Abstract: A reverse flow combustor for a gas turbine engine having an outer combustor liner and an inner combustor liner defining an annular combustion chamber, and a compound-angle frustoconical portion in the outer liner having a first and second conical slopes towards an engine centerline.

Abstract: A combustor for a gas turbine engine having an annular combustion chamber includes a plurality of main fuel injection and air swirler assemblies and a plurality of pilot fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of a forward bulkhead. The plurality of pilot fuel injection and air swirler assemblies are interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies. Fuel being supplied to the combustor is selectively distributed between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies in response to the level of power demand on the gas turbine engine.

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 turbomachine combustion chamber in which a chamber end wall presents an opening for receiving a pre-vaporization bowl, and including a device for injecting air and fuel mounted on the axis thereof, the bowl being floatingly mounted relative to the chamber end wall to move in a predetermined radial direction and flaring downstream so as to form a collar, a deflector forming a thermal shield being made integrally with the bowl beside the chamber end wall so that the floating movement of the bowl-and-deflector assembly takes place in a sliding plane situated between the deflector and the chamber end wall.

Abstract: A gas turbine engine combustion chamber including at least one deflector mounted on the chamber end wall and including an opening for a carburetted air supply device. The deflector includes an opening, corresponding to the chamber end wall opening, with an annular cylindrical part for attachment to the wall, the cylindrical part including a mechanical attachment mechanism collaborating with a complementary attachment mechanism on a metal sleeve secured to the wall and a cylindrical centering cup fixed by one end to the sleeve and housed inside the cylindrical part of the deflector.

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: A system includes a multi-tube fuel nozzle having an inlet plate and a plurality of tubes adjacent the inlet plate. The inlet plate includes a plurality of apertures, and each aperture includes an inlet feature. Each tube of the plurality of tubes is coupled to an aperture of the plurality of apertures. The multi-tube fuel nozzle includes a differential configuration of inlet features among the plurality of tubes.

Abstract: A head part of an annular gas turbine combustion chamber including: an end wall with a passage opening for accommodating a burner, the end wall including a set back portion adjacent the passage opening; a heat shield covering a back side of the end wall which faces towards the combustion chamber, the heat shield including a protruding portion shaped to cooperate with the set back portion of the end wall; and a burner collar adapted to fit within the passage opening and receive a burner, the burner collar including a protruding portion radially protruding from an outer surface of the burner collar; wherein the head part of the annular gas turbine combustion chamber is configured such that in an installed configuration the protruding portion of the burner collar is held between the protruding portion of the heat shield and the set back portion of the end wall.

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 system includes a turbine combustor containing a first wall portion disposed about a combustion zone, a second wall portion disposed about the first wall portion, and a plurality of vanes disposed within an annulus between the first and second wall portions. The plurality of vanes is configured to swirl air flowing within the annulus.

Abstract: The present invention discloses a novel apparatus and way for controlling a velocity of a fuel-air mixture entering a gas turbine combustion system. The apparatus comprises a hemispherical dome assembly which directs a fuel-air mixture along a portion of the outer wall of a combustion liner and turns the fuel-air mixture to enter the combustion liner in a manner coaxial to the combustor axis and radially outward of a pilot fuel nozzle so as to regulate the velocity of the fuel-air mixture.

Abstract: A baffle for the bottom of a combustion chamber of a gas turbine engine is disclosed. The baffle includes a flat portion of wall with an opening for an injector of the combustion chamber to pass through, two longitudinal edges for the assembly to two adjacent baffles and two transverse edges. At least one of the edges includes a joint cover arranging a housing along the edge for an edge of an adjacent baffle so as to seal the junction between the two edges. A combustion chamber incorporating the baffles is also disclosed.

Abstract: A combustion chamber head of a gas turbine has a confinement enclosing a dampening volume (207) and including a combustion chamber-opposite confinement (206) and a combustion chamber-side confinement (210). The combustion chamber-side confinement (210) is provided as perforated wall (210). In the edge area of the combustion chamber-side confinement (210), cooling air can be routed onto the combustion chamber-side confinement (210) via recesses (203) in the confinement (206). This cooling air, which flows along the combustion chamber-side confinement (210), crosses the cooling air flow through the perforated wall (210) in the combustion chamber (101) without mixing with the latter, as both are separated by walls.

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 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: 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 combustor assembly for a turbine engine includes a combustor liner and a flow sleeve which surrounds the combustor liner. Compressed air flows through an annular space located between an outer surface of the combustor liner and an inner surface of the flow sleeve. A plurality of cooling holes are formed through the flow sleeve to allow compressed air to flow from a position outside the flow sleeve, through the cooling holes, and into the annular space. The height of the annular space may vary along the length of the combustor assembly. Thus, the flow sleeve may have reduced diameter portions which result in the height of the annular space being smaller in certain locations than at other locations along the length of the combustor assembly.

