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Combustor liner

Subclass of:

060 - Power plants

060390010 - COMBUSTION PRODUCTS USED AS MOTIVE FLUID

060722000 - Combustion products generator

Patent class list (only not empty are listed)

Deeper subclasses:

Class / Patent application numberDescriptionNumber of patent applications / Date published
060754000 Porous 43
060755000 Having means to direct flow along inner surface of liner 35
060753000 Ceramic 18
060760000 Air outside liner flows counter to combustion products flow within liner 10
060758000 Air introduced within liner counter to flow of combustion products 3
20120111014Low calorific fuel combustor for gas turbine - A low calorific value fuel-fired can combustor for a gas turbine include a generally cylindrical housing, and a generally cylindrical liner disposed coaxially within the housing to define with the housing a radial outer flow passage for combustion air, the liner also defining inner combustion and a dilution zone, the dilution zone being axially distant a closed housing end relative to the combustion zone. A nozzle assembly disposed at the closed housing end includes an air blast nozzle and surrounding swirl vanes. An impingement cooling sleeve coaxially disposed in the combustion air passage between the housing and the liner impingement cools the portion of the liner defining the combustion zone. The combustion liner has an L/D ratio of in the range 1≦L/D≦4, and a ratio of the combustion zone volume (m05-10-2012
20110203287COMBUSTOR LINER FOR A TURBINE ENGINE - A combustor liner for a combustor of a turbine engine includes a single-walled generally cylindrical liner that extends from a head end to an aft end. A venturi is formed on the combustor liner at a location between the head and aft ends. The venturi is formed by first and second straight portions of the combustor liner that angle inward to form a reduced diameter portion. The combustor liner may also include an air deflector to deflect air down onto the exterior of the reduced diameter portion of the single-walled combustor liner. The combustor liner may also include dilution holes, and turbulator rings.08-25-2011
20100257864REVERSE FLOW CERAMIC MATRIX COMPOSITE COMBUSTOR - A gas turbine engine has an annular reverse-flow combustor with a combustor inner liner enclosing a combustion chamber. The inner liner having a dome portion at an upstream end of the combustor and a downstream combustor exit defined between a small exit duct portion and a large exit duct portion. At least one of the dome portion, the small exit duct portion and the large exit duct portion is made of a separately formed hemi-toroidal shell composed of a ceramic matrix composite.10-14-2010
060759000 Air scoop extends into air flowing outside liner 1
20120247112TURBINE COMBUSTION SYSTEM COOLING SCOOP10-04-2012
Entries
DocumentTitleDate
20080264064Floatwall dilution hole cooling - A combustor for a gas turbine engine is provided, the combustor having an outer shell with an outer surface exposed to cooling air and an inner surface, and at least one floatwall panel attached to the inner surface of the outer shell and having a trailing edge. At least one dilution hole is in the floatwall panel near the trailing edge and in communication with the outer surface of the outer shell, and at least one local air impingement hole is in the outer shell downstream of each at least one dilution hole, that directs the cooling air towards the trailing edge of the floatwall panel.10-30-2008
20100011774METHOD OF REFURBISHING A SEAL LAND ON A TURBOMACHINE TRANSITION PIECE AND A REFURBISHED TRANSITION PIECE - A method of refurbishing a seal land on a transition piece of a turbomachine includes applying a wear strip to a wall surface of the seal land, and covering the wear strip with a slot protector.01-21-2010
20080256955COMBUSTOR LINER WITH IMPROVED HEAT SHIELD RETENTION - A combustor liner including a dome section having a positioning hole defined at a first radial distance and sized to receive a first heat shield fastener to at least substantially prevent radial and circumferential motion of the first fastener, a circumferential slot sized to receive a second heat shield fastener to at least substantially prevent radial motion of the second fastener while allowing limited circumferential motion of the second fastener, and a clearance hole defined at a second radial distance and sized to receive a third heat shield fastener to allow limited radial and circumferential motion of the third fastener.10-23-2008
20130081399CASING FOR A GAS TURBINE ENGINE - A casing for a can annular gas turbine engine, including: a compressed air section (04-04-2013
20130081398GAS PATH LINER FOR A GAS TURBINE ENGINE - An exhaust nozzle for a gas turbine engine includes a gas path liner connected to an interior surface of an exhaust nozzle flap. The gas path liner includes a liner backbone that extends along a liner axis, and a plurality of liner panels sequentially connected to the liner backbone along the axis. Each liner panel includes a panel length that extends axially from a first panel end to a second panel end, where the panel length thermally expands independent of the liner backbone04-04-2013
20130031908ASSEMBLIES AND APPARATUS RELATED TO INTEGRATING LATE LEAN INJECTION INTO COMBUSTION TURBINE ENGINES - A transfer tube for use in a late lean injection system of a combustor, wherein the combustor includes an inner radial wall, which defines a primary combustion chamber downstream of a primary fuel nozzle, and an outer radial wall, which surrounds the inner radial wall forming a flow annulus therebetween, the outer radial wall including a late lean nozzle, the transfer tube including flow directing structure that defines a fluid passageway. At a first end, the flow directing structure may include an inlet and attachment means that attach the transfer tube to the late lean nozzle. The flow directing structure may have a configuration such that the fluid passageway spans the flow annulus and positions the outlet at a desirable injection point in the inner radial wall.02-07-2013
20100162715METHOD AND SYSTEM FOR ENHANCING HEAT TRANSFER OF TURBINE ENGINE COMPONENTS - A method and system for enhancing the heat transfer of turbine engine components is disclosed that includes applying a metallic coating having a high thermal conductivity to the cold side of a turbine component to enhance heat transfer away from the component. The metallic coating may be roughened to improve heat transfer. The metal coating may be a Ni—Al bond coating having an aluminum content greater than about 50 weight percent.07-01-2010
20100043441METHOD AND APPARATUS FOR ASSEMBLING GAS TURBINE ENGINES - A method for assembling a gas turbine engine includes coupling a transition piece between a combustor liner and a nozzle assembly. The method also includes extending a first portion of a flow sleeve from the transition piece about at least a portion of the combustor liner. The method further includes coupling a second portion of the flow sleeve to the first portion of the flow sleeve such that the flow sleeve second portion extends from the flow sleeve first portion and at least partially about at least a portion of the transition piece. The flow sleeve second portion includes a scoop that cooperates with the transition piece to at least partially define a unitary cooling air passage that includes a unitary scoop-shaped opening. The scoop is oriented to introduce a substantially uniform cooling air flow to the transition piece.02-25-2010
20130031909COMBUSTOR HEAT SHIELD WITH INTEGRATED LOUVER AND METHOD OF MANUFACTURING THE SAME - A combustor dome heat shield and a louver are separately metal injection molded and then fused together to form a one-piece combustor heat shield.02-07-2013
20130133330DEVICE TO LOWER NOx IN A GAS TURBINE ENGINE COMBUSTION SYSTEM - An emissions control system for a gas turbine engine including a flow-directing structure (05-30-2013
20100107645Combustor liner cooling flow disseminator and related method - A combustor component includes a hollow cylindrical body, at least a section of which is provided with a plurality of annular, axially spaced shoulders. A plurality of rings are provided on an interior side of the section of the cylindrical body, aligned with the shoulders to thereby create a like plurality of annular slots. A plurality of cooling holes are formed in section of the cylindrical body, radially overlying the rings, and adapted to supply cooling air to the annular slots. A plurality of flow disseminators are provided on a radially outer side of the rings, aligned with the cooling holes, and configured to spread the cooling air flowing through the cooling holes.05-06-2010
20100095677PILOT NOZZLE HEAT SHIELD HAVING INTERNAL TURBULATORS - A pilot nozzle heat shield includes a body having a first end for receiving a pilot nozzle and a second end including a flow tip. The body includes a plurality of internal turbulators circumferentially disposed about the internal peripheral surface of the body. The flow tip includes a proximal periphery and a distal periphery. A plurality of flow ports are circumferentially spaced about the proximal periphery of the flow tip. The flow tip includes a plurality of slots. Each slot extends distally from one of the flow ports to the distal periphery of the flow tip, which defines an aperture. The plurality of slots define a plurality of tangs; each tang is defined between a pair of neighboring slots. A plurality of turbulators can be disposed about the inner peripheral surface of the heat shield body at the tangs.04-22-2010
20090120093Turbulated aft-end liner assembly and 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.05-14-2009
20120180492APPARATUS FOR VIBRATION SUPPORT IN COMBUSTORS AND METHOD FOR FORMING APPARATUS - A sleeve component assembly for a combustor, and a method for forming the sleeve component assembly for the combustor, are disclosed. The sleeve component assembly includes a sleeve component, the sleeve component comprising one of an inner sleeve component or an outer sleeve component. The sleeve component assembly further includes at least one support feature extending from the sleeve component, the at least one support feature configured to contact and provide vibratory support to an adjacent sleeve component. The at least one support feature is integral with the sleeve component.07-19-2012
20090044540COMBUSTION LINER STOP IN A GAS TURBINE - A liner stop for retaining a liner within a sleeve of a combustion system, includes: a female component including a saddle, the female component adapted for coupling to one of the liner and the sleeve; a male component including a tab for insertion into the saddle, the male component adapted for coupling to an opposite one of the liner and the sleeve; and an insert adapted for being attached to the tab and ensuring dampening of vibration when the tab is disposed within the saddle. A method and a system are also disclosed.02-19-2009
