, the aft end having a reduced diameter portion and a cooling and dilution sleeve overlying the reduced diameter portion thereby establishing a cooling plenum therebetween. A plurality of cooling and dilution air entry holes are formed in the cooling and dilution sleeve and a plurality of cooling and dilution air exit holes formed adjacent an aft edge of the liner such that, in use, cooling and dilution air flows through the cooling and dilution air entry holes, and through the plenum, exiting the cooling and dilution air exit holes, thereby cooling and dilution tuning the aft end of the combustor liner without having to remove the transition piece. | 11-19-2009 |
| 20110247341 | COMBUSTOR LINER HELICAL COOLING APPARATUS - A combustor liner is provided. The combustor liner may include an upstream portion and a downstream end portion. The upstream portion may have a radius and a length along a generally longitudinal axis. The downstream end portion may have a radius and a length along the generally longitudinal axis. The downstream end portion may define a plurality of channels. Each of the plurality of channels may extend helically through the length of the downstream end portion. Each of the plurality of channels may be configured to flow an air flow therethrough, cooling the downstream end portion. | 10-13-2011 |
| 20150338103 | TURBINE ENGINE WALL HAVING AT LEAST SOME COOLING ORIFICES THAT ARE PLUGGED - A turbine engine wall having a cold side and a hot side and including a plurality of cooling orifices for enabling air flowing on the cold side of the wall to penetrate to the hot side at least some of the cooling orifices being plugged by a plugging material so as to define a minimum level of porosity for the wall corresponding to putting the turbine engine into service, and the plugged cooling orifices being suitable for being unplugged progressively throughout the lifetime of the turbine engine in order to define a maximum level of porosity for the wall corresponding to an end of lifetime for the turbine engine, the plugging being performed by alternating at least one of the following rows or lines: circumferential rows, axial rows, diagonal lines, so as to lie in the range one-third to one-half of the maximum porosity. | 11-26-2015 |
| 20170234226 | Cooled Combustor Case with Over-Pressurized Cooling Air | 08-17-2017 |
