Abstract: The present application provides for a combustor for combusting a flow of fuel and a flow of air. The combustor may include a number of fuel nozzles, a lean pre-nozzle fuel injection system positioned upstream of the fuel nozzles, and a premixing annulus positioned between the fuel nozzles and the lean pre-nozzle fuel injection system to premix the flow of fuel and the flow of air.

Abstract: A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having an inner surface, an outer surface, a region, an inlet, and an outlet. The inner surface defines an inner volume of the conduit. The region extends between the inner and outer surfaces and includes cooling fluid structure defining a plurality of cooling passageways. The inlet extends inwardly from the outer surface and provides fluid communication between the inlet and the passageways. The outlet extends from the passageways to the inner surface to provide fluid communication between the passageways and the inner volume. At least one first cooling passageway intersects with at least one second cooling passageway such that cooling fluid flowing through the first cooling passageway interacts with cooling fluid flowing through the second cooling passageway.

Abstract: A nut (64) is affixed to an outer surface of a transition impingement sleeve forward ring (50) that encircles, and is affixed to, a forward end (44) of a tubular transition impingement sleeve (45). The nut has a threaded hole (63) aligned with a hole (66) in the impingement sleeve forward ring. A machine screw (68) is threaded into the nut and extends through the hole (66), and has a radially inner end with a wear pad (70), and a radially outer end with a turning tool engagement element (72). The wear pad contacts an outer surface of an aft portion of a transition piece forward outer ring (52) that is surrounded by the transition impingement sleeve forward ring (50). The rotational position of the machine screw (68) sets a radial gap (76) between the transition impingement sleeve forward ring and the transition piece forward outer ring.

Abstract: A gas turbine combustor including a fuel nozzle for injecting mixed gas of fuel and air, a cylindrical liner for burning and reacting the mixed gas of fuel and air in a combustion chamber, a transition piece which is a flow path for leading combustion gas generated in the liner to turbine blades, and a transition piece flow sleeve for wrapping an outside surface of the transition piece, wherein a plurality of air introduction holes for introducing air into the transition piece flow sleeve are formed in regions of the transition piece flow sleeve excluding regions which are corner portions of the transition piece flow sleeve in a sectional direction thereof.

Abstract: A venturi assembly for a turbine combustor includes a first outer annular wall and a second intermediate annular wall radially spaced from each other in substantially concentric relationship. The first outer annular wall and said second intermediate annular wall shaped to define a forward, substantially V-shaped throat region, and an aft, axially extending portion. A third radially innermost annular wall is connected to the second intermediate annular wall at an aft end of said throat region. A first plurality of apertures is provided in the first outer annular wall in the substantially V-shaped throat region, and a second plurality of apertures is provided in the aft, axially extending portion of said second intermediate annular wall so that cooling air flows through the first and second pluralities of apertures to impingement cool the third radially innermost annular wall.

Abstract: A system includes a gas turbine combustor, which includes a combustion liner disposed about a combustion region, a flow sleeve disposed about the combustion liner, an air passage between the combustion liner and the flow sleeve, and an aerodynamic mounting assembly disposed in the air passage. The aerodynamic mounting assembly is configured to retain the combustion liner within the flow sleeve. The aerodynamic mounting assembly includes a flow sleeve mount coupled to the flow sleeve and a liner stop coupled to the combustion liner. The flow sleeve mount includes a first portion of an aerodynamic shape and the liner stop includes a second portion of the aerodynamic shape, which is configured to direct an airflow into a wake region downstream of the aerodynamic mounting assembly. The flow sleeve mount and the liner stop couple with one another to define the aerodynamic shape.

Abstract: A system includes a gas turbine combustor, which includes a combustion liner disposed about a combustion region, a flow sleeve disposed about the combustion liner, an air passage between the combustion liner and the flow sleeve, and a structure between the combustion liner and the flow sleeve. The structure obstructs an airflow through the air passage. The gas turbine combustor also includes a wake reducer disposed adjacent the structure. The wake reducer directs a flow into a wake region downstream of the structure.