20100005804COMBUSTOR STRUCTURE - A combustor includes at least one combustor liner defining a combustion chamber capable of directing combustion products toward a turbine. At least one combustor sleeve is located outside of the combustion chamber and is capable of reducing a magnitude of acoustic waves in the combustion chamber. The at least one combustor liner and the at least one combustor sleeve define at least one flow channel therebetween. Further, a combustor includes at least one combustor liner defining a combustion chamber capable of directing combustion products toward turbomachinery. At least one combustor sleeve disposed outside of the combustion chamber and is capable of controlling distribution of fluid flow in the combustor to modify a uniformity of the fluid flow to the combustion chamber.01-14-2010
20130047621MULTI-PIERCED COMBUSTION CHAMBER WITH COUNTER-ROTATING TANGENTIAL FLOWS - A combustion chamber, for example for a turbine engine, presenting an inner annular wall, an outer annular wall, and a pierced annular chamber end wall extending around an axis, the chamber end wall including at least one opening receiving a fuel injector, the opening being substantially centered on a circular line defining a first chamber end wall portion extending radially between the circular line and the inner annular wall, and a second chamber end wall portion extending radially between the circular line and the outer annular wall. First and second channels are inclined relative to a normal vector normal to the chamber end wall while extending tangentially, the first channels arranged to provide a flow of air in a first rotary direction, the second channels arranged to provide a flow of air in a second rotary direction opposite to the first rotary direction.02-28-2013
20090308077HOLE PATTERN FOR GAS TURBINE COMBUSTOR - A combustor for a turbine engine includes an outer liner having a row of circumferentially distributed outer combustion air holes and an inner liner circumscribed by the outer liner and having a row of circumferentially distributed inner combustion air holes. The inner and outer liners each include at least a major air hole having a first hole size, an intermediate air hole having a second hole size, and a minor air hole having a third hole size. The first, second, and third hole sizes are all different from each other.12-17-2009
20100269510Machine Component and Gas Turbine - A machine component including a base body produced from a base material is provided. The base body is equipped in a partial region of the surface thereof with a plating made of an application material having a greater hardness and/or viscosity as compared to the base material. The plating is foamed by a number of plating elements that are applied to the base body in the longitudinal direction thereof in a tilted manner relative to the main flow direction of a hot gas flowing through the base body.10-28-2010
20090094986Combustion Chamber Wall, Gas Turbine Installation and Process for Starting or Shutting Down a Gas Turbine Installation - In a combustion chamber wall for a combustion chamber having a combustion chamber outlet through which a hot combustion exhaust gas can exit the combustion chamber, the combustion chamber wall comprises an outlet end which surrounds the combustion chamber outlet, and the outlet end is provided with a tempering device.04-16-2009
20090301093SLIDEABLE LINER ANCHORING ASSEMBLY - An example gas turbine engine includes an engine casing and an engine liner within the engine casing. One of the engine casing or the engine liner includes a first attachment structure. The other of the engine casing or the engine liner defines a track guide. A slideable member is moveable within the track guide between an engaged position and a disengaged position. The slideable member includes a second attachment structure engageable with the first attachment structure to secure the engine liner relative the engine casing when the slideable member is in the engaged position.12-10-2009
20090019856CHAMBER-BOTTOM BAFFLE, COMBUSTION CHAMBER COMPRISING SAME AND GAS TURBINE ENGINE FITTED THEREWITH - The present invention relates to a baffle (01-22-2009
20090013694COMBUSTION CHAMBER COMPRISING CHAMBER END WALL HEAT SHIELDING DEFLECTORS AND GAS TURBINE ENGINE EQUIPPED THEREWITH - The present invention relates to an annular combustion chamber for a gas turbine engine comprising an external wall, an internal wall, a wall connecting these two walls and constituting a chamber end wall, the chamber end wall being provided with openings for the fuel injection systems, heat shielding deflectors being fixed to the wall, the deflectors comprising a flat wall portion with an opening centered on said openings for the fuel injection systems, two longitudinal edges and two transverse edges, wherein, at least along one of the longitudinal edges, a deflector comprises a tongue forming a joint cover, creating a housing along said edge for the edge of the adjacent deflector so as to seal the join between the two adjacent edges of the two deflectors, said tongue being spaced away from the chamber end wall so as to create a space supplied with cooling air via orifices in the chamber end wall.01-15-2009
20100095679DUAL 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 via a plurality of threaded studs and includes a plurality of forward heat shield panels and a plurality of aft heat shield panels. Each of the plurality of forward heat shield panels and aft heat shield panels includes a plurality of side rails, a forward rail, and an aft rail including a plurality of controlled openings, that when coupled to the outer wall defines a single cavity there between. A plurality of cavities being formed by the plurality of forward and aft heat shield panels.04-22-2010
20090056339Retaining Element and Heat Shield Element for a Heat Shield and Combustion Chamber Provided with a Heat Shield - A retaining element for retaining a heat shield element on a support structure comprises at least one fixing section adapted to fix the retaining element to the support structure and at least one retaining section adapted to engage with an engaging groove present on a periphery of the heat shield element. A projection is arranged on the retaining element in such a manner that it projects in the direction of the heat shield element when retaining a heat shield element.03-05-2009
20090235666CRACK RESISTANT COMBUSTOR - A combustion chamber louver assembly includes an aft louver having a forward panel 09-24-2009
20120234012IMPINGEMENT 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, a method for designing an impingement sleeve is disclosed. The method includes determining a desired operational value for a transition piece, inputting a combustor characteristic into a processor, and utilizing the combustor characteristic in the processor to determine a cooling hole pattern for the impingement sleeve, the cooling hole pattern comprising a plurality of cooling holes, at least a portion of the plurality of cooling holes being generally longitudinally asymmetric, the cooling hole pattern providing the desired operational value.09-20-2012
20110283707System for Cooling Turbine Combustor Transition Piece - An air cooling hood offset from an exterior of the transition piece to define an air cooling passage between the air cooling hood and the exterior of the transition piece, wherein the air cooling hood comprises a plurality of air outlets disposed along the exterior of the transition piece, and the plurality of air outlets is configured to expel an airflow from the air cooling passage away from the exterior of the transition piece.11-24-2011
20110296842CONTROL OF AXIAL THRUST BY GUIDANCE OF THE AIR DRAWN OFF FROM A CENTRIFUGAL COMPRESSOR - Inner combustion chamber casing of a turbomachine, which casing is intended to be placed downstream of a centrifugal compressor, the said casing having the shape of a disc, pierced by a central circle, and comprising on its disc at least one guide vane of the drawn-off air, the said guide vane extending longitudinally over the said disc between the periphery of the disc and the central circle and spreading out axially from the disc so as to form with the downstream face of the said centrifugal compressor a guide channel for the air which is drawn off upon exit from the said compressor,12-08-2011
20090094985Non-Rectangular Resonator Devices Providing Enhanced Liner Cooling for Combustion Chamber - Embodiments of the present invention provide resonators (04-16-2009
20100083665Open-cooled component for a gas turbine, combustion chamber, and gas turbine - An open-cooled component for a gas turbine is provided. The component includes an outer wall that is subjected to a hot gas and at least partially defines a first cavity for a first means. The outer wall is provided with through-openings which open up into the cavity on one side and into a hot gas chamber on the other. The inventive component also includes at least one second cavity for admixing a second means, said second cavity is connected to the through-openings in a fluid-connected manner. A combustion chamber for a gas turbine, and to a gas turbine including one such component is also provided.04-08-2010
20090100838WALL ELEMENT FOR USE IN COMBUSTION APPARATUS - A tile for use as part of an inner wall of a gas turbine engine combustor, the tile being provided with deformable cooling pedestals extending into a space between the tile and the outer wall. The pedestals are deformed in contact with the outer wall of the combustor.04-23-2009
20110197590Burner inserts for a gas turbine combustion chamber and gas turbine - A burner insert for a gas turbine combustion chamber is provided. The burner insert includes a burner insert wall including a cold side and a hot side, an edge delimiting the burner insert wall. The edge includes an edge bar extending at least partially circumferentially and projecting beyond the cold side. A burner opening for inserting a burner is formed in the burner insert wall.08-18-2011
20100083664METHOD AND APPARATUS FOR ASSEMBLING GAS TURBINE ENGINE - A method for assembling a gas turbine engine is provided. The method includes providing a combustor having a combustor liner assembly defining a combustion chamber. An outer combustor liner includes a radially extending first end that defines a combustion chamber input opening. An axially extending second end of the outer combustor liner defines a combustion chamber output opening. The first end transitions into the second end to form an arcuate cross-sectional shape of the outer combustor liner. A dome assembly is coupled to the first end of the combustor liner that extends downstream from the dome assembly. A fuel nozzle is positioned within a cyclone formed on the dome assembly and configured in a radial configuration.04-08-2010
20100077762Off Center Combustor Liner - The present application provides a liner for a combustor. The combustor liner may include a mouth, one or more angled transition zones, and an off center exit.04-01-2010
20100077761IMPINGEMENT COOLED COMBUSTOR SEAL - Disclosed is a combustor seal including a seal support locatable at a first combustor component and having a plurality of through impingement holes. A wave-shaped seal located at the seal support and defining at least one seal cavity between the wave-shaped seal and the seal support. A peak of the wave-shaped seal is locatable at a second combustor component. The wave-shaped seal includes at least one through passageway located upstream of the peak capable of flowing cooling fluid therethrough into the at least one seal cavity and through the plurality of impingement holes thereby cooling the first combustor component. Further disclosed is a combustor including a combustor seal and a method for cooling a first combustor component.04-01-2010
