| Class / Patent application number | Description | Number of patent applications / Date published |
|---|
| 060754000 | Porous | 86 |
| 20080264065 | Gas-turbine combustion chamber wall - A gas-turbine combustion chamber wall for a gas-turbine has a combustion chamber wall | 10-30-2008 |
| 20080271457 | Cooling Holes For Gas Turbine Combustor Having A Non-Uniform Diameter Therethrough - A gas turbine combustor liner, including a shell having a first end adjacent to an upstream end of the combustor and a second end adjacent to a downstream end of the combustor, where the shell also has a hot side, a cold side, and a centerline axis therethrough. A plurality of small, closely-spaced film cooling holes are formed in the shell through which air flows for providing a cooling film along the hot side of the shell. Each cooling hole has a non-uniform diameter as it extends through the shell. In particular, each cooling hole includes a first opening located at the cold side of the shell having a first diameter and a second opening located at the hot side of the shell having a second diameter, wherein the second diameter of the second opening is larger than the first diameter of the first opening. It is preferred that the shape of each cooling hole be substantially frusto-conical. | 11-06-2008 |
| 20080314044 | HEAT SHIELDS FOR USE IN COMBUSTORS - A combustor includes an inner liner; an outer liner circumscribing the inner liner and forming a combustion chamber with the inner liner; a combustor dome coupled to the inner and outer liners; and a plurality of heat shields coupled to combustor dome. Each of the heat shields includes a heat shield plate defined by a first edge facing the inner liner and a second edge facing the outer liner; and a plurality of baffles extending from the heat shield plate. Each of the plurality of baffles includes two ribs and a connection portion connecting the two ribs to form a closed portion and an opposite open portion. The open portion of each of the plurality of baffles faces the first edge or the second edge. | 12-25-2008 |
| 20090013695 | Floatwell Panel Assemblies and Related Systems - Floatwall panel assemblies and related systems are provided. A floatwall panel assembly includes a panel formed of porous ceramic material, the porous ceramic material exhibiting a porosity gradient along at least one of a length, a width and a depth of the panel, the panel lacking a substrate, formed of a material other than porous ceramic material, for supporting the porous ceramic material. | 01-15-2009 |
| 20090071161 | COMBUSTORS AND COMBUSTION SYSTEMS FOR GAS TURBINE ENGINES - A combustion system includes a combustor having a forward annular liner having a first plurality of effusion holes, and an aft annular liner having a second plurality of effusion holes and forming a combustion chamber with the forward annular liner. The first plurality of effusion holes and the second plurality of effusion holes are adapted to receive compressed air from a compressor and contribute to a single toroidal recirculation air flow pattern in the combustion chamber. The combustion system further includes a rotary fuel slinger further adapted to receive a flow of fuel from a fuel source and to centrifuge the received fuel into the combustion chamber; and an igniter extending at least partially into the combustion chamber to ignite the fuel and compressed air in the combustion chamber, to thereby generate combusted gas. | 03-19-2009 |
| 20090071162 | PERIPHERAL COMBUSTOR FOR TIP TURBINE ENGINE - A tip turbine engine ( | 03-19-2009 |
| 20090071163 | SYSTEMS AND METHODS FOR INSTALLING COOLING HOLES IN A COMBUSTION LINER - A system and method for modification of cooling holes in a combustion liner to prevent cracking. The invention provides a portable combustion liner modification system for a gas turbine able to allow for the correct placement of cooling holes on the combustion liner preventing cracks from forming around the mixing hole. | 03-19-2009 |
| 20090100839 | COMBUSTION CHAMBER WALL WITH OPTIMIZED DILUTION AND COOLING, AND COMBUSTION CHAMBER AND TURBOMACHINE BOTH PROVIDED THEREWITH - The invention applies to the field of turbomachines and relates to a combustion chamber in which the supply of dilution air and cooling air is optimized. The invention is concerned more particularly with optimizing the position of the dilution holes present on the walls of the combustion chamber. | 04-23-2009 |
| 20090100840 | COMBUSTION CHAMBER WITH OPTIMISED DILUTION AND TURBOMACHINE PROVIDED WITH SAME - The invention relates to the field of turbomachines and concerns a combustion chamber ( | 04-23-2009 |
| 20090120095 | Combustion liner thimble insert and related method - A gas turbine combustor liner has at least one circumferential row of air holes adapted to supply air in a radial direction into a combustion chamber within the liner. One or more of the air holes have a thimble fixed therein, the thimble having a substantially circular body and a pair of lips extending from an interior end of the thimble in diametrically opposed upstream and downstream directions. | 05-14-2009 |
| 20090133404 | SYSTEMS AND METHODS FOR COOLING GAS TURBINE ENGINE TRANSITION LINERS - A gas turbine engine assembly includes a combustor configured to combust an air-fuel mixture to produce combustion gases in a first direction; a transition liner coupled to the combustor and adapted to receive the combustion gases from the combustor and to redirect the combustion gases in a second direction; and a turbine coupled to the transition liner and adapted to receive the combustion gases from the transition liner. The transition liner has a plurality of effusion holes that include a first group that extend at least partially in a tangential direction. | 05-28-2009 |
| 20090188256 | EFFUSION COOLING FOR GAS TURBINE COMBUSTORS - A combustor for a turbine engine includes an outer liner and an inner liner circumscribed by the outer liner to form a combustion chamber therewith in which an air and fuel mixture is combusted to form streamlines of combustion gases. A first plurality of effusion cooling holes is formed in at least one of the inner or outer liners, with the first plurality of effusion cooling holes being oriented as a function of the streamlines. | 07-30-2009 |
| 20100018211 | Gas turbine transition piece having dilution holes - A gas turbine transition piece includes a duct body having a forward end and an aft end, the duct body defining an enclosure for confining a flow of combustion products from a combustor to a turbine first stage nozzle. A plurality of dilution holes are formed in the duct body, located at selected X, Y, Z coordinates measured from a zero reference point at a center of an exit plane of the transition piece. | 01-28-2010 |