Abstract: A turbomachine combustor assembly includes a combustor body, and a combustor liner arranged within the combustor body and defining a combustion chamber. The combustor liner includes a venturi portion arranged within the combustion chamber. A fluid passage is defined between the combustor body and the combustor liner, and at least one turbulator is arranged in the fluid passage. The at least one turbulator is configured and disposed to create vortices in the fluid passage. A vortex modification system is arranged at the fluid passage and is configured and disposed to disrupt the vortices in the fluid passage.

Abstract: The present application provides a combustor with a radial penetration. The combustor may include a combustion chamber, a liner surrounding the combustion chamber, a flow sleeve surrounding the liner, a penetration tube extending through the liner and the flow sleeve with the radial penetration positioned within the penetration tube, a flange extending from the penetration tube about the flow sleeve, and a ferrule positioned about the flange and the radial penetration so as to limit leakage about the radial penetration.

Abstract: A gas turbine combustor comprising a fuel nozzle for injecting mixed gas of fuel and air, a cylindrical liner for burning and reacting the mixed gas of fuel and air in a combustion chamber, a transition piece which is a flow path for leading combustion gas generated in the liner to turbine blades, and a transition piece flow sleeve for wrapping an outside surface of the transition piece, wherein a plurality of air introduction holes for introducing air into the transition piece flow sleeve are formed in regions of the transition piece flow sleeve excluding regions which are corner portions of the transition piece flow sleeve in a sectional direction thereof.

Abstract: An annular combustor includes an annular outer shell that includes a first flange that defines an inner diameter IDOS and an annular inner shell that includes a second flange that defines an outer diameter ODIS. An annular hood includes a radially outer hood flange and a radially inner hood flange. A bulkhead includes a radially outer bulkhead flange that defines an outer diameter ODB and a radially inner bulkhead flange that defines an inner diameter IDB. The first flange is secured at a radially outer joint between the radially outer hood flange and the radially outer bulkhead flange. The second flange is secured at a radially inner joint between the radially inner hood flange and the radially inner bulkhead flange. The IDOS and the ODB define a ratio R1 of IDOS/ODB that is 0.998622-1.001129, and the IDB and the ODIS define a ratio R2 of IDB/ODIS that is 0.998812-1.001388.

Abstract: It is required for a gas turbine combustor to exhaust low NOx. The gas turbine combustor is provided with a combustion tube which has a cooling passage through which cooling air flows in a double wall structure. The cooling passage has a main cooling air supply opening opened to a side of a combustion zone. The cooling air supplied from the main cooling air supply opening is guided to a direction along an inner wall surface of the combustion tube by a guide. The cooling air flows through the cooling passage inside the combustion tube, and then is reused for film cooling along the inner wall surface. Thus, it is possible to save cooling air. Therefore, a more part of the air supplied from a compressor can be used as air for combustion and it becomes possible to exhaust low NOx.

Abstract: Methods and apparatus of an improved cooling structure having cooling passages are provided. A method includes forming pilot holes with an electrical discharge machine (EDM) drill near a heated flow surface. The pilot holes are then shaped with a wire EDM into cooling passages having a desired shape.

Abstract: Floatwall panel assemblies and related systems are provided. A floatwall panel assembly includes a panel formed of porous ceramic material, the porous ceramic material exhibiting a porosity gradient along at least one of a length, a width and a depth of the panel, the panel lacking a substrate, formed of a material other than porous ceramic material, for supporting the porous ceramic material.

Abstract: The present invention provides in one embodiment a unique gas turbine engine. Another embodiment is a unique gas turbine engine combustion system. Still another embodiment is a unique gas turbine engine combustor. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine engine combustion systems and combustors.

Abstract: 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.

Abstract: A combustor-to-transition seal (301) includes a spring clip assembly (322, 422, 522) and a first spring metal ring (340, 441) or a ring assembly (440,540) that includes the first spring metal ring (441) and a second spring metal ring (442). Such combustor-to-transition seal (301) provides a first seal (321) and a second seal (331, 531) when at a junction of a combustor (108, 308, 508) and a transition (114, 314, 514). In an embodiment, the first spring metal ring (340) has a plurality of apertures (342) through the ring (340) that provide effusion cooling of the ring (340). The plurality of apertures (342) regulates a flow of cooling fluid to maintain an acceptable temperature of the ring (340). In an alternative embodiment, the second spring metal ring (442) has grooves (446) that regulate a flow of cooling fluid to maintain an acceptable temperature of the ring assembly (440).