20090090110FACETED DOME ASSEMBLIES FOR GAS TURBINE ENGINE COMBUSTORS - A dome assembly for a gas turbine engine includes a plurality of faceted segments and a plurality of openings. The plurality of faceted segments are coupled together to form a dome structure that is configured to be disposed between an inner liner and an outer liner that circumscribes the inner liner. The plurality of openings are each opening formed within a respective faceted segment, and are configured to at least partially house an atomizer therein. Each faceted segment is at least substantially flat.04-09-2009
20090277180Combustor dynamic attenuation and cooling arrangement - Disclosed is a combustor casing with an inner casing which defines a combustion chamber, an outer casing spaced apart from the inner casing for defining a passage between the inner and the outer casing, first and second effusion holes arranged in the inner casing, and dividing ribs connecting the inner and outer casings and forming at least first and second volumes for receiving part of a flow injected into the passage.11-12-2009
20090120094Impingement cooled can combustor - A can combustor includes a generally cylindrical housing having an interior, an axis, and a closed axial end. The closed axial end includes means for introducing fuel to the housing interior. A generally cylindrical combustor liner is disposed coaxially within the housing and configured to define with the housing respective radially outer passages for combustion air and for dilution air, and also respective radially inner volumes for a combustion zone and a dilution zone. The combustion zone is disposed axially adjacent the closed housing end, and the dilution zone is disposed axially distant the closed housing end. The can combustor also includes an impingement cooling sleeve coaxially disposed between the housing and the combustor liner and extending axially from the closed housing end for a substantial length of the combustion zone. The sleeve has a plurality of apertures sized and distributed to direct combustion air against the radially outer surface of the portion of the combustor liner defining the combustion zone, for impingement cooling. Essentially all of the combustion air flows through the impingement cooling apertures prior to admission to the combustion zone. A small portion of the impingement cooling air may be used for film cooling of the liner proximate the closed housing end.05-14-2009
20090282832COLD GAS DYNAMIC SPRAYING OF HIGH STRENGTH COPPER - A process for forming an article, such as a combustion chamber liner, comprises the steps of providing a mandrel formed from a material, such as an aluminum containing material, having a net shape of the article to be made, depositing a powdered metal material onto the mandrel without melting the powdered metal material, and removing the material forming the mandrel to leave a free standing monolithic article. In a preferred embodiment of the present invention, the powdered metal material comprises powdered GRCop-84. Alternatively, the powdered metal material may be GRCop-42.11-19-2009
20100275606COMBUSTOR LINER - A combustor liner for a turbine engine is disclosed herein. The combustor liner includes an inner liner surface operable to define at least part of a combustion chamber in a turbine engine. The inner liner surface extends along a portion of a chordal arc on a first side of the chordal arc. The combustor liner also includes a bearing surface operable to support a floating dome panel. At least part of the bearing surface is spaced from the chordal arc on a second side of the chordal arc opposite the first side.11-04-2010
20090255266SURFACE TREATMENTS FOR PREVENTING HYDROCARBON THERMAL DEGRADATION DEPOSITS ON ARTICLES - A method of preventing thermal hydrocarbon degradation deposits on a surface of a gas turbine component, the method includes providing the turbine component comprising the surface configured for contacting a hydrocarbon fluid, wherein the substrate comprises a material having a nominal liquid wettability sufficient to generate, with reference to an oil, a nominal contact angle, disposing a plurality of features on the substrate to form an anti-deposition surface texture, wherein the plurality of features have a size, shape, and orientation selected such that the surface has an effective wettability sufficient to generate, with reference to an oil, an effective contact angle of greater than the nominal contact angle, and the features comprise a width dimension (a), and a spacing dimension (b), wherein the features prevent the hydrocarbon fluid from penetrating into the surface texture and thereby reduce the adhesion of the thermal hydrocarbon deposits to the surface.10-15-2009
20080307793Brazed Joint Between a Metal Part and a Ceramic Part - An assembly including a metal piece, a piece made of ceramic material, and at least one intermediate connection element assembled to each of the pieces by brazing. The intermediate connection element includes a deformable sheet presenting at least two flat zones brazed to respective ones of the pieces, the two flat zones being interconnected by a deformable zone presenting at least two free undulations oriented in alternation towards the metal piece and towards the piece made of ceramic material.12-18-2008
20100205973GAS TURBINE WITH WELDED COMBUSTOR LINERS - In order to provide an annular combustion chamber (08-19-2010
201102032863D NON-AXISYMMETRIC COMBUSTOR LINER - A combustor liner with an input end and an output end includes an annular inner wall and an annular outer wall. At least one of the inner wall and outer wall is three-dimensionally contoured. The inner wall and the outer wall form a combustion chamber with the contours creating alternating expanding and constricting regions inside the chamber causing combustion gases to flow in the circumferential and axial directions.08-25-2011
20100139283COMBUSTOR LINER WITH INTEGRATED ANTI-ROTATION AND REMOVAL FEATURE - A combustor liner for a gas turbine engine including first and second annular liner portions engaged to one another with an interference fit. The first liner portion includes at least one anti-rotation feature extending therefrom and engaging the second liner portion and having a threaded hole defined therethrough at least substantially perpendicularly to an adjacent radially extending surface of the second liner portion for threadingly receiving a pushing tool for pushing against that surface.06-10-2010
20090199564DEVICE FOR MOUNTING AN IGNITER PLUG IN A COMBUSTION CHAMBER OF A GAS TURBINE ENGINE - The present invention relates to a device for mounting an igniter plug in a combustion chamber of a gas turbine engine contained in a casing, in which the combustion chamber has an axis YY, the mounting device comprising a hollow shaft, of axis XX, a floating igniter plug sleeve absorbing expansion along an axis perpendicular to the axis XX of the hollow shaft. The device is one which further comprises a hollow shaft sleeve such that the igniter plug sleeve is housed in the hollow shaft sleeve, and means of inclining said hollow shaft sleeve relative to the axis XX. Inclining the hollow shaft sleeve allows the chamber to be inclined relative to the axis XX.08-13-2009
20100005803COMBUSTION LINER FOR A GAS TURBINE ENGINE - A combustion duct assembly has a transition duct and a combustion liner having a hula seal at a downstream end that is forced within an inner wall of the transition duct. The combustion liner is held within the transition duct by the hula seal, but allowed to move relative to the transition duct. The combustion liner is formed with heat transfer columns adjacent the downstream end, and radially inwardly of the hula seal.01-14-2010
20100162716PANELED COMBUSTION LINER - The present invention relates to combustors, and provides a combustion liner that may at least partially contain a combustion process. The combustion liner may include a support structure that supports a plurality of panels. The panels may be in the form of thermal barrier panels, e.g., that are attached to the support structure and/or other parts of the machine.07-01-2010
20130213047COMBUSTION LINER GUIDE STOP AND METHOD FOR ASSEMBLING A COMBUSTOR - A combustor for a gas turbine including a casing, a flow sleeve at least partially disposed within the casing, a combustion liner at least partially disposed within the flow sleeve, a liner stop feature extending from the combustion liner, and a liner guide stop including a first end separated from a second end, the second end configured to be at least partially engaged with the liner stop feature, wherein the liner guide stop extends through the casing and the flow sleeve.08-22-2013
20120102959SUBSTRATE 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.05-03-2012
20100192587COMBUSTOR ASSEMBLY FOR USE IN A GAS TURBINE ENGINE AND METHOD OF ASSEMBLING SAME - A combustor assembly for use in a gas turbine engine and method of assembly is described. The combustor assembly includes a combustor liner having a slot that at least partially circumscribes the combustor liner. The slot is defined adjacent to a venturi throat region defined within the liner. The combustor assembly also includes a restrictor plate having at least one aperture defined therein. The restrictor plate is removably coupled within the combustor assembly such that the restrictor plate is inserted within the slot and extends at least partially across the venturi throat region.08-05-2010
20120186260TRANSITION PIECE IMPINGEMENT SLEEVE FOR A GAS TURBINE - An impingement sleeve for a transition piece of a gas turbine is disclosed. The impingement sleeve generally includes a first casing configured to surround a portion of an inner duct of the transition piece and a second casing configured to surround a portion of the inner duct. Additionally, the impingement sleeve may include a joint defined between the first and second casings. The joint may include a plurality of fasteners configured to attach the first casing to the second casing.07-26-2012
20090077974Heat Shield Arrangement for a Component Guiding a Hot Gas in Particular for a Combustion Chamber in a Gas Turbine - The invention relates to a heat shield arrangement for a hot gas (m)-guiding component, which comprises a number of heat shield elements arranged side-by-side on a supporting structure while leaving a gap there between. A heat shield element can be mounted on the supporting structure whereby forming an interior space which is delimited in areas by a hot gas wall to be cooled, with an inlet channel for admitting a coolant into the interior space. According to the invention, a coolant discharge channel is provided for the controlled discharge of coolant from the interior space and, from the interior space, leads into the gap. Coolant can be saved and efficiently used by the specific coolant discharge via the coolant discharge channel, and reduction in pollutant emissions can also be achieved. The heat shield arrangement is particularly suited for linking a combustion chamber of a gas turbine.03-26-2009