| 20100037622 | Contoured Impingement Sleeve Holes - A combustor for use with a gas turbine. The combustor may include liner, an impingement sleeve, and with the liner and the impingement sleeve defining an airflow channel. The impingement sleeve may include a number of contoured holes therethrough. | 02-18-2010 |
| 20100050650 | GAS TURBINE ENGINE REVERSE-FLOW COMBUSTOR - A gas turbine engine reverse-flow sheet metal combustor for an aero gas turbine engine has a louverless single-piece liner design which includes a plurality of effusion patches, containing multiple rows of effusion cooling holes, bounded by cooled effusionless bands to aid in repairability of the combustor. The cooling of the bands improves the durability of the combustor and provides a method of repair for the combustor. | 03-04-2010 |
| 20100077763 | COMBUSTOR WITH IMPROVED COOLING HOLES ARRANGEMENT - A gas turbine engine combustor liner with at least one of the inner and outer liners that is effusion cooled and has a row of groups of circumferentially spaced apart dilution holes defined therethrough. Each group is located within a respective zone of the combustor liner defined by an overlap of adjacent conical sections corresponding to the sprays of adjacent fuel nozzles. | 04-01-2010 |
| 20100077764 | Structures with adaptive cooling - Disclosed are exemplary structures with adaptive cooling and methods of adaptively cooling structures. A liner is affixed to a support and the liner deflects away from the support when exposed to a localized hot spot in a hot fluid stream. The liner deflection creates a chamber between the liner and support, allowing cooling air to impinge against the liner, thus mitigating the effects of the hot spot. By providing impingement cooling only where needed, the amount of air needed for cooling is reduced. | 04-01-2010 |
| 20100095680 | DUAL WALL STRUCTURE FOR USE IN A COMBUSTOR OF A GAS TURBINE ENGINE - A dual wall structure for a combustor of a gas turbine engine including an inner liner and an outer liner coupled to a combustor dome and defining a combustion chamber there between. Each of the inner liner and the outer liner include an outer wall and an inner wall. Each of the outer walls includes a plurality of impingement holes formed therein for allowing a coolant to flow therethrough. Each of the inner walls is coupled to the outer wall and a combustor dome and includes a plurality of heat shield panels. Each of the plurality of heat shield panels extends a longitudinal length of the combustor chamber and includes a plurality of side rails, an aft rail and a forward flange that when coupled to the outer wall defines a single cavity there between. A plurality of cavities being formed by the plurality of heat shield panels. | 04-22-2010 |
| 20100122537 | COMBUSTORS WITH INSERTS BETWEEN DUAL WALL LINERS - A combustor 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 the first liner is a first dual wall liner having a first hot wall facing the combustion chamber and a first cold wall. The first cold wall has a first cold wall orifice and the first hot wall has a first hot wall orifice. A first insert is mounted in the first hot wall orifice and is configured to receive a first air jet of pressured air from the first cold wall orifice, and guide the first jet through the first hot wall orifice and into the combustion chamber. | 05-20-2010 |
| 20100170256 | RING COOLING FOR A COMBUSTION LINER AND RELATED METHOD - A gas turbine combustor includes a liner having a forward end and an aft end; a flow sleeve surrounding the liner, the flow sleeve also having forward and aft ends, the aft end of the flow sleeve supporting an annular ring formed with a plurality of cooling bores and extending through the flow sleeve, at least some of the plurality of cooling bores formed at an acute angle relative to a longitudinal axis of the liner. | 07-08-2010 |
| 20100170257 | COOLING A ONE-PIECE CAN COMBUSTOR AND RELATED METHOD - A cooling arrangement for cooling a single-piece, combined combustor liner/transition piece substantially enclosed within a surrounding flow sleeve, with a cooling annulus radially between the flow sleeve and the single-piece combined combustor liner/transition piece, the cooling arrangement including a first plurality of impingement cooling holes in the flow sleeve, the plurality of impingement cooling holes having first diameters and arranged to direct cooling air onto designated areas of the single-piece, combined combustor liner/transition piece; and a second plurality of effusion cooling holes in the single-piece, combined combustor liner/transition piece having second diameters smaller than the first diameters, and located to cool by effusion other areas of the single-piece, combined combustor liner/transition piece. | 07-08-2010 |
| 20100218503 | PLUNGED HOLE ARRANGEMENT FOR ANNULAR RICH-QUENCH-LEAN GAS TURBINE COMBUSTORS - A combustor may include an outer liner having a first row and a second row of circumferentially distributed air admission holes. The second row of the outer liner may be downstream of the first row of the outer liner, and the air admission holes of the first row of the outer liner may be larger than the air admission holes of the second row of the outer liner. An inner liner is circumscribed by the outer liner and has a third and fourth row of circumferentially distributed air admission holes. The air admission holes of the third row of the inner liner may be larger than the air admission holes of the fourth row of the inner liner. The inner and outer liners may form a combustion chamber, and at least a portion of the air admission holes of the first, second, third, or fourth rows may be plunged. | 09-02-2010 |
| 20100218504 | ANNULAR RICH-QUENCH-LEAN GAS TURBINE COMBUSTORS WITH PLUNGED HOLES - A combustor may include an outer liner having a first group of air admission holes and defining a plurality of outer liner regions. The combustor may further include an inner liner having a second group of air admission holes and defining a plurality of inner liner regions. The first group of air admission holes within a respective outer liner region may include a first plunged air admission hole approximately axially aligned with the respective fuel injector, a second plunged air admission hole approximately on the outer boundary line between the respective outer liner region and a first adjacent outer liner region, the first air admission hole being downstream of the second air admission hole, and a third plunged air admission hole approximately on the outer boundary line between the respective outer liner region and a second adjacent outer liner region. | 09-02-2010 |