Abstract: 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.

Abstract: An annular seal for use between coupled combustor components includes a segmented annular solid edge portion and a plurality of alternating spring fingers and slots extending from the solid edge portion and arranged about a circumference of the solid edge portion, wherein one or more of the plurality of spring fingers or one or more of the plurality of straight and angled slots have non-uniform width dimensions.

Abstract: A combustor liner grommet is disclosed. The grommet may include a peripheral wall defining a hole in a combustor liner and further including at least one cooling air flow channel. The cooling air flow channel in the grommet wall may be a slot or a hole. The channel may increase cooling flow to the grommet and the combustor liner around the grommet to prevent cracking from heat stress.

Abstract: A combustor assembly includes a convergent segment followed by a divergent segment to advantageously improve combustion. The combustor assembly includes a first segment beginning at a forward end that transitions to a second segment past a transition segment in a direction along a combustor axis toward an aft end. The reduction in cross-sectional area within the first segment provides desirable fuel and air mixing properties. The convergent first segment in combination with the divergent second segment decreases residence time of fuel-air mixture within the combustor chamber that decreases production of undesirable emissions from the combustor assembly.

Abstract: 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.

Abstract: 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.

Abstract: An object of the present invention is to provide a gas turbine combustor that supports hydrogen-containing gas having a high burning velocity and is capable of performing low NOx combustion without reducing reliability of a burner. A first fuel nozzle is provided upstream of a combustion chamber and supplies fuel for activation and hydrogen-containing gas. The combustor has a primary combustion zone, a reduction zone and a secondary combustion zone. In the primary combustion zone, the fuel supplied from the first fuel nozzle is combusted under a fuel rich condition to form a burned gas containing a low concentration of oxygen. In the reduction zone, a hydrogen-containing gas is injected into the combustion chamber through a second fuel injection hole from a second fuel nozzle so that NOx generated in the primary combustion zone is reduced by an oxygen reaction of the hydrogen.

Abstract: A combustor liner is provided. The combustor liner may include an upstream portion and a downstream end portion. The upstream portion may have a radius and a length along a generally longitudinal axis. The downstream end portion may have a radius and a length along the generally longitudinal axis. The downstream end portion may define a plurality of channels. Each of the plurality of channels may extend helically through the length of the downstream end portion. Each of the plurality of channels may be configured to flow an air flow therethrough, cooling the downstream end portion.

Abstract: An annular combustion chamber for a gas turbine engine includes radially inner and outer walls connected together by a chamber end wall including openings, each of which receives a fuel injection system. Heat protection deflectors are fastened to the chamber end wall. Holes are formed through the chamber end wall to pass cooling air onto an upstream face of each deflector. The inner or outer edge of a deflector presents a sealing rim engaging the respective inner or outer wall of the chamber.

Abstract: A combustion chamber arrangement is described for operating a gas turbine, with a combustion chamber wall which encloses the combustion chamber space and in the region of the combustion chamber outlet encloses a flow passage for hot gases which develop inside the combustion chamber, has a combustion chamber wall edge which freely terminates in the axial flow direction of the hot gases. The combustion chamber wall edge is formed with a profile which blocks or at least inhibits a diffuser effect when a cooling air flow, which is guided axially through the flow passages into the annular spatial area, flows over the combustion chamber wall edge.

Abstract: A combustor for a gas turbine engine comprises an inner annular liner and an outer annular liner. First and second combustion stages are defined between the liners. Each combustion stage has a plurality of fuel injection bores distributed in a liner wall defining the respective stage. A lobed mixer extends into the combustor, the lobed mixer arranged to receive combustion gases from each combustion stage for mixing flows of said combustion gases.

Abstract: 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.