20090077975Heat shield element for lining a combustion chamber wall, combustion chamber and gas turbine - The invention relates to a heat shield element for lining a combustion chamber wall, to a combustion chamber and to a gas turbine. The heat shield element comprises a hot side that can be exposed to a hot medium, a wall side opposite said hot side, and a peripheral side adjoining the hot side and the wall side and having a peripheral side surface. Relief slots are introduced into the material in an area of the heat shield element that is susceptible to material cracks induced by thermal stress, thereby limiting crack propagation.03-26-2009
20090077976ANNULAR COMBUSTION CHAMBER FOR A GAS TURBINE ENGINE - The present invention relates to an annular gas turbine engine combustion chamber comprising an outer wall and an inner wall connected by a wall forming the chamber bottom, the walls delimiting sources of combustion with axes inclined relative to the axis of the chamber, the chamber-bottom wall, of frustoconical shape, being pierced with orifices for the fuel injection systems, the planes of the orifices being perpendicular to the axes of the sources of combustion, heat-protection baffles centered on each of the orifices comprising a shoulder by which they rest against a flat surface portion along the periphery of the orifices. The chamber is characterized in that the chamber-bottom wall is conformed in a succession of adjacent flat facets having a common edge, with one facet per orifice, the shoulder of the deflectors pressing against the plane of the facets.03-26-2009
20100192589Method for Production of a Coating System - Conventionally, cavities and cracks are filled with a solder metal which forms brittle phases with a subsequently applied coating, which have a negative effect on the mechanical properties. According to the invention, the components which form brittle phases are removed from the solder metal. The above is achieved, whereby a second material is applied which reacts with said component and which is removed again with the brittle phases, before the coating.08-05-2010
20100223930INJECTION DEVICE FOR A TURBOMACHINE - A turbomachine includes a compressor, a combustor including a first end operatively connected to the compressor and a second end, a transition piece mounted to the second end of the combustor, and at least one injection device mounted to one of the combustor and the transition piece. The at least one injection device includes a first end portion that extends to a second end portion through an intermediate portion. The intermediate portion includes a flow conditioning mechanism. Combustion air from the compressor enters the first end portion passes through the flow conditioning mechanism and into the one of the combustion liner and transition piece. The flow conditioning mechanism creates an air flow disturbance in the combustion air to promote mixing of combustion gases.09-09-2010
20100212324DUAL WALLED COMBUSTORS WITH IMPINGEMENT COOLED IGNITERS - A combustor for a gas turbine engine includes an inner liner and an outer liner circumscribing the inner liner and forming a combustion chamber with the inner liner. The outer liner is a dual walled liner with a first wall and a second wall. The combustor includes a fuel igniter comprising a tip portion configured to ignite an air and fuel mixture in the combustion chamber and an igniter tube positioning the fuel igniter relative to the combustion chamber. The igniter tube includes a plurality of holes configured to direct cooling air toward the tip portion of the fuel igniter.08-26-2010
20100212323MICRO-COMBUSTOR FOR GAS TURBINE ENGINE - An improved gas turbine combustor (08-26-2010
20100236248Combustion Liner with Mixing Hole Stub - A combustor liner for a gas turbine combustor includes a cooling hole formed in the liner that delivers cooling air into a combustion zone of the combustor. A stub is secured in the cooling hole and is structured to provide added stiffness to an inside edge of the cooling hole. The added stiffness reduces cracking caused by thermal fatigue and provides resistance against high cycle fatigue failures at high frequencies.09-23-2010
20100236249Systems and Methods for Reintroducing Gas Turbine Combustion Bypass Flow - A system and method for reintroducing gas turbine combustion bypass flow. The system may include a combustor body, wherein the combustor body includes a reaction zone for primary combustion of fuel and air, and a casing enclosing the combustor body and defining an annular passageway for carrying compressor discharge air into the combustor body at one end. The system further may include a reintroduction manifold for receiving combustor bypass air extracted from the compressor discharge air in the annular passageway, and one or more reintroduction slots in communication with the reintroduction manifold for injecting the combustor bypass air into the combustor body downstream of the reaction zone. The method may include extracting combustor bypass air from the annular passageway, transporting the combustor bypass air to a reintroduction manifold, and reintroducing the combustor bypass air into the combustor body through one or more reintroduction slots in communication with the reintroduction manifold.09-23-2010
20100242485COMBUSTOR LINER - A combustor within which a combustion zone is defined is provided and includes an annular liner having a first mixing hole defined therein at a first axial position, a flow sleeve, having a second mixing hole defined therein at a second axial position, the flow sleeve surrounding the liner to form a first flow space at an exterior of the liner, a port, coupled to the flow sleeve at the second axial position, which is configured to remove air from the first flow space via the second mixing hole, and a shield, having a third mixing hole defined therein at the second axial position, the shield being disposed to shield the liner and to form a second flow space within the liner, which is communicable with the combustion zone via the third mixing hole and with the first flow space via the first mixing hole.09-30-2010
20100126177Augmentor Pilot - A gas turbine engine augmenter has a gas flowpath. A number of vanes extend into the gas flowpath. A number of augmenter fuel conduits have outlets along at least some of the vanes. At least one burner discharge outlet is along at least one of the vanes for discharging a pilot gas.05-27-2010
20100300106SYSTEM AND METHOD FOR THERMAL CONTROL IN A CAP OF A GAS TURBINE COMBUSTOR - A system comprises a turbine combustor cap. The turbine combustor cap includes a plurality of segments. Each segment of the plurality of segments has edges abutting at least two fuel nozzle receptacles. Moreover, each segment of the plurality of segments does not completely surrounding any one fuel nozzle receptacle.12-02-2010
20110126544AIR FLOW PASSAGE LINER - A liner is provided for an air flow passage of a gas turbine engine. The liner comprises an acoustic attenuation layer which forms an air-washed surface of the flow passage, and a heat exchanger layer which extends as a backing to the acoustic attenuation layer. The heat exchanger layer is configured to carry a heated fluid flow and to transport heat from the fluid flow to the acoustic attenuation layer from where the heat is transferred to air flowing through the air flow passage.06-02-2011
20110016869COOLING STRUCTURE FOR GAS TURBINE COMBUSTOR - To provide a gas turbine combustor having an improved cooling structure effective to efficiently suppress a possible occurrence of buckling in the combustion liner while exhibiting a convection cooling effect to the combustion liner, the gas turbine combustor includes a combustion liner having a combustion chamber defined therein and an outer peripheral surface forming a path of a compressed air, and a heat transfer enhancement structure provided on the outer peripheral surface of the combustion liner. The heat transfer enhancement structure referred to above is of a honeycomb construction defined by ribs protruding outwardly from the outer surface of the combustion liner. The honeycomb construction may be of a geometry in which hexagonal shapes, rhombic shapes, parallelogrammic shapes, bent rectangular shapes or triangular shapes are deployed next to each other.01-27-2011
20110030377COMBUSTOR - A combustor includes a first wall, a second wall, an injector grommet, a combustor longitudinal axis, and a connection assembly indirectly connecting the first wall to the second wall. A portion of the injector grommet is disposed within a groove of the connection assembly.02-10-2011
20100218502EFFUSION COOLED ONE-PIECE CAN COMBUSTOR - A combustor for an industrial turbine includes a single transition piece transitioning directly from a combustor head-end to a turbine inlet. The transition piece includes an inner surface and an outer surface. The inner surface bounds an interior space for combusted gas flow from the combustor head-end to the turbine inlet. The outer surface at least partially defines an area for compressor discharge air flow. The transition piece includes a plurality of apertures configured to allow compressor discharge air flow into the interior space. Each of the plurality of apertures extends from an entry portion on the outer surface to an exit portion on the inner surface.09-02-2010
20110239654ANGLED SEAL COOLING SYSTEM - An aft liner seal for a combustor for a gas turbine includes an inner shell having an inner and outer surface and central axis. The aft liner seal has an outer shell positioned over the inner shell that has an inner and outer surface and central axis coaxial with the inner shell. One of the outer surface of the inner shell or the inner surface of the outer shell has grooves angled relative to the central axis. The angled grooves and adjacent portions of the inner surface of the outer shell or the outer surface of the inner shell form cooling passages. The cooling air exits the cooling passages at an exit angle that is matched to a swirl angle of the combustor flow. Matching of the exit angle and the swirl angle minimizes shear of cooling air with respect to flow exiting the combustor liner.10-06-2011
20110061393Gas Turbine Transition Duct Profile - A transition duct having a panel assembly with an inlet end of generally circular cross section and an outlet end having a generally rectangular arc-like cross section is disclosed. The panel assembly has an uncoated internal profile substantially in accordance with coordinate values X, Y, and Z as set forth in Table 1. The coordinates are taken at a sweep angle Θ wherein Θ is an angle measured from the inlet end and X, Y, and Z are coordinates define the panel assembly profile at each angle Θ. An alternate embodiment of the invention defines an envelope for the uncoated internal profile of the panel assembly.03-17-2011
20100037617TRANSITION WITH A LINEAR FLOW PATH WITH EXHAUST MOUTHS FOR USE IN A GAS TURBINE ENGINE - A transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine is disclosed. The transition duct may have an internal passage extending between an inlet to an outlet. An axis of the transition duct body may be generally linear such that gases expelled from the transition duct body flow in a proper direction into the downstream turbine blades. The linear transition duct may include an outlet with exhaust mouths that are configured such that sides of the transition duct are coplanar with adjacent transition ducts, thereby eliminating destructive turbulence between adjacent, linear transition ducts.02-18-2010