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 method for assembling a gas turbine engine combustor is provided. The method includes providing a heat shield defined by a perimeter. The perimeter includes a radially inner edge, a radially outer edge, an axially inner edge, an axially outer edge, and an opening that extends from an upstream side of the heat shield to a downstream side of the heat shield. The method further includes coupling the heat shield to a domeplate such that the perimeter of the heat shield is positioned a distance downstream from an edge of the heat shield defining the opening. The method additionally includes coupling at least one fuel injector to the domeplate such that a portion of the fuel injector extends through the heat shield opening.

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 combustor liner cap assembly is provided for use in a multiple fuel nozzle combustor of a gas turbine. The combustor liner cap assembly includes a tube plate having a plurality of fuel nozzle openings and a plurality of open ended premix tubes extending aft from the tube plate. Each premix tube has a forward end having a forward edge, the forward end being received in a corresponding one of the fuel nozzle openings. Each premix tube is secured to the tube plate aft of the forward edge of the forward end of the premix tube, for example, metallurgically bonded, such as by welding or brazing, to the tube plate or threaded into the tube plate.

Abstract: The present application provides a fuel nozzle spring support system. The fuel nozzle spring support system may include a fuel nozzle, a cap assembly, and a spring support positioned between the fuel nozzle and the cap assembly.

Abstract: 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.

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: 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 section panel assembly for a bulkhead of a gas turbine combustor includes a section panel defining a rear or cold side and defining a front or hot side. The panel has a circumferential inner edge, a circumferential outer edge and radially extending first and second side edges each extending from the inner edge to the outer edge such that the inner and outer edges subtend a predetermined arc. The panel defines at least one opening extending from the rear side to the front side for receiving cooling air. The rear side includes a central lip extending about a periphery of the opening, and a circumferential rail spaced radially outwardly of and extending about the opening for defining a cavity region including the opening on the rear side of the section panel.

Abstract: A device for supplying flow to a combustor includes a flow sleeve configured to circumferentially surround the combustor, wherein the flow sleeve defines a first annular passage around the combustor. A first section of the first annular passage converges at a first convergence rate. A second section of the first annular passage downstream from the first section converges at a second convergence rate that is less than the first convergence rate. A method for supplying flow to a combustor includes flowing a first portion of a working fluid substantially axially through a first annular passage, converging the first annular passage at a first convergence rate, and converging the first annular passage at a second convergence rate downstream from the first convergence rate, wherein the second convergence rate is less than the first convergence rate.

Abstract: A device for supplying flow across a combustor includes an axial fluid injector configured to circumferentially surround at least a portion of the combustor. An inner annular passage extends through the axial fluid injector and provides fluid communication through the axial fluid injector and into a first annular passage that surrounds the combustor. An outer annular passage extends through the axial fluid injector radially outward from the inner annular passage and provides axial flow into the first annular passage. A method for supplying flow to a combustor includes flowing a first portion of a working fluid through a first axial flow path and flowing a second portion of the working fluid through a second axial flow path.

Abstract: The present application provides a fuel nozzle assembly. The fuel nozzle assembly may include an end cap assembly, a number of fuel nozzles positioned within the end cap assembly, and one or more cup seals. The cup seals may be positioned between the end cap assembly and the fuel nozzles.

Abstract: A combustion chamber having a burner and a burner insert surrounding the burner while leaving a gap open towards the combustion chamber interior is provided. The burner insert comprises a carrier and a burner insert wall located in front of the carrier towards the combustion chamber interior, wherein a flow passage connected to a cooling-fluid source is formed between the carrier and the burner insert wall. The flow passage opens out into the gap between the burner and the burner insert and is otherwise sealed off from the combustion chamber interior.

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 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: Systems and methods for an oxy-fuel type combustion reaction are provided. In one or more embodiments, a combustion system can include at least two mixing zones, where a first mixing zone at least partially mixes oxygen and carbon dioxide to produce a first mixture and a second mixing zone at least partially mixes the first mixture with a fuel to produce a second mixture. The combustion system can also include a combustion zone configured to combust the second mixture to produce a combustion product. In one or more embodiments, the first mixture can have a spatially varied ratio of oxygen-to-carbon dioxide configured to generate a hot zone in the combustion zone to increase flame stability in the combustion zone.

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: An annular combustion chamber for a gas turbine engine is disclosed. The combustion chamber includes an external wall, an internal wall, a chamber end wall connecting these two walls and provided with openings for the fuel injection systems, and heat shielding deflectors fixed to the chamber end wall. The deflectors include a flat wall portion with an opening centered on said openings for the fuel injection systems, two longitudinal edges, and two transverse edges. Along one of the longitudinal edges, the deflector includes a tongue forming a joint cover, creating a housing along the edge for the edge of the adjacent deflector so as to seal the join between the two adjacent edges of the two deflectors. The tongue is spaced away from the chamber end wall so as to create a space supplied with cooling air via orifices in the chamber end wall.