Abstract: A thermal machine, especially a gas turbine, includes an annular combustor which is outwardly delimited by an outer shell and an inner shell (33) and through which a hot gas axially flows. The outer shell and inner shell (33) are each provided with a concentric cooling shroud (31) which is attached at a distance on their outer side, forming a cooling passage (32) through which cooling passage (32) cooling air flows in a direction which is opposite to the hot gas flow. The cooling of the combustor is improved by at least one of the cooling shrouds (31), on the side on which the cooling air enters the cooling passage (32), having an outwardly curved, rounded inlet edge (37) for improving the inflow conditions.

Abstract: 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.

Abstract: 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.

Abstract: A system includes an air flow conditioner configured to mount in an air chamber separated from a combustion chamber of a turbine combustor. The air flow conditioner comprises a perforated annular wall configured to direct an air flow in both an axial direction and a radial direction relative to an axis of the turbine combustor. In addition, the air flow conditioner is configured to uniformly supply the air flow into air inlets of one or more fuel nozzles.

Abstract: Embodiments for an apparatus and associated method for providing a cooling fluid to a gas turbine transition duct in order to lower the effective operating temperatures of the transition duct are disclosed. The transition duct has an inner liner and an impingement sleeve positioned radially outward with a passageway formed therebetween. The impingement sleeve has a plurality of openings where a portion of the openings each have a feed tube extending through the opening and into the passageway. The feed tubes are oriented at an angle relative to the impingement sleeve, such that an inlet to the feed tube is directed generally towards an oncoming flow of cooling fluid. The feed tubes direct a portion of the cooling fluid toward the inner liner and into the passageway for cooling of the transition duct.

Abstract: A gas turbine engine includes an exhaust liner having cold and hot sheets radially spaced from one another and interconnected by a band arranged in a cavity between the hot and cold sheets. In one example, the band is Z-shaped to permit thermal growth of the hot sheets relative to the cold sheets in the axial and radial directions. The hot sheets include axially adjacent portions that provide slots that are in fluid communication with the space. Cooling fluid is provided to the cavity through impingement jets in the cold sheet. The slots are arranged radially between the hot sheet portions.

Abstract: A gas-turbine combustion chamber wall for a gas-turbine has a combustion chamber wall 9, on the inner side of which several tiles 10 are arranged, with an interspace 14 being formed between the tiles 10 and the combustion chamber wall 9, into which cooling air is introduced via impingement-cooling holes 8 provided in the combustion chamber wall 9 and from which the cooling air flows into the combustion chamber via effusion-cooling holes 11, 23 provided in the tile 10. The tile 10 includes a surface structure 19, 22 on the side facing the combustion chamber wall 9. The area of the impingement-cooling holes 8 and the area of the effusion-cooling holes 11 do not coincide.

Abstract: 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.

Abstract: A combustor liner is provided. The combustor liner may include an upstream portion and a downstream end portion. The upstream portion may have a radius and a length along a generally longitudinal axis. The downstream end portion may have a radius and a length along the generally longitudinal axis. The downstream end portion may define a plurality of channels. Each of the plurality of channels may extend helically through the length of the downstream end portion. Each of the plurality of channels may be configured to flow an air flow therethrough, cooling the downstream end portion.

Abstract: A transition duct (30) for a gas turbine engine (2) having improved cooling and reduced stress levels. The transition duct may be formed of two panels ((36, 38) joined together with welds (40) disposed remote from the bent corner regions (34) of the panels. Cooling channels (32) extending longitudinally in the direction of flow of the hot combustion gas carried by the duct are formed within each panel, including the corner regions. Because the entire annular width (W) of the transition duct is cooled, the gap (G) separating adjacent ducts around the inlet to the turbine (4) may be reduced when compared to prior art designs. Two-panel construction with welds remote from the corner regions is facilitated by maintaining the minimum bend radius in the corners (R2) and in the direction of flow (R4) to be greater than in prior art designs.

Abstract: A gas turbine engine diffuser defined between an external casing and an internal casing of the engine and supplied with air via an upstream annular diffuser duct is disclosed. The diffuser includes a combustion chamber of the convergent type, forming an external annular duct with the external casing and an internal annular duct with the internal casing, and a cowling partially closing off the external annular duct. The cowling is positioned toward the closed end of the combustion chamber.