20110247339COMBUSTOR HAVING A FLOW SLEEVE - A combustor is provided and includes a liner through which fluid fed from at least two injection points flows from a head end to an interior of a transition piece, a first one of the at least two injection points being axially proximate to a fluid impenetrable coupling between the liner and the transition piece and defining apertures disposed in fluid communication with a first passage leading to the head end, and a second one of the at least two injection points being disposed axially between the apertures and the head end and upstream from the apertures relative to a direction of fluid flow through the first passage, the second one of the at least two injection points being formed of openings disposed in fluid communication with a second passage leading to the head end.10-13-2011
20100037618TRANSITION WITH A LINEAR FLOW PATH FOR USE IN A GAS TURBINE ENGINE - A transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine is disclosed. The transition duct may have an internal passage extending between an inlet to an outlet. An axis of the transition duct body may be generally linear such that gases expelled from the transition duct body flow in a proper direction into the downstream turbine blades.02-18-2010
20100037620Impingement and effusion cooled combustor component - A cooling arrangement for cooling a first turbine combustor component surrounded by a second component includes a first plurality of impingement cooling holes in the second component, the impingement cooling holes directing cooling air onto designated areas of the first turbine combustor component; and a second plurality of effusion cooling holes in the first turbine combustor component located to cool by effusion other areas of the first turbine combustor component.02-18-2010
20100037619CANTED OUTLET FOR TRANSITION IN A GAS TURBINE ENGINE - A transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine has an internal passage from an inlet to an outlet. The outlet may include canted sides that reduce formation of damaging vibration in downstream turbine blades caused by downstream wake between adjacent transition ducts and by pressure differentials between adjacent transition ducts that include turning sections.02-18-2010
20100071379EFFUSION COOLING TECHNIQUES FOR COMBUSTORS IN ENGINE ASSEMBLIES - A combustor for an engine assembly includes a cylindrical wall forming a combustion chamber in which an air and fuel mixture is combusted; and a plurality of effusion cooling holes formed in the cylindrical wall, the plurality of effusion cooling holes oriented such that cooling air flowing therethrough cools the cylindrical wall with effusion cooling, convection cooling, and impingement cooling.03-25-2010
20120240586LOW EMISSION AND FLASHBACK RESISTANT BURNER TUBE AND APPARATUS - A burner tube to provide combustible materials to a combustor is provided and includes an annular shroud and a center body, having a cavity defined therein, disposed within the annular shroud to form an annular passage, the annular passage being communicable with a combustion zone of the combustor at an aft portion thereof and including a fore portion in which fuel is injected into the annular passage. The center body includes a surface having a passage defined therein through which air is to be supplied to the annular passage from the cavity at a position, which is downstream from the fuel injection and upstream from the combustion zone. Also provided is a contouring of the centerbody.09-27-2012
20110173984GAS TURBINE TRANSITION PIECE AIR BYPASS BAND ASSEMBLY - An air bypass band assembly includes a transition piece of a gas turbine, the transition piece having at least one opening therein to allow a flow of air to pass through the at least one opening. The air bypass band assembly also includes a band that is movable between at least two positions, a first one of the at least two positions being a closed position where the at least one opening is closed to prevent the flow of air from flowing through the at least one opening, a second one of the at least two positions being an open where the at least one opening is opened to allow the flow of air to flow through the at least one opening. The air bypass band assembly further includes a mechanism that moves the band between the at least two positions.07-21-2011
20110061394METHOD OF HEAT TREATING A NI-BASED SUPERALLOY ARTICLE AND ARTICLE MADE THEREBY - A method of heat treating an Ni-base superalloy article is disclosed. The method includes hot-working an article comprising an NiCrMoNbTi superalloy comprising, in weight percent, at least about 55 Ni to produce a hot-worked microstructure; solution treating the article at a temperature of about 1600° F. to about 1750° F. for about 1 to about 12 hours to form a partially recrystallized warm-worked microstructure; and cooling the article. The method also includes precipitation aging the article at a first precipitation aging temperature of about 1300° F. to about 1400° F. for a first duration of about 4 hours to about 12 hours; cooling the article to a second precipitation aging temperature; precipitation aging the article at a second precipitation aging temperature of about 1150° F. to about 1200° F. for a second duration of about 4 hours to about 12 hours; and cooling the article from the second precipitation aging temperature to an ambient temperature.03-17-2011
20110067405Integrated Ion Transport Membrane and Combustion Turbine System - Integrated gas turbine combustion engine and ion transport membrane system comprising a gas turbine combustion engine including a compressor with a compressed oxygen-containing gas outlet; a combustor comprising an outer shell, a combustion zone in flow communication with the compressed oxygen-containing gas outlet, and a dilution zone in flow communication with the combustion zone and having one or more dilution gas inlets; and a gas expander. The system includes an ion transport membrane oxygen recovery system with an ion transport membrane module that includes a feed zone, a permeate zone, a feed inlet to the feed zone in flow communication with the compressed oxygen-containing gas outlet of the compressor, a feed zone outlet, and a permeate withdrawal outlet from the permeate zone. The feed zone outlet of the membrane module is in flow communication with any of the one or more dilution gas inlets of the combustor dilution zone.03-24-2011
20110247340APPARATUS AND METHOD FOR MINIMIZING AND/OR ELIMINATING DILUTION AIR LEAKAGE IN A COMBUSTION LINER ASSEMBLY - A combustion liner assembly for a gas turbine includes an outer liner, the outer liner having a flange at a forward end. An inner liner is disposed within the outer liner. The inner liner has a first inner wall. A venturi includes a second inner wall, a venturi throat, and the first inner wall of the inner liner. A slip joint is connected to the second inner wall. The slip joint receives the flange of the outer liner. Alternatively, or additionally, the combustion liner assembly includes a slip joint between the inner or outer liner and an aft section.10-13-2011
20100037621Thermal Machine - A thermal machine, especially a gas turbine, includes an annular combustor which is outwardly delimited by an outer shell and an inner shell (02-18-2010
20090000303COMBUSTOR HEAT SHIELD WITH INTEGRATED LOUVER AND METHOD OF MANUFACTURING THE SAME - A combustor dome heat shield and a louver are separately metal injection molded and then fused together to form a one-piece combustor heat shield.01-01-2009
20110162378TUNABLE TRANSITION PIECE AFT FRAME - A tunable transition piece aft-frame is disclosed for tuning the exit profile of combustion products as such products flow from a transition piece to a combustion product receiving apparatus. The aft-frame includes a generally rectilinear shaped body which has a laterally extending flange. Dilution holes are formed in the flange and are configured to allow dilution air to penetrate into the flow of the combustion products to shape the exit temperature profile of the combustion products.07-07-2011
20110120133DUAL WALLED COMBUSTORS WITH IMPROVED LINER SEALS - A combustor for a turbine engine is provided. The combustor includes a first liner and a second liner forming a combustion chamber. The combustion chamber is configured to receive an air-fuel mixture for combustion therein and having a longitudinal axis that defines axial and radial 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 first liner cavity having first and second ends. A first liner seal is configured to seal the second end of the first liner cavity and to accommodate relative movement of the first hot wall and first cold wall generally in the axial and radial directions.05-26-2011
20090293488Combustor - A gas turbine engine combustor has forward bulkhead extending between inboard and outboard walls and cooperating therewith to define a combustor interior volume or combustion chamber. At least one of the walls has an exterior shell and an interior shell including a number of panels. Each panel has interior and exterior surfaces and a perimeter having leading and trailing edges and first and second lateral edges. A number of cooling passageways have inlets on the panel exterior surface and outlets on the panel interior surface. The shell has a plurality of holes for directing air to a space between the shell and heat shield and adapted for preferentially directing said air toward leading edge portions of first stage vanes of a turbine section.12-03-2009
20090293487TURBOMACHINE COMBUSTION CHAMBER - Annular combustion chamber for a turbomachine, comprising two cylindrical walls these being a radially internal and a radially external wall and fixed by bolting at their upstream ends (12-03-2009
20090293486COMBUSTORS WITH IGNITERS HAVING PROTRUSIONS - A combustor for a gas turbine engine is provided, and includes an inner case; an outer case circumscribing the inner case and forming an annular pressure vessel therebetween; an inner liner positioned within the annular pressure vessel; an outer liner circumscribing the inner liner and forming a combustion chamber with the inner liner; and an igniter coupled to the outer case and extending to the outer liner such that the igniter is positioned to ignite an air and fuel mixture in the combustion chamber. The igniter includes a protrusion for coupling the igniter to the outer liner.12-03-2009
20110185736GAS TURBINE COMBUSTOR WITH VARIABLE AIRFLOW - An annular combustor and a method for operating a gas turbine engine over a power demand range facilitate combustion in a lean direct ignition (LDI) mode over an extended range of operating fuel air ratios. The flow primary combustion air admitted into the primary combustion zone is varied in response to power demand from a maximum air flow rate of high power demand to a minimum flow air rate of low power demand, while the flow of dilution air into a quench zone downstream of the primary combustion zone is increased from a minimum air flow rate at high power demand to a maximum air flow rate at low power demand.08-04-2011