| 20110005233 | COMBUSTION CHAMBER HEAD OF A GAS TURBINE - A combustion chamber head of a gas turbine has a confinement enclosing a dampening volume ( | 01-13-2011 |
| 20110048024 | GAS TURBINE COMBUSTOR WITH QUENCH WAKE CONTROL - A gas turbine engine has a combustor module including an annular combustor having a liner assembly that defines an annular combustion chamber and includes a circumferential row of a plurality of relatively large combustion dilution air admission holes and a circumferential row of a plurality of smaller quench air admission holes disposed downstream with respect to the flow of combustion gas products. The plurality of quench air admission holes are arranged with respect to the plurality of relatively large dilution air admission holes disposed upstream thereof such that there is associated with each dilution air admission hole a first quench air admission hole and a second quench air hole, the first quench air hole being offset laterally in a first lateral direction and the second quench air hole being offset laterally in a second lateral direction opposite to the first direction. | 03-03-2011 |
| 20110067406 | IMPINGEMENT COOLED CROSSFIRE TUBE ASSEMBLY - A crossfire tube assembly is configured for connecting adjacent combustion cans in a gas turbine, and includes a first tube segment having a first end and an opposite female end. A second tube segment has a first end and an opposite male end fitted concentrically within the female end with an overlap region between the female and male ends. Each of the first ends of the tube segments is configured for securing to a liner of a respective combustion can. Oppositely oriented first and second impingement sleeves extend from the female end of the first tube segment to the respective first ends of the tube segments, Combustion cooling air is directed through metering holes in the impingement sleeves and flows axially along concentric cavities defined between the impingement sleeves and the first and second tube segments. The combustion cooling air vents from the cavities into an axial combustion air flow stream between the respective combustion can liners and sleeves. | 03-24-2011 |
| 20110120134 | GAS TURBINE COMBUSTOR - An annular combustor for a gas turbine has a combustion chamber having an interior volume that, in longitudinal section, includes a forward volume, an intermediate volume and an aft volume. The forward volume represents from about 30% to about 40% of the combustor interior volume, the intermediate volume represents from about 10% to about 20% of the combustor interior volume, and the aft volume represents from about 40% to about 60% of the combustor interior volume. | 05-26-2011 |
| 20110185738 | GAS TURBINE ENGINE COMPONENT CONSTRUCTION - A gas turbine engine component is disclosed having a construction that permits a working fluid to flow from one side of the component to the other. In one embodiment the construction includes multiple layers, and in one particular embodiment the construction includes three layers. One or more layers can have different properties. Additionally, one or more layers of the construction can include a cellular structure. In one embodiment the component is a gas turbine engine combustor liner. | 08-04-2011 |
| 20110214428 | HYBRID VENTURI COOLING SYSTEM - A venturi device for a turbine combustor includes a substantially annular outer liner; a substantially annular inner liner; a venturi channel located between the substantially annular outer and inner liners; the substantially annular outer and inner liners being substantially V-shaped in axial cross-section, thereby defining a throat region; the substantially annular outer liner formed with an array of impingement holes and the substantially annular inner liner formed with a plurality of vortex generators facing the substantially annular outer liner. | 09-08-2011 |
| 20120036858 | COMBUSTOR LINER COOLING SYSTEM - A combustor liner is disclosed. The combustor liner includes an upstream portion, a downstream end portion extending from the upstream portion along a generally longitudinal axis, and a cover layer associated with an inner surface of the downstream end portion. The downstream end portion includes the inner surface and an outer surface, the inner surface defining a plurality of microchannels. The downstream end portion further defines a plurality of passages extending between the inner surface and the outer surface. The plurality of microchannels are fluidly connected to the plurality of passages, and are configured to flow a cooling medium therethrough, cooling the combustor liner. | 02-16-2012 |
| 20120144834 | Gas Turbine Combustion Chamber - A gas turbine combustion chamber has a housing ( | 06-14-2012 |
| 20120167574 | GAS TURBINE ENGINE AND COMBUSTION LINER - One embodiment of the present invention is a unique gas turbine engine combustion liner. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine engine combustion liners. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. | 07-05-2012 |
| 20120255308 | COOLED DOUBLE WALLED ARTICLE - A gas turbine engine combustion chamber includes a first wall and a second wall. The second wall is arranged within and spaced from the first wall to define a cavity between the first wall and the second wall. The first wall has a plurality of impingement apertures extending there-through and the second wall has a plurality of effusion apertures extending there-through. The impingement apertures have a first diameter, a first pitch, and a first area. The effusion apertures have a second diameter, a second pitch, and a second area. The ratio of the first diameter to the second diameter is at least 3, the ratio of the first pitch to the second pitch is at least 4 and the ratio of the first area to the second area is at least 9. This arrangement increases the cooling performance of the effusion apertures in the second wall. | 10-11-2012 |
| 20120272653 | INTERNAL COMBUSTION ENGINE HOT GAS PATH COMPONENT WITH POWDER METALLURGY STRUCTURE - A hot gas path component ( | 11-01-2012 |