Abstract: A turbomachine combustion chamber including a chamber bottom with at least one opening designed to receive a combustion bowl in the axis of which an air and fuel injection device is mounted is disclosed. The downstream end of the flared bowl includes a divergent with a double partition delimiting an annular cavity. The first outer partition includes inlet orifices arranged to cool the second inner partition by impact and the second inner partition includes outlet orifices. The inlet orifices, distributed in at least two circular rows on the periphery of the divergent, are in staggered rows with the outlet orifices.

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 deflector takes the form of a plate provided with a hole. The plate is a portion of a conical surface of revolution about a cone axis with a concave face and a convex face. The plate includes a contour with four sides. Two of these the sides are concentric circular arcs centered on the cone axis. The other two sides are segments of generatrices of the cone which connect the first sides. The invention is applicable to a deflector of a chamber endwall of a combustion chamber of a turbine engine.

Abstract: A combustion chamber of a combustion system comprises a combustion space, a support structure, a heat shield, and at least one through-opening. The heat shield has at least two segments, and each segment has an edge region, a gap communicating with the combustion space being formed between edge regions of adjacent segments, and a retaining device. The adjacent segments include a support element disposed in a bottom region of the gap. The retaining device fixes the respective liner element on the support structure via the respective support element. The at least one through-opening communicates with the gap so as to enable a cooling gas to flow through the through-opening. The at least one through-opening is disposed in at least one of the respective edge region and the support element at the bottom region of the gap.

Abstract: A gas turbine engine reverse-flow sheet metal combustor for an aero gas turbine engine has a louverless single-piece liner design which includes a plurality of effusion patches, containing multiple rows of effusion cooling holes, bounded by cooled effusionless bands to aid in repairability of the combustor. The cooling of the bands improves the durability of the combustor and provides a method of repair for the combustor.

Abstract: An annular wall which transversely connects longitudinal walls of an annular combustion chamber of a turbine engine is disclosed. The wall is essentially flat, inclined in relation to a longitudinal axis of the turbine engine, and includes a plurality of deflectors, each formed by an essentially rectangular flat sheet. The deflectors are mounted on the wall and each includes an aperture for the installation of a fuel injection system, a plurality of multi-perforation holes formed in relation to the deflectors around their aperture so as to allow a passage of air intended for the cooling of the deflectors, and a flow forcing unit which forces the flow of air for cooling the deflectors to flow radially around the fuel injection systems.

Abstract: A combustion chamber which includes a chamber endwall which takes the form of a substantially frustoconical component including a succession of adjacent planar portions is disclosed. The invention can be applied to a combustion chamber of a turbine engine.

Abstract: A turbomachine combustion chamber has primary and dilution air inlet orifices formed by die stamping to have edges that project into the inside of the combustion chamber, and stress relaxation and/or reduction means in the edges or in the vicinity of the edges of said orifices, said means comprising, for each orifice, one, two, or three slots formed in the edge or around a fraction of the edge of said orifice.

Abstract: A method and apparatus to facilitate reducing NOx emissions in turbine engines is provided. The method includes providing an annular shell including a plurality of circumferentially extending panels. The plurality of circumferentially extending panels includes a first panel positioned at an upstream end of the shell and a second panel positioned downstream from, and adjacent to, the first panel. The method also includes forming a plurality of primary dilution holes in the first panel and forming a plurality of secondary dilution holes in the second panel. The dilution holes are configured to discharge dilution air into the shell.

Abstract: A gas turbine combustor is provided with a dome heat shield having a fuel nozzle opening, the opening receiving a floating collar assembly for permitting relative movement between nozzle and heat shield. The floating collar is provided with a louver to provide film cooling to the face of the combustor heat shield and, thus, improve cooling thereof.

Abstract: 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.

Abstract: 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.

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 combustion chamber having a chamber end wall that presents a plurality of air and fuel injector devices, each of the devices including a bowl and a deflector mounted on a main axis forming a heat shield disposed around the bowl which flares in the gas flow direction and possesses an annular portion, the interface between the bowl and the deflector being situated on this annular portion, at least a portion of the interface being situated towards the end of the annular portion that is downstream in the gas flow direction, such that an annular cavity is defined between the deflector and the downstream annular portion of the interface. First bowl holes are formed in the annular portion to open out facing the edge of the chamber end wall that is in contact with the deflector in order to sustain a flow of air from upstream to downstream in the annular cavity.

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.