20090178412APPARATUS AND METHOD FOR A GAS TURBINE ENTRAINMENT SYSTEM - This invention relates to an apparatus for an entrainment system of a vortex burning combustion chamber or a vortex burning inter-turbine burner in a gas turbine. The entrainment system rapidly and thoroughly mixes hot combustion gases with non-combustion gases to reduce the gas temperature before entering a turbine. The entrainment system includes a plurality of helical vanes forming trenches and resulting in a highly helical flow path. The highly helical flow path provides an increased residence time for mixing of the combustion gases and non-combustion gases. Radial cavities in the helical vanes, canted vane angles and varying geometries further facilitate mixing while reducing losses. This invention also includes a method of mixing combustion and non-combustion gases in an entrainment system.07-16-2009
20110113785THERMAL MACHINE - A thermal machine is provided, in particular a gas turbine, which includes an annular combustion chamber which is bounded on the outside by an outer shell and an inner shell. The outer shell and the inner shell are each split on a separating plane into an upper half and a lower half, which are mechanically interlocked by welding one to the other on the separating plane. Increased mechanical robustness and a longer life of the combustion chamber are achieved in that an additional mechanical interlock is provided on the separating planes in order to absorb tensile and shear forces acting on the separating planes.05-19-2011
20120036857COMBUSTION LINER STOP BLOCKS HAVING INSERTABLE WEAR FEATURES AND RELATED METHODS - A combustion liner stop block having insertable wear features and methods for installing wear features within a combustion liner stop block are disclosed. The methods may generally include enlarging a slot defined in the liner stop block to form a cavity, locating an insert block in the cavity and providing at least one wear feature associated with the insert block.02-16-2012
20120304656COMBUSTION LINER AND TRANSITION PIECE - An apparatus is disclosed. The apparatus may include a body configured to flow hot gases of combustion between a forward end and an aft end. Additionally, the body may include a plurality of raised sections spaced apart circumferentially around an outer perimeter of the body. The raised sections may generally extend lengthwise between the forward and aft ends.12-06-2012
20120304657LOCK LEAF HULA SEAL - A flexible annular seal for insertion between concentrically assembled turbine combustor components includes an annular inner seal portion having a first solid annular edge and first plurality of spring fingers extending axially from the first solid annular edge; and an annular outer seal portion having a second solid annular edge and a second plurality of spring fingers extending axially from the second solid edge and overlying the first plurality of spring fingers such that the inner and outer seal portions are substantially fully engaged along an entire length dimension of the flexible annular seal. The second plurality of spring fingers are circumferentially offset from the first plurality of spring fingers, and free ends of the first plurality of spring fingers are bent around and over free ends of the second plurality of spring fingers.12-06-2012
20120304658Segment component in high-temperature casting material for an annular combustion chamber, annular combustion chamber for an aircraft engine, aircraft engine and method for the manufacture of an annular combustion chamber - The present invention relates to a segment component in high-temperature casting material for an annular combustion chamber of an aircraft engine, characterized by a combustion-chamber wall which in operation shields a fuel flame extending along a burner axis from the environment, with the combustion-chamber wall having a bulge which points in a direction facing away from the burner axis. The invention furthermore relates to an annular combustion chamber, an aircraft engine with an annular combustion chamber as well as a method for the manufacture of an annular combustion chamber.12-06-2012
20120304655TURBOMACHINE COMBUSTOR ASSEMBLY INCLUDING A LINER STOP - A turbomachine combustor assembly includes a combustor housing and a combustor body. The combustor body defines a combustor liner having a first end portion that extends to a second end portion through a combustion chamber. A cap assembly is mounted at the combustor housing. The cap assembly includes an endcover, a plurality of fuel nozzles supported by the end cover, and an outer barrel member. The outer barrel member extends about the plurality of fuel nozzles. A liner stop is arranged on one of the outer barrel member and the first end portion of the combustor liner. The liner stop includes a stepped lip portion that receives the other of the outer barrel member and the first end portion of the combustor liner to form substantially smooth liner to outer barrel interface. The liner stop restricts axial movement of the combustor liner relative to the outer barrel member.12-06-2012
20110314829LINER AFT END SUPPORT MECHANISMS AND SPRING LOADED LINER STOP MECHANISMS - A gas turbine includes a liner, a casing surrounding the liner, a hula seal flexibly connected to an aft end of the liner and a liner aft support mechanism. The liner is configured to receive compressed gas and fuel at an upstream end, the mixture of the compressed gas and the fuel being burned in a combustion core area of the liner to yield hot exhaust gasses. The liner aft end support mechanism is located downstream from an area where a highest temperature on an outer surface of the liner is attained, and upstream to a portion where the hula seal is connected to the liner, and is configured to movably support the liner inside the casing. The liner aft end support mechanism includes at least three individual support elements configured to allow a part of the individual support elements to move in the flow direction relative to at least one of the liner or the casing.12-29-2011
20120006031REPLACEABLE ORIFICE FOR COMBUSTION TUNING AND RELATED METHOD - A boss and orifice plate assembly comprising an annular boss adapted to be secured in a hole formed in a combustor component, said boss formed with an annular seat supporting a replaceable orifice plate, and an annular retaining ring groove adjacent said seat, said seat extending radially inwardly of said annular retaining ring groove; and a retaining ring seated in said retaining ring groove and at least partially and resiliently engaged between a surface of said groove and a surface of said orifice plate.01-12-2012
20100205972ONE-PIECE CAN COMBUSTOR WITH HEAT TRANSFER SURFACE ENHACEMENTS - A method, system, and apparatus is provided for transferring heat from a can combustor associated with a gas turbine by providing a one-piece can combustor body having surface features that facilitate the transfer of heat away from the can combustor body, and by directing air flow to the surface features so that heat from the can combustor body is transferred from the surface features to the air flow.08-19-2010
20120060503TRANSITIONAL REGION FOR A COMBUSTION CHAMBER OF A GAS TURBINE - A gas turbine including a combustion chamber and a first row of guide vanes, arranged essentially directly downstream thereof, of a turbine. The outer and/or inner limitation of the combustion chamber defined by at least one outer and/or inner heat shield, mounted on at least one combustion chamber structure arranged radially outside and/or inside. The hot gases flow path in the region of the guide vane row being restricted radially on the outside and/or inside by an outer and/or inner vane platform, mounted at least indirectly on at least one turbine carrier. A minimal gap size directly upstream of the first row of guide vanes is achieved by mounting at least indirectly on the turbine carrier at least one mini heat shield, arranged upstream of the first row of guide vanes and essentially adjacent the vane platform and in the flow direction between the heat shield and the vane platform.03-15-2012
20100095678Heat Shield Sealing for Gas Turbine Engine Combustor - A combustor heat shield sealing arrangement comprises a sealing rail extending from the combustor liner shell at the exit of the combustor for sealing engagement with a rail-less downstream end portion of the combustor heat shield. The sealing rail is offset relative to the downstream vane passage. Doing so may minimize the combustor/vane waterfall and, thus, minimize the horseshoe vortex effect at the leading edge of the turbine vanes.04-22-2010
20090133403Internal manifold air extraction system for IGCC combustor and method - A combustor for a turbine including a combustor liner; a first flow sleeve surrounding the combustor liner with a first flow annulus therebetween, the first flow sleeve having at least one cooling aperture formed about a circumference thereof for directing compressor discharge air as cooling air into the first flow annulus; a casing surrounding first flow sleeve with a second flow annulus therebetween, the first flow sleeve having at least one air extraction opening formed about a circumference thereof for directing compressor discharge air from the first flow annulus as extraction air into the second flow annulus; and an extraction port operatively coupled to the casing for extracting the extraction air from the second flow annulus.05-28-2009
20110179798SEALING BETWEEN A COMBUSTION CHAMBER AND A TURBINE NOZZLE IN A TURBOMACHINE - A turbomachine including an annular combustion chamber, a sectorized turbine nozzle arranged at an outlet from the chamber, and a sealing mechanism interposed axially between the chamber and the nozzle, the sealing mechanism including an annular gasket that is axially resilient. The gasket includes a first axial bearing mechanism for bearing against a downstream end of the chamber and a downstream annular lip that is sectorized, each sector of the downstream lip being in alignment with a sector of the nozzle and including a second axial bearing mechanism for bearing against an upstream end of the nozzle sector.07-28-2011
20090293489COMBUSTOR LINER CAP ASSEMBLY - A liner cap assembly is disclosed for use in a gas turbine engine combustor. The assembly includes an outer ring that extends along an axis. Multiple struts are circumferentially arranged about an inner diameter of the outer ring and extend radially inwardly therefrom. A plate is supported by and axially aligned with the struts. The plate includes multiple circumferential openings that support a collar and a premix tube at each of the openings. The plate is arranged between leading and trailing edges of the struts to provide a stiffened liner cap assembly that is robust and resistant to the vibrations typically found in dry low NOx systems.12-03-2009
20120159954TRANSITION PIECE AND GAS TURBINE - A transition piece 06-28-2012
20120159955MOUNTING/DISMOUNTING JIG FOR COMBUSTOR TAIL PIPE AND TAIL PIPE INSTALLATION METHOD - Provided is a tail cylinder attaching and detaching fixture that attaches and detaches a tail cylinder to and from a casing, the tail cylinder being included in a combustor inserted into the casing so that the front end is connected to an inlet portion of a combustion gas passageway, the tail cylinder attaching and detaching fixture including: a guide portion of which the front end is disposed inside the casing and the front end and the base end are respectively supported by the casing and which supports the tail cylinder so as to be movable in the axial direction of the combustor; and an advancing and retracting mechanism that advances and retracts the tail cylinder in the axial direction.06-28-2012
20120247111TURBINE COMBUSTION SYSTEM LINER - A combustion chamber liner (10-04-2012