Abstract: A gas turbine engine has a combustor module including an annular combustor having a liner assembly that defines an annular combustion chamber. The liner assembly includes an inner liner and an outer liner that circumscribes the inner liner. The inner liner has a plurality of first combustion air admission holes passing therethrough and the second liner has a plurality of second combustion air admission holes passing therethrough. Each of the first and the second combustion air admission holes has a swirler disposed therein.

Abstract: A combustor liner (231) for a gas turbine engine combustor (200) comprises an inner wall (232), an outer wall (238), a flow channel (244) formed there between, and an end-capping ring (246). The end-capping ring (246) is sealingly attached to the downstream end of the inner wall (232). In operation air passes within the end-capping ring (246), into the flow channel (244), and through holes (250) disposed in the inner wall (232). In some embodiments, an end-capping ring variation, a flow-diverting ring (357) comprises a plurality of holes (360) that, during gas turbine engine operation, may additionally dispense a flow of cooling air. One or more surfaces may be coated with a thermal barrier coating (237) to provide additional protection from thermal damage.

Abstract: A gas turbine engine combustor having a dome heat shield includes a cooling scheme having a plurality of impingement cooling holes extending through the combustor and a plurality of adjacent ejector holes for directing cooling air past the heat shield lips of the dome heat shields. The impingement and ejector holes are preferably staggered to reduce interaction therebetween.

Abstract: A method facilitates assembling a gas turbine engine including a combustor assembly and a nozzle assembly. The method comprises providing a transition piece including a first end, a second end, and a body extending therebetween, where the body includes an inner surface, an opposite outer surface, coupling the first end of the transition piece to the combustor assembly, and coupling the second end of the transition piece to the nozzle assembly such that a turbulator extending helically over the outer surface of the transition piece extends from the transition piece first end to the transition piece second end to facilitate inducing turbulence to cooling air supplied to the combustor assembly.

Abstract: 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.

Abstract: A combustor for a gas turbine is provided having a nozzle assembly located at one end and a combustion chamber defining a second end of the combustor. A venturi is positioned within the combustor, between the nozzle and the combustion chamber. The venturi defines a passageway therein having a first side facing the nozzle and a second side facing the combustion chamber. Compressed air is directed into an inlet in fluid communication with the first and second sides of the venturi passageway. The venturi passageway directs the compressed air from the inlet in opposite directions within the first and second sides of the passageway for cooling the venturi.

Abstract: A combustor for a gas turbine is provided having a nozzle assembly located at one end and a combustion chamber defining a second end of the combustor. A venturi is positioned within the combustor, between the nozzle and the combustion chamber. The venturi defines an internal passageway therein having a first side facing the nozzle and a second side facing the combustion chamber. Compressed air is directed into an inlet in fluid communication with the first and second sides of the internal passageway. The internal passageway directs the compressed air from the inlet in opposite directions within the first and second sides of the internal passageway for cooling the venturi.

Abstract: An downstream plasma boundary layer shielding system includes film cooling apertures disposed through a wall having cold and hot surfaces and angled in a downstream direction from a cold surface of the wall to an outer hot surface of the wall. A plasma generator located downstream of the film cooling apertures is used for producing a plasma extending downstream over the film cooling apertures. Each plasma generator includes inner and outer electrodes separated by a dielectric material disposed within a groove in the outer hot surface. The wall may be part of a hollow airfoil or an annular combustor or exhaust liner. A method for operating the downstream plasma boundary layer shielding system includes forming a plasma extending in the downstream direction over the film cooling apertures along the outer hot surface of the wall. The method may further include operating the plasma generator in steady state or unsteady modes.

Abstract: A combustor liner includes a forward end and an aft end, the aft end having a reduced diameter portion and a cooling and dilution sleeve overlying the reduced diameter portion thereby establishing a cooling plenum therebetween. A plurality of cooling and dilution air entry holes are formed in the cooling and dilution sleeve and a plurality of cooling and dilution air exit holes formed adjacent an aft edge of the liner such that, in use, cooling and dilution air flows through the cooling and dilution air entry holes, and through the plenum, exiting the cooling and dilution air exit holes, thereby cooling and dilution tuning the aft end of the combustor liner without having to remove the transition piece.

Abstract: 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.