| 20120304659 | IMPINGEMENT SLEEVE AND METHODS FOR DESIGNING AND FORMING IMPINGEMENT SLEEVE - An impingement sleeve and methods for designing and forming an impingement sleeve are disclosed. In one embodiment, the impingement sleeve includes a body configured to at least partially surround a transition piece of the combustor. The impingement sleeve further includes a plurality of cooling holes defined in the body, the plurality of cooling holes having a cooling hole pattern configured to provide a desired operational value for the transition piece. At least one of the plurality of cooling holes has a chamfer extending at least partially between an inlet and an outlet of the at least one of the plurality of cooling holes. At least a portion of the plurality of cooling holes are generally longitudinally asymmetric. | 12-06-2012 |
| 20130036742 | COMBUSTOR LINER COOLING SYSTEM - A turbine engine with a combustor that includes a hollow wall about a combustor liner. The combustor liner includes an inner surface facing inwardly toward a combustion chamber. The turbine engine includes a first air flow path in an upstream direction through the hollow wall toward a head end of the combustor. The first air flow path includes a plurality of bypass openings extending through the combustor liner to the inner surface to supply a first cooling film to a downstream end portion of the combustor liner. The turbine engine further includes a second flow path in a second direction opposite the upstream direction through the hollow wall. The second flow path includes a plurality of film holes extending through the combustor liner to the inner surface to supply a second cooling film to the downstream end portion of the combustor liner downstream of the first cooling film. | 02-14-2013 |
| 20130074507 | COMBUSTION LINER FOR A TURBINE ENGINE - A combustion liner for a combustor of a turbine engine includes a plurality of undulations which extend around the exterior circumference of the combustion liner. A plurality of rows of cooling holes are formed through the combustion liner. Each row of cooling holes is located in one of the undulations which extends around the exterior circumference of the combustion liner. The cooling holes admit a flow of cooling air into the interior of the combustion liner. The cooling holes are located and oriented to help the flow of cooling air form a film along the inner surface of the combustion liner. | 03-28-2013 |
| 20130180252 | COMBUSTOR ASSEMBLY WITH IMPINGEMENT SLEEVE HOLES AND TURBULATORS - A combustor assembly for use with a gas turbine. The combustor assembly may include a liner, an impingement sleeve disposed about the liner, and an airflow channel defined between the liner and the impingement sleeve. One or more holes may be disposed through the impingement sleeve, and one or more tubulators may be disposed within the airflow channel. | 07-18-2013 |
| 20130205790 | MULTI-LOBED COOLING HOLE AND METHOD OF MANUFACTURE - A gas turbine engine component includes a cooling hole. The cooling hole includes an inlet, an outlet, a metering section and a diffusing section. The diffusing section extends from the metering section to the outlet and includes a first lobe diverging longitudinally and laterally from the metering section, a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section, and a transition region having a portion that extends between the first and second lobes and an end adjacent the outlet. | 08-15-2013 |
| 20130205791 | COOLING HOLE WITH CURVED METERING SECTION - A gas turbine engine component includes a cooling hole. The component includes a first wall having an inlet, a second wall having an outlet and a metering section extending downstream from the inlet and having a substantially convex first surface and a substantially concave second surface. The component also includes a diffusing section extending from the metering section to the outlet. A gas turbine engine wall includes first and second surfaces and a cooling hole extending between an inlet at the first surface and an outlet at the second surface. The cooling hole includes a metering section commencing at the inlet and a diffusing section in communication with the metering section and terminating at the outlet. The metering section includes a top portion having a first arcuate surface and a bottom portion having a second arcuate surface. The first and second arcuate surfaces have arcs extending in substantially similar directions. | 08-15-2013 |
| 20130205792 | COOLING HOLE WITH ASYMMETRIC DIFFUSER - A gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section. | 08-15-2013 |
| 20130205793 | GAS TURBINE ENGINE COMPONENT WITH IMPINGEMENT AND DIFFUSIVE COOLING - A gas turbine engine component includes a gas path wall having a first surface and second surface and an impingement baffle having impingement holes for directing cooling fluid onto the first surface of the gas path wall. A cooling hole extends through the gas path wall. The cooling hole continuously diverges from an inlet in the first surface to an outlet in the second surface such that cross-sectional area of the cooling hole increases continuously from the inlet to the outlet. A longitudinal ridge divides the cooling hole into lobes. | 08-15-2013 |
| 20130205794 | GAS TURBINE ENGINE COMPONENT WITH IMPINGEMENT AND LOBED COOLING HOLE - A gas turbine engine component includes a gas path wall having a first and second opposing surfaces and a baffle positioned along the gas path wall. The baffle has impingement holes for directing cooling fluid onto the first surface of the gas path wall. A cooling hole is formed in the gas path wall and extends from a metering section having an inlet in the first surface through a transition to a diffusing section having an outlet in the second surface. A longitudinal ridge extends along the cooling hole between the transition and the outlet. The longitudinal ridge divides the diffusing section of the cooling hole into first and second lobes. | 08-15-2013 |
| 20130232980 | SYSTEM FOR SUPPLYING A WORKING FLUID TO A COMBUSTOR - A system for supplying a working fluid to a combustor includes a combustion chamber, a liner that circumferentially surrounds at least a portion of the combustion chamber, and a flow sleeve that circumferentially surrounds at least a portion of the liner. A tube provides fluid communication for the working fluid to flow through the flow sleeve and the liner and into the combustion chamber, and the tube spirals between the flow sleeve and the liner. | 09-12-2013 |
| 20130255265 | COMBUSTORS WITH QUENCH INSERTS - A combustor for a turbine engine is provided. The combustor includes a first liner and a second liner forming a combustion chamber with the first liner, the combustion chamber configured to receive an air-fuel mixture for combustion therein. The first liner defining an air admission hole for directing a first jet of pressurized air into the combustion chamber, and the air admission hole may have a non-circular shape. The combustor further includes an insert positioned within the air admission hole to guide the first jet through the air admission hole and into the combustion chamber. | 10-03-2013 |