20100287941ADVANCED QUENCH PATTERN COMBUSTOR - A combustor for a gas turbine engine is provided. The combustor includes a forward bulkhead, an inner radial combustor wall, and an outer radial combustor wall. The bulkhead includes a plurality of circumferentially disposed injector apertures. The inner radial combustor wall includes a plurality of inner quench aperture sets. Each inner quench aperture set includes a first inner quench aperture and a second inner quench aperture separated from each other by an inner interset distance. Each inner quench aperture set is separated from an adjacent inner quench aperture set by an inner intraset distance. The inner intraset distance is different than the inner interset distance. The outer radial combustor wall includes a plurality of circumferentially disposed outer quench apertures. The outer radial combustor wall is disposed radially outside of the inner radial combustor wall, thereby defining an annular combustion region therebetween.11-18-2010
20120131922COATING A PERFORATED SURFACE - An example method of coating a surface includes rotating a sprayer about an axis and directing spray away from the axis using the sprayer. The method coats a surface with the spray. The method moves a fluid through apertures established in the surface to limit movement of spray into apertures. The apertures are configured to direct the fluid toward the axis.05-31-2012
20100192588METHOD FOR THE PROVISION OF A COOLING-AIR OPENING IN A WALL OF A GAS-TURBINE COMBUSTION CHAMBER AS WELL AS A COMBUSTION-CHAMBER WALL PRODUCED IN ACCORDANCE WITH THIS METHOD - A cooling-air opening in produced in a wall (08-05-2010
20120167573Thermal Barrier Coatings and Methods of Application - A coated part is exposed to a gas flow. The gas flow has a characteristic gas flow direction distribution over a surface of the coated part. The coated part has a substrate having a substrate surface and a coating over the substrate surface. The coating comprises at least one coating layer. A first such layer is columnar and has a column boundary direction distribution. The column boundary direction distribution is selected for partial local alignment with the gas flow direction distribution.07-05-2012
20110185737COMBUSTOR LINER SEGMENT SEAL MEMBER - A combustor liner segment seal member is provided that includes a center section, a forward flange, and an aft flange. The center section includes a base surface, a gas path surface, a forward side surface, and an aft side surface. The forward flange extends outwardly from the forward side surface, and includes a width, a height, a shell side surface, and a liner side surface. The aft flange extends outwardly from the aft side surface, and includes a width, a height, a shell side surface, and a liner side surface.08-04-2011
20100050649COMBUSTOR DEVICE AND TRANSITION DUCT ASSEMBLY - A combustor device and transition duct assembly is provided for use in a gas turbine engine. The combustor device comprises combustor structure having an exit portion; spring clips mounted to the exit portion of the combustor structure; and a burner assembly. The transition duct comprises a conduit having inlet and outlet sections and an abradable material layer provided along a circumferential portion of the inlet section of the transition duct conduit. The transition duct conduit inlet section may be coupled to the combustor structure exit portion such that the spring clips engage the abradable material layer.03-04-2010
20100031664Combustor liner replacement panels - A replacement panel for repairing a liner for a gas turbine engine combustor, the combustor having a combustion zone formed by inner and outer liners, the replacement panel comprising a sheet of material suitable for use in a combustor liner, at least one opening in the sheet of material, a thermal barrier material applied to the sheet of material adjacent the at least one opening; and a peripheral edge free of thermal barrier material.02-11-2010
20100011775Combustion apparatus - A coated turbine engine article having a first wall and a second wall, wherein the first wall has a shielding feature integral with the article and which inhibits entry of a coating material into an open passage extending between the first wall and the second wall when the coating material is directed towards the article using a line of sight process.01-21-2010
20100011773COMBUSTOR LINER AND METHOD OF FABRICATING SAME - A combustor liner includes a liner having an upstream end and a downstream end having a longitudinal axis extending therethrough, and a plurality of cooling holes formed in the liner, the cooling holes are arranged along the longitudinal axis into a plurality of circumferentially extending rows that are variably spaced apart along the longitudinal axis.01-21-2010
20120073303METAL INJECTION MOLDING PROCESS AND COMPONENTS FORMED THEREWITH - A process of producing a metallic component having a desired shape that includes at least one nonuniform section, as well as metallic components produced by such a process. The process uses a composition containing a mixture of a polymeric binder and a metal powder that includes particles of an alloy having a reactive element that renders the alloy uncastable. The composition is metal injection molded to yield a green compact having a shape corresponding to the shape of the metallic component, including its at least one nonuniform section. A majority of the binder is then removed from the green compact, and then the green compact is sintered to remove a remainder of the binder and fuse particles of the metal powder together to form the metallic component and the nonuniform section thereof.03-29-2012
20120260659INTERFACE BETWEEN A COMBUSTOR BASKET AND A TRANSITION OF A GAS TURBINE ENGINE - An interface (10-18-2012
20110120132DUAL WALLED COMBUSTORS WITH IMPINGEMENT COOLED IGNITERS - A combustor for a gas turbine engine includes an inner liner and an outer liner circumscribing the inner liner and forming a combustion chamber therewith. The outer liner is a dual walled liner with a first wall and a second wall. A fuel igniter includes a tip portion configured to ignite an air and fuel mixture in the combustion chamber. An igniter support assembly positions the fuel igniter relative to the combustion chamber. The igniter support assembly defines a plurality of holes configured to direct cooling air toward the tip portion of the fuel igniter. The igniter support assembly includes first and second floating seals that are configured to accommodate radial and axial relative movements.05-26-2011
20120317987HOT GAS PATH COMPONENT - A hot gas path component is provided and includes a body having a surface and being formed to define a cavity, the cavity employing coolant flow through a pin-fin bank with coolant discharge through film-cooling holes defined on the surface, the pin-fin bank including first and second pluralities of pin-fins, the first plurality of pin-fins and the second plurality of pin-fins each being aligned with a determined flow streamline, and any two pin-fins of the first and second pluralities of pin-fins being separated from one another by a gap as a function of a film-cooling hole dimension.12-20-2012
20100229565WALL OF A ROCKET ENGINE - A wall configured to withstand a high thermal load, said wall having a surface that is turned away from a heat source and provided with at least one strengthening element that, from a base portion thereof, protrudes from said surface. Said at least one strengthening element presents a decreasing width from its base portion towards a top portion thereof. The wall is an engine wall of a thrust nozzle of a rocket engine.09-16-2010
20100229564COMBUSTOR LINER COOLING SYSTEM - A system, in one embodiment, includes a turbine engine. The turbine engine includes a combustor that includes a hollow annular wall having a combustor liner. The turbine engine also includes first flow path in a first direction through the hollow annular wall. The turbine engine further includes a second flow path in a second direction that is opposite the first direction through the hollow annular wall. The second flow path may include one or more film holes configured to supply a cooling film to a downstream end portion of the combustor liner.09-16-2010
20100229563Wall elements for gas turbine engine combustors - The invention relates to a wall element for a wall structure of a gas turbine engine combustor. The wall element has an inner, in use, hot surface, and an outer, in use, cooler surface. A plurality of projections is provided on the outer surface to facilitate heat transfer to a coolant flow. The wall element comprises a means to direct more coolant flow at a hot-spot on an adjacent tile than the remainder of the adjacent tile and thereby reducing the thermal gradient across the adjacent tile.09-16-2010
20120324898COMBUSTOR ASSEMBLY FOR USE IN A TURBINE ENGINE AND METHODS OF ASSEMBLING SAME - A combustor assembly for use with a turbine engine that includes a rotor assembly. The combustor assembly includes a casing that includes a plenum and a combustor liner that is spaced a distance from the plenum and that defines a combustion chamber therein. A transition nozzle extends between the combustor liner and the rotor assembly for channeling combustion gases from the combustion chamber to the rotor assembly. The transition nozzle includes a transition portion and a nozzle portion integrally formed with the transition portion. An annular flowsleeve is coupled radially outward from the transition nozzle such that an annular flow path is defined between the flowsleeve and the transition nozzle. The flowsleeve includes a plurality of openings extending through an outer surface of the flowsleeve for providing flow communication between the plenum and the annular flow path to facilitate impingement cooling of the flowsleeve.12-27-2012
20120324897METHODS AND SYSTEMS FOR TRANSFERRING HEAT FROM A TRANSITION NOZZLE - Methods and systems are provided for transferring heat from a transition nozzle. The transition nozzle includes a transition portion, a nozzle portion integrally formed with the transition portion, and at least one surface feature configured to transfer heat away from the transition portion and/or the nozzle portion. The transition portion is oriented to channel the combustion gases towards the nozzle portion.12-27-2012
20100199678Corrosion-Resistant Pressure Vessel Steel Product, a Process for Producing It and a Gas Turbine Component - A corrosion-resistant pressure vessel steel is provided. The corrosion-resistant pressure vessel is formed by aluminization of a pressure vessel steel and deliberate oxidation of the layer which has been enriched in this way. Also provided is a process for producing a corrosion-resistant pressure vessel product. The process includes alitizing the pressure vessel steel and then oxidizing the steel before it is used for the first time. Lastly, a gas turbine component using the pressure vessel steel product is provided.08-12-2010
20100199677Transition Duct Assemblies and Gas Turbine Engine Systems Involving Such Assemblies - Transition duct assemblies and gas turbine engine systems involving such assemblies are provided. In this regard, a representative a transition duct assembly for a gas turbine engine includes: an impingement sheet having cooling holes formed therethrough, an inlet end and a non-flanged outlet end, the impingement sheet being operative to be positioned about an exterior of a transition duct such that cooling air is directed to flow about the transition duct; the non-flanged outlet end of the impingement sheet being operative to attach the impingement sheet to the transition duct such that the inlet end is positioned adjacent to an intake end of the transition duct and the outlet end is positioned adjacent to an exhaust end of the transition duct.08-12-2010
20120137697COMBUSTION CHAMBER FOR A TURBOMACHINE INCLUDING IMPROVED AIR INLETS - A combustion chamber for a turbomachine, including two coaxial walls including air inlets, each of which is configured such that its orthogonal projection, in a plane passing through the axis of the injection system closest to the inlet and perpendicular to an axial plane passing through this axis and through the axis of the combustion chamber, has an upstream edge of convex shape when seen from downstream.06-07-2012
20130014510COATED GAS TURBINE COMPONENTSAANM Pater; Christopher M.AACI TollandAAST CTAACO USAAGP Pater; Christopher M. Tolland CT US - A gas turbine component subject to extreme temperatures and pressures includes a wall defined by opposite first and second surfaces. An airflow aperture through the wall is defined by an aperture wall surface which extends from a first opening in the first surface to a second opening in the second surface. The aperture wall surface is flared at a juncture with the first surface, such that the first opening has a greater cross-sectional flow area than the second opening. A high-pressure, high-temperature coating is adhered to the first surface, and adhered to at least a portion of the aperture wall surface.01-17-2013
20130167543METHODS AND SYSTEMS FOR COOLING A TRANSITION NOZZLE - A transition nozzle for use with a turbine assembly is provided. The transition nozzle includes a liner defining a combustion chamber therein, a wrapper circumscribing the liner such that a cooling duct is defined between the wrapper and the liner, a cooling fluid inlet configured to supply a cooling fluid to the cooling duct, and a plurality of ribs coupled between the liner and the wrapper such that a plurality of cooling channels are defined in the cooling duct.07-04-2013
20110232290PRESS-FITTING CORROSION RESISTANT LINERS IN NOZZLES AND CASINGS - An apparatus and method for protecting an inner radial surface of a housing of a turbomachine from corrosion. The method includes tapering the inner radial surface of the housing and a corresponding outer radial surface of a corrosion-resistant liner, and heating the housing to increase a diameter of the inner radial surface of the housing. The method also includes inserting the corrosion-resistant liner at least partially into the housing, and attaching the corrosion-resistant liner to the inner radial surface of the housing a solid-state bonding process.09-29-2011
20130174562GAS TURBINE ENGINE, COMBUSTOR AND DOME PANEL - One embodiment of the present invention is a unique dome panel for a gas turbine engine combustor. Another embodiment is a unique gas turbine combustor. Yet another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and combustion systems and components. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.07-11-2013
20130111912FLEXIBLE METALLIC SEAL FOR TRANSITION DUCT IN TURBINE SYSTEM - A turbine system is disclosed. In one embodiment, the turbine system includes a transition duct. The transition duct includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct further includes an interface member for interfacing with a turbine section. The turbine system further includes a flexible metallic seal contacting the interface member to provide a seal between the interface member and the turbine section.05-09-2013
20130111911LEAF SEAL FOR TRANSITION DUCT IN TURBINE SYSTEM - A turbine system is disclosed. In one embodiment, the turbine system includes a transition duct. The transition duct includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct further includes an interface member for interfacing with a turbine section. The turbine system further includes a leaf seal contacting the interface member to provide a seal between the interface member and the turbine section.05-09-2013
20130111910TRANSITION PIECE AFT FRAME - A transition piece aft frame is provided and includes a manifold having an interior that is receptive of fuel and formed to define fuel injection holes configured to inject the received fuel from the manifold interior toward a main flow of products of combustion flowing through the manifold. The manifold includes a main body having an interior facing surface that faces the main flow of the products of the combustion and along which the fuel injection holes are defined.05-09-2013
20130111909Combustion System Having A Venturi For Reducing Wakes In An Airflow - A combustion system is provided having a liner, a flow sleeve, a flow-obstructing element, and a venturi. The liner is disposed around a combustion region. The flow sleeve is disposed around the liner. The liner and the flow sleeve cooperate to create an air passage having an airflow located between the liner and the flow sleeve. The flow-obstructing element is disposed within the air passage, and generally obstructs the airflow in the air passage to create wakes in the airflow. The venturi is disposed downstream from the flow-obstructing element, and generally restricts and diffuses the airflow in the air passage to reduce wakes in the airflow.05-09-2013
20130091847COMBUSTOR LINER - A combustor liner is provided. The combustor liner has a first end, including a forward end liner formed to define a converging interior through which a main flow is directed to flow and a second end. The second end is fluidly coupled to an aft portion of the first end and includes an aft end liner formed to define a diverging interior receptive of the main flow and through which the main flow is directed to continue to flow.04-18-2013
20130091848ANNULAR FLOW CONDITIONING MEMBER FOR GAS TURBOMACHINE COMBUSTOR ASSEMBLY - A turbomachine combustor assembly includes a combustor body, a combustor liner arranged within the combustor body and defining a combustion chamber, a fluid passage defined between the combustor body and the combustor liner, and an annular flow conditioning member arranged in the fluid passage and extending about the combustor liner.04-18-2013
20130125551RETRACTABLE EXHAUST LINER SEGMENT FOR GAS TURBINE ENGINES - A retractable exhaust liner segment between a gas turbine engine and an exhaust duct.05-23-2013
20110219774CIRCUMFERENTIALLY VARIED QUENCH JET ARRANGEMENT FOR GAS TURBINE COMBUSTORS - A combustor for a turbine engine is provided. The combustor includes a first liner; a second liner positioned relative to the first liner to form a combustion chamber therebetween, the combustion chamber configured to receive a fuel-air mixture; an igniter positioned relative to the combustion chamber and configured to ignite the fuel-air mixture; a first group of air admission holes positioned in the first liner and forming a regular circumferential pattern around the first liner; and a second group of air admission holes positioned in the first liner at a first circumferential position corresponding to the igniter, the second group of air admission holes departing from the regular circumferential pattern.09-15-2011
20120255307GAS TURBINE ENGINE TRANSITION DUCTS - A gas turbine engine transition duct 10-11-2012
20100293957SYSTEM AND METHOD FOR COOLING A WALL OF A GAS TURBINE COMBUSTOR - A system, in one embodiment, includes an engine wall. The engine wall includes a cold-side and a hot-side. The engine wall includes one or more dilution holes, wherein each of the one or more dilution holes includes a first opening on the cold-side, a second opening on a hot-side, and a section of thermal barrier coating (TBC) applied on the cold-side and having an opening that generally circumscribes the first opening.11-25-2010
20100307162Heat shield element arrangement and method for installing a heat shield element - A heat shield element arrangement including a heat shield element for a heat shield arranged on a supporting structure is provided. On each of the two opposing sides running parallel to the installation grooves the heat shield element includes a continuous screw head opening which penetrates the cold side and the hot side of the heat shield element substantially vertically and through which the screw head of the respective screw is arranged to be accessible to the supporting structure, and a spring element may be arranged under the respective screw, which spring element may be extended along the hot side of the heat shield element and the parallel installation groove of the supporting structure. An outer end of the spring element is embodied as a clamping retaining hook which is provided for the purpose of engaging in the laterally recessed retaining groove of the heat shield element.12-09-2010
20110302924COOLED CONDUIT FOR CONVEYING COMBUSTION GASES - A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having a central axis and defining an inner volume of the conduit for permitting hot combustion gases to pass through the conduit. The wall structure includes a forward end, an aft end axially spaced from the forward end, the aft end defining a combustion gas outlet for the hot combustion gases passing through the conduit, and a plurality of generally radially outwardly extending protuberances formed in the wall structure. The protuberances each include at least one cooling fluid passage formed therethrough for permitting cooling fluid to enter the inner volume. At least one of the protuberances is shaped so as to cause cooling fluid passing through it to diverge in a circumferential direction as it enters into the inner volume.12-15-2011
20120017596COMBUSTORS WITH QUENCH INSERTS - A combustor for a turbine engine is provided. 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 configured to receive an air-fuel mixture for combustion therein; and an insert including a body portion extending through the first liner, a shoulder circumscribing the body portion and abutting the first hot side, and an inlet portion coupled to the body portion and abutting the first cold side such that the inlet portion and the shoulder capture the second liner therebetween to retain the insert.01-26-2012
20130192233HEAT SHIELD FOR A COMBUSTOR - A heat shield includes a panel having a forward face and an aft face, a circumferential outer side and a circumferential inner side such that the panel subtends a predetermined arc, and a first radially extending side and a second radially extending side each extending from the circumferential outer side to the circumferential inner side. The panel defines an opening extending between the forward face and the aft face, a lip projecting from the forward face and extending around the opening, a rail projecting from the forward face and extending around the lip to define a cavity region between the lip and the rail, and a plurality of holes in the cavity region. Each of the plurality of holes extends from the forward face to the aft face. A region of the panel extending from the lip to the circumferential outer side is free of any aft-projecting features.08-01-2013

Patent applications in class Combustor liner

Patent applications in all subclasses Combustor liner