| 20130255266 | Transition Nozzle Combustion System - The present application provides a combustion system for use with a cooling flow. The combustion system may include a head end, an aft end, a transition nozzle extending from the head end to the aft end, and an impingement sleeve surrounding the transition nozzle. The impingement sleeve may define a first cavity in communication with the head end for a first portion of the cooling flow and a second cavity in communication with the aft end for a second portion of the cooling flow. The transition nozzle may include a number of cooling holes thereon in communication with the second portion of the cooling flow. | 10-03-2013 |
| 20130269354 | Substrate with Shaped Cooling Holes and Methods of Manufacture - A substrate having one or more shaped effusion cooling holes formed therein. Each shaped cooling hole has a bore angled relative to an exit surface of the combustor liner. One end of the bore is an inlet formed in an inlet surface of the combustor liner. The other end of the bore is an outlet formed in the exit surface of the combustor liner. The outlet has a shaped portion that expands in only one dimension. Also methods for making the shaped cooling holes. | 10-17-2013 |
| 20130283806 | COMBUSTOR AND A METHOD FOR REPAIRING THE COMBUSTOR - A combustor generally includes a liner that at least partially defines a combustion chamber within the combustor. The liner may generally include a hole that extends through the liner and that at least partially defines a mating surface. An insert having an outer surface extends at least partially through the hole. A projection that at least partially defines a mating surface at least partially circumferentially surrounds the insert outer surface. A joint between the projection mating surface and the hole mating surface provides a connection point between the projection mating surface and the liner hole mating surface. | 10-31-2013 |
| 20130340437 | TURBINE ENGINE COMBUSTOR WALL WITH NON-UNIFORM DISTRIBUTION OF EFFUSION APERTURES - A turbine engine combustor wall includes support shell and a heat shield. The support shell includes shell quench apertures, first impingement apertures, and second impingement apertures. The combustor heat shield includes shield quench apertures fluidly coupled with the shell quench apertures, first effusion apertures fluidly coupled with the first impingement apertures, and second effusion apertures fluidly coupled with the second impingement apertures. The shield quench apertures and the first effusion apertures are configured in a first axial region of the heat shield, and the second effusion apertures are configured in a second axial region of the heat shield located axially between the first axial region and a downstream end of the heat shield. A density of the first effusion apertures in the first axial region is greater than a density of the second effusion apertures in the second axial region. | 12-26-2013 |
| 20140007580 | RETAINING COLLAR FOR A GAS TURBINE COMBUSTION LINER - A gas turbine combustion liner having a novel and improved system for receiving a peripheral device through a floating collar is disclosed. The retaining collar assembly provides planar movement for receiving the peripheral device while preventing the floating collar from rotational movement through an anti-rotation tab on the collar and a collar cap, resulting in reduced wear to the peripheral device. The collar cap can include different configurations for reducing any vibratory effects caused by an oncoming airflow directly contacting the floating collar. | 01-09-2014 |
| 20140020393 | COMBUSTOR FOR GAS TURBINE ENGINE AND GAS TURBINE - A combustor for gas turbine engine, wherein: a plurality of impingement-cooling holes are opened passing through an external wall of a liner for cooling air to blow out towards the outer surface of an internal wall of the liner; a plurality of pin-fins are formed on the outer surface of the internal wall of the liner; a plurality of effusion cooling holes are opened passing through the internal wall of the liner for cooling air to blow out along the inner surface of the internal wall of the liner. The top surface of each pin-fin is not in contact with the inner surface of the external wall of the liner and the ratio of the height of the pin-fin to the equivalent diameter of the impingement-cooling hole is set to be 1.0-3.0. | 01-23-2014 |
| 20140047845 | COMBUSTOR LINER COOLING ASSEMBLY - A combustor liner cooling assembly includes a combustor liner defining a combustor chamber. Also included is a cover sleeve spaced radially outwardly from and at least partially surrounding an aft end of the combustor liner, the cover sleeve and the combustor liner defining a cooling annulus. Further included is at least one aperture extending through the cover sleeve for routing a cooling flow to the cooling annulus. Yet further included is a perforated sleeve disposed between the cover sleeve and the combustor liner, wherein the perforated sleeve comprises a plurality of holes for impinging the cooling flow toward the combustor liner. | 02-20-2014 |
| 20140090384 | GAS TURBINE ENGINE COOLING HOLE WITH CIRCULAR EXIT GEOMETRY - A gas turbine engine component includes a structure having an exterior surface. A cooling hole extends from a cooling passage to the exterior surface to provide an exit area on the exterior surface that is substantially circular in shape. A gas turbine engine includes a compressor section and a turbine section. A combustor is provided between the compressor and turbine sections. A component in at least one of the compressor and turbine sections has an exterior surface. A film cooling hole extends from a cooling passage to the exterior surface to provide an exit area that is substantially circular in shape. A method of machining a film cooling hole includes providing a component having an internal cooling passage and an exterior surface, machining a film cooling hole from the exterior surface to the internal cooling passage to provide a substantially circular exit area on the exterior surface. | 04-03-2014 |
| 20140096527 | GAS TURBINE ENGINE COMBUSTOR LINER - A liner for a combustor of a turbine engine includes a cooling feature which projects from a backside and an effusion hole that communicates through the liner | 04-10-2014 |
| 20140116058 | ASSEMBLIES AND APPARATUS RELATED TO COMBUSTOR COOLING IN TURBINE ENGINES - A combustor of a combustion turbine engine is described. The combustor may include an inner radial wall, which defines a combustion chamber downstream of a primary fuel nozzle, and an outer radial wall, which surrounds the inner radial wall so to form a flow annulus therebetween, and the combustor may include a socket extending from the outer radial wall into the flow annulus. The socket may include: a mouth formed through the outer radial wall; a floor offset a predetermined distance from an outboard surface of the inner radial wall; impingement ports formed through the floor; and an axial nozzle. | 05-01-2014 |
| 20140116059 | HOT GAS SEGMENT ARRANGEMENT - A hot gas segment arrangement, especially for a combustion chamber of a gas turbine, that includes at least one hot gas segment, which is removably mounted on a carrier, and is subjected at its outside to hot gas and impingement-cooled at its inside, whereby an impingement plate with a plurality of distributed impingement holes is arranged in a distance at the inside of the impingement plate. A cooling air supply means is provided for loading the impingement plate with pressurized cooling air in order to generate through the impingement holes jets of cooling air, which impinge on the inside of the hot gas segment. The cooling efficiency and lifetime are increased by the impingement plate being part of a closed receptacle, which is supplied with the pressurized cooling air, and by the receptacle with the impingement plate being mounted on the carrier independently of the hot gas segment. | 05-01-2014 |
| 20140144147 | TRANSITION PIECE OF COMBUSTOR, AND GAS TURBINE HAVING THE SAME - A transition piece of a combustor that sends high temperature combustion gas to a turbine includes a cylindrical body wall and cooling air passages. The passages are formed in the body wall so as to extend in an axial direction thereof, and each of the passages has cooling air inlet ports opened at an outer circumferential surface of the transition piece and cooling air outlet ports opened at an inner circumferential surface of the transition piece. The cooling air outlet ports form a plurality of lines in a direction oblique with respect to the axial direction of the body wall. A first distance between a first line of the cooling air outlet ports and a second line of the cooling air outlet ports adjacent to the first line is larger than a second distance between the cooling air outlet ports adjacent to each other. | 05-29-2014 |
| 20140174092 | Closure of Cooling Holes with a Filling Agent - A method for filling cooling holes in a component of a gas turbine engine is disclosed. The component may include a plurality of first cooling holes extending through the wall of the component. The method may comprise the steps of exposing the outer surface of the component, filling the plurality of first cooling holes with a polyimide, curing the polyimide to block the passage of cooling fluid through the plurality of first cooling holes, and applying a thermal bather coating over the outer surface of the component. The method may further include the step of installing a second plurality of cooling holes in the wall of the component wherein the plurality of second cooling holes penetrate the thermal barrier coating and the wall of the component and allow cooling fluid to pass therethrough. | 06-26-2014 |
| 20140216042 | COMBUSTOR COMPONENT WITH COOLING HOLES FORMED BY ADDITIVE MANUFACTURING - Combustor liners made using additive manufacturing techniques can employ cooling hole patterns which are not possible, or at least time consuming or expensive, to make using traditional subtractive manufacturing techniques. By additively manufacturing floatwall panels, cooling holes may be placed along axes that transect features on the floatwall panel, such as mounting studs, spacers, cooling pedestals, and rails. | 08-07-2014 |
| 20140238029 | FLOW CONDITIONER IN A COMBUSTOR OF A GAS TURBINE ENGINE - A combustor in a gas turbine includes a liner having an interior volume defining a main combustion zone, a fuel injection system for delivering fuel into the main combustion zone, and a flow sleeve that defines, with the liner, a passageway for air to flow on its way to be mixed with fuel from the fuel injection system, wherein the mixture is burned in the main combustion zone to create hot combustion gases. The combustor further includes a flow conditioner including at least one panel having a configuration such that air is able to pass through the panel(s) on its way to the passageway, wherein at least a substantial portion of the air that enters the passageway for being burned in the main combustion zone passes through the panel(s). | 08-28-2014 |
| 20140238030 | IMPINGEMENT-EFFUSION COOLED TILE OF A GAS-TURBINE COMBUSTION CHAMBER WITH ELONGATED EFFUSION HOLES - The present invention relates to a gas-turbine combustion chamber having a combustion chamber wall including a tile carrier, on which wall tiles are mounted at a distance to form an impingement cooling gap, where the tile carrier has impingement cooling holes and the tile is provided with effusion holes, where the tile has on its side facing the tile carrier a surface structure which raises from the surface of the tile and extends in the direction of the tile carrier. | 08-28-2014 |
| 20140238031 | COMBUSTOR LINER - The invention is a combustor liner ( | 08-28-2014 |
| 20140250894 | DUAL-WALL IMPINGEMENT, CONVECTION, EFFUSION (DICE) COMBUSTOR TILE - A gas turbine engine includes a combustor having a dual-wall impingement convention effusion combustor tile assembly. The dual-wall tile assembly provides a cooling air flow channel and attachments for securing the tile to the cold skin liner of the combustor. Cooling is more efficient in part due to the dual wall construction and in part due to reduced parasitic leakage, and the design is less sensitive to attachment features which operate at lower temperatures. | 09-11-2014 |
| 20140260282 | GAS TURBINE ENGINE COMBUSTOR LINER - A gas turbine engine variable porosity combustor liner has a laminated alloy structure. The laminated alloy structure has combustion chamber facing holes on one side and cooling plenum facing holes on a radially opposite side. The combustion chamber facing holes are in fluid communication with the cooling plenum facing holes via axially and circumferentially extending flow passages sandwiched between metal alloy sheets of the laminated alloy structure. Porous zones having respective different cooling flow amounts are formed in the laminated alloy structure based on at least one of an arrangement of the combustion chamber facing holes, an arrangement of the cooling plenum facing holes, and an arrangement of the flow passages. | 09-18-2014 |
| 20140290258 | METHOD FOR THE ARRANGEMENT OF IMPINGEMENT COOLING HOLES AND EFFUSION HOLES IN A COMBUSTION CHAMBER WALL OF A GAS TURBINE - A method for the arrangement of effusion holes and impingement cooling holes in a combustion chamber wall including: distribution of the effusion holes in the surface to be cooled in accordance with pattern, diameter and dimension selections made; distribution of the impingement cooling holes in accordance with pattern, diameter and dimension selections made; checking of the number of matches and their spacing from one another, taking into account the component and assembly tolerances; comparison with the permitted number and their minimum spacing; and, if quality requirements are not met, taking corrective actions, including selecting alternative diameters and patterns. | 10-02-2014 |
| 20140338346 | COMBUSTOR SKIN ASSEMBLY FOR GAS TURBINE ENGINE - A combustor assembly includes a hot skin of a combustion chamber wall having an inner face exposed to the combustion chamber and an opposite outer face, a receiving skin having a securing portion affixed to the hot skin outer face in an air-tight manner and a receiving flange, extending from the securing portion, that is offset from the hot skin outer face to form a female recess, a cold skin having a cold wall portion spaced from the hot skin and forming a cooling cavity therebetween, a securing portion extending from a first end of the cold wall portion affixed to the hot skin outer face in an air-tight manner and a male flange extending from a second end of the cold wall portion opposite the first end, the male flange snugly received in the female recess and forming a sliding engagement therebetween. | 11-20-2014 |
| 20140338347 | COMBUSTORS WITH COMPLEX SHAPED EFFUSION HOLES - A combustor is provided for a turbine engine. The combustor includes a first liner having a first side and a second side and a second liner having a first side and a second side. The second side of the second liner forms a combustion chamber with the second side of the first liner, and the combustion chamber is configured to receive an air-fuel mixture for combustion therein. The first liner defines a plurality of effusion cooling holes configured to form a film of cooling air on the second side of the first liner. The plurality of effusion cooling holes including a first effusion cooling hole extending from the first side to the second side with a non-linear line of sight. | 11-20-2014 |
| 20140366545 | GAS TURBINE ENGINE - In a gas turbine engine ( | 12-18-2014 |
| 20140373549 | COOLING HOLE WITH CURVED METERING SECTION - A gas turbine engine component includes a cooling hole. The component includes a first wall having an inlet, a second wall having an outlet and a metering section extending downstream from the inlet and having a substantially convex first surface and a substantially concave second surface. The component also includes a diffusing section extending from the metering section to the outlet. A gas turbine engine wall includes first and second surfaces and a cooling hole extending between an inlet at the first surface and an outlet at the second surface. The cooling hole includes a metering section commencing at the inlet and a diffusing section in communication with the metering section and terminating at the outlet. The metering section includes a top portion having a first arcuate surface and a bottom portion having a second arcuate surface. The first and second arcuate surfaces have arcs extending in substantially similar directions. | 12-25-2014 |
| 20150007573 | ANNULAR-COMBUSTION-CHAMBER BYPASS - An annular combustion chamber ( | 01-08-2015 |
| 20150013340 | GAS TURBINE ENGINE COMBUSTOR LINER - A gas turbine engine variable porosity combustor liner has a laminated alloy structure. The laminated alloy structure has combustion chamber facing holes on one side and cooling plenum facing holes on a radially opposite side. The combustion chamber facing holes are in fluid communication with the cooling plenum facing holes via axially and circumferentially extending flow passages sandwiched between metal alloy sheets of the laminated alloy structure. Porous zones having respective different cooling flow amounts are formed in the laminated allow structure based on at least one of an arrangement of the combustion chamber facing holes, an arrangement of the cooling plenum facing holes, and an arrangement of the flow passages. | 01-15-2015 |
| 20150027128 | HEAT-SHIELD ELEMENT FOR A COMPRESSOR-AIR BYPASS AROUND THE COMBUSTION CHAMBER - A heat-shield element ( | 01-29-2015 |
| 20150040570 | GAS TURBINE ENGINE COMPONENT HAVING FOAM CORE AND COMPOSITE SKIN WITH COOLING SLOT - In one embodiment, a gas turbine engine component includes a foam based core and a composite skin member. Both the foam based core and the composite skin member can be used to structurally support the gas turbine engine component. The composite skin member can be a CMC material and is used to partially encapsulate the foam core. The gas turbine engine component can take the form of an airfoil member such as a blade or a vane, a combustor liner, etc. A first portion of the composite skin member includes a first surface extending past an edge of the component creating a step approximate an edge section. In another embodiment, composite skin members can be used to form a continuous shape for the edge section such that the foam core forms part of a gas path surface. | 02-12-2015 |
| 20150101335 | HOLE ARRANGEMENT OF LINERS OF A COMBUSTION CHAMBER OF A GAS TURBINE ENGINE WITH LOW COMBUSTION DYNAMICS AND EMISSIONS - A gas turbine combustion chamber with an inner housing and an outer housing the inner housing having an inner wall element with a first hole arrangement and a second hole arrangement is provided. The inner wall element envelopes a burner volume. The first hole arrangement has first holes arranged in a first areal density, the second hole arrangement has second holes arranged in a second areal density. The outer housing has an outer wall element with a further first hole arrangement and a further second hole arrangement. The outer wall element of the outer housing envelops the inner wall element of the inner housing so that a gap in between is formed. The further first hole arrangement has further first holes arranged in a further first areal density, the further second hole arrangement has further second holes arranged in a further second areal density. | 04-16-2015 |
| 20150121885 | Gas Turbine Combustor - There is provided a gas turbine combustor that achieves improved product reliability and a reduced increase in pressure loss through improvements made on a cooling characteristic and structural intensity. | 05-07-2015 |
| 20150128602 | HEAT SHIELD FOR A GAS TURBINE COMBUSTION CHAMBER - The present invention relates to a combustion chamber heat-shielding element of a gas-turbine, having a bolt for mounting the combustion chamber heat-shielding element on a combustion chamber wall or a combustion chamber head, where the combustion chamber heat-shielding element is designed substantially plate-like and where on one side at least one bolt, which is designed as a separate component, is anchored on it by means of a bonded connection. | 05-14-2015 |
| 20150369486 | Heat-Transfer Device and Gas Turbine Combustor with Same - Disclosed is a heat-transfer device adapted to enhance uniformity of cooling characteristics to be given to a heat transfer object, and thereby to extend a life of the heat transfer object. The heat-transfer device for facilitating heat exchange between combustion air (heat transfer medium) flowing along an outer surface of a combustor liner (heat transfer object), and the combustor liner, the heat-transfer device including at least one longitudinal vortex generating device protruding toward a annular passage (flow passage) of the combustion air and formed to generate a longitudinal vortex E with a central axis in a flow direction of the combustion air, and stir the combustion air flowing in the annular passage; and at least one radiator fin provided in a region A on the outer surface of the combustor liner, the region A being where a flow of the vortex E, on a swirling plane thereof, that are generated by the vortex generating device is directed from a side of the combustor liner, toward a side of a flow sleeve, the fin exchanging heat with the combustion air stirred by the vortex generating device. | 12-24-2015 |
| 20150377134 | COMBUSTOR COOLING STRUCTURE - The invention relates to a transition piece assembly for a gas turbine. A transition piece having one end adapted for connection to a gas combustor and an opposite end adapted for connection to a first turbine stage. The transition piece having at least one external liner and at least one internal liner. The internal liner forms the hot gas flow channel. A first section of the transition piece assembly upstream of a first turbine stage has a plurality of cooling apertures. Cooling medium through the cooling apertures enters the plenum, which is created between external and internal liners and a cooling medium flows along at least the first section of the transition piece assembly. At least one second section upstream of the first section with respect to the hot gas flow has at least one additional air inlet system. | 12-31-2015 |
| 20150377488 | CONICAL-FLAT HEAT SHIELD FOR GAS TURBINE ENGINE COMBUSTOR DOME - A gas turbine engine combustor conical-flat heat shield includes an annular conical section extending upstream from and being integral with a flat section with a flat downstream facing surface which may be generally perpendicular to or canted with respect to a centerline. The flat section includes radially outer and inner edges at least one of which is circular and circumscribed about a centerline and circumferentially spaced apart clockwise and counter-clockwise radial edges having an origin on the centerline. A gas turbine engine combustor includes conical-flat heat shields in one or more circular rows arranged in a non-symmetrical or asymmetrical pattern. Two or more groups (A, B, C) of the conical-flat heat shields in the circular rows may be mounted on a domeplate and one or more of the conical-flat heat shields is different in one or more of the groups (A, B, C). | 12-31-2015 |
| 20160025345 | LINER ELEMENT FOR A COMBUSTOR - There is disclosed a liner element in the form of an impingement/effusion tile for a gas turbine combustor having a structural wall. The liner element has a unitary construction defining a cooling side and combustion side, and a plurality of effusion holes extending between a cooling side surface of the element and a combustion side surface of the element. The liner element is configured to be affixed to the structural wall of a combustor with its cooling side surface spaced from the structural wall to define a chamber between the cooling side surface and the structural wall, and the liner element further includes integrally formed and internally threaded protuberances on its cooling side, the protuberances being arranged to engage the structural wall. | 01-28-2016 |
| 20160061448 | HEAT SHIELD LABYRINTH SEAL - A seal for sealing a combustor heat shield against an interior surface of a combustor shell, the seal comprising: a first sealing surface on the interior surface of the combustor shell; and a second sealing surface on a rail on an edge of a heat shield,
| 03-03-2016 |
| 20160084167 | SELF-MODULATED COOLING ON TURBINE COMPONENTS - Systems and methods are disclosed herein for passively managing cooling air in a gas turbine engine. A cooling air supply line may supply cooling air to a component in the gas turbine engine. A metering coupon may have a negative coefficient of thermal expansion. The metering coupon may allow more airflow through the metering coupon and through the component in response to an increase in temperature. | 03-24-2016 |
| 20160102860 | LINER ELEMENT FOR A COMBUSTOR, AND A RELATED METHOD - A liner element for a gas turbine combustor having a structural wall with fixing apertures provided therethrough. The liner element has a unitary construction defining a cooling side and combustion side, and a plurality of effusion holes extending between a cooling side surface of the element and a combustion side surface of the element. The liner element is configured to be affixed to the structural wall of a combustor with its cooling side surface spaced from the wall to define a chamber between the cooling side surface and the wall, and the liner element further includes integrally formed and non-threaded protuberances on its cooling side, the protuberances being arranged to engage and extend through respective fixing apertures in the combustor wall. Also disclosed is a method of thermally insulating the liner element. | 04-14-2016 |
| 20160169512 | COOLED WALL ASSEMBLY FOR A COMBUSTOR AND METHOD OF DESIGN | 06-16-2016 |
| 20160251968 | GAS TURBINE ENGINE AIRFOIL COOLING CONFIGURATION WITH PRESSURE GRADIENT SEPARATORS | 09-01-2016 |
| 20170234537 | Surface Contouring | 08-17-2017 |
