Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a liner at least partially defining a combustion chamber of the gas turbine engine, wherein the liner comprises a looped feature.

Abstract: In accordance with an embodiment of the disclosure, a system includes a combustor liner disposed about a combustion chamber of a combustor of a gas turbine system. The combustor liner includes an inner wall portion exposed to the combustion chamber, an outer wall portion disposed about the inner wall portion, and multiple channels between the inner and outer wall portions of the combustor liner. Each channel of the multiple channels is configured to direct a coolant along the combustor liner to convectively cool the combustor liner, and each channel of the multiple channels includes a cross-sectional area that progressively changes along a length of each channel of the plurality of channels.

Abstract: To provide a gas turbine combustor that can suppress a generation amount of NOx and maintain a flame holding property, while suppressing burn damage around a pilot nozzle including the pilot nozzle. A gas turbine combustor includes a pilot nozzle that can inject fuel F and cooling air A for cooling a nozzle tip, a flow regulating valve that can adjust a flow rate of cooling air to be supplied to the pilot nozzle, a detection sensor that detects a combustion state of fuel, and a control device that controls the flow regulating valve based on a detection result of the detection sensor.

Abstract: A method for perforating a wall, includes calculating mechanical stresses applied on the wall for a use of the wall. The method further includes perforating an orifice in a first determined zone of the wall, the perforation being made using a tool with a cross-section dependent on the mechanical stresses calculated in the first determined zone.

Abstract: A combustor assembly (17) including guide vanes (44) located between an inner cylinder (24) and a flow sleeve (25). Each guide vane (44) includes a circumferentially angled flow directing portion (60) adjacent to a leading edge (46). The leading edge (46) of at least one guide vane (44) can be located radially inward along the longitudinal axis (54) relative to the leading edge (46) of at least one other of the guide vanes (44). The length of the guide vanes (44) may vary, and the circumferential spacing between a first pair of the guide vanes (44) can be different from a spacing between a second pair of the guide vanes (44).

Abstract: The present invention relates to a gas-turbine combustion chamber having a head plate and an outer and an inner combustion chamber wall, characterized in that the combustion chamber is designed in one piece by means of a DLD method, or is assembled from segments which are welded to one another and manufactured in one piece by means of a DLD method, as well as to a method for manufacturing the gas-turbine combustion chamber.

Abstract: A cooling arrangement (56) having: a duct (30) configured to receive hot gases (16) from a combustor; and a flow sleeve (50) surrounding the duct and defining a cooling plenum (52) there between, wherein the flow sleeve is configured to form impingement cooling jets (70) emanating from dimples (82) in the flow sleeve effective to predominately cool the duct in an impingement cooling zone (60), and wherein the flow sleeve defines a convection cooling zone (64) effective to cool the duct solely via a cross-flow (76), the cross-flow comprising cooling fluid (72) exhausting from the impingement cooling zone. In the impingement cooling zone an undimpled portion (84) of the flow sleeve tapers away from the duct as the undimpled portion nears the convection cooling zone. The flow sleeve is configured to effect a greater velocity of the cross-flow in the convection cooling zone than in the impingement cooling zone.

Abstract: According to one aspect of the invention, a gas turbine engine includes a combustor, a fuel nozzle placed in an end of the combustor, and a passage configured to receive an air flow from a compressor discharge casing, wherein the passage directs the air flow into a chamber downstream of the nozzle, wherein a chamber pressure is lower than a compressor discharge casing pressure. The gas turbine engine also includes a flow control device configured to control the air flow from the compressor discharge casing into the passage.

Abstract: A transition piece aft frame assembly is provided, and includes a transition piece aft frame and a heat shield. The transition piece aft frame has an aft face. At least a portion of the aft face is exposed to an exhaust gas stream. The heat shield is connected to the transition piece aft frame. The heat shield is oriented to generally deflect the exhaust gas stream away from the aft face of the transition piece aft frame.

Abstract: A gas turbine engine combustion chamber includes a first wall and a second wall. The second wall is arranged within and spaced from the first wall to define a cavity between the first wall and the second wall. The first wall has a plurality of impingement apertures extending there-through and the second wall has a plurality of effusion apertures extending there-through. The impingement apertures have a first diameter, a first pitch, and a first area. The effusion apertures have a second diameter, a second pitch, and a second area. The ratio of the first diameter to the second diameter is at least 3, the ratio of the first pitch to the second pitch is at least 4 and the ratio of the first area to the second area is at least 9. This arrangement increases the cooling performance of the effusion apertures in the second wall.

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 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 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 combustor dome heat shield and a louver are separately metal injection molded and then fused together to form a one-piece combustor heat shield.

Abstract: 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.

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: A system for supplying a working fluid to a combustor includes a fuel nozzle, a combustion chamber disposed downstream from the fuel nozzle, an inner flow sleeve that circumferentially surrounds the combustion chamber and a plurality of injectors circumferentially arranged around the inner flow sleeve. The plurality of injectors provide for fluid communication through the inner flow sleeve and into the combustion chamber downstream from the fuel nozzle. The system further includes an outer air shield that defines an injection air plenum that surrounds the plurality of injectors. An inlet passage extends through the outer air shield to define a flow path into the injection air plenum. An outer sleeve is slidingly engaged with the outer air shield. The outer sleeve has a first position that restricts flow through the inlet passage and a second position that increases flow through the inlet passage.

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: 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: A combustor assembly for a turbine engine includes a combustor liner, a flow sleeve and a baffle ring. The flow sleeve surrounds the combustor liner. An annulus is formed between the flow sleeve and the combustor liner. A plurality of row of cooling holes are formed in the flow sleeve. The baffle ring radially surrounds the combustor liner and is located in the annulus.

Abstract: A method of using a counterflow injection mechanism disposed in a combustion liner. The combustion liner including an outer casing comprising a plurality of openings formed along a length and configured as compressed air inlets, an inner casing having a closed rear end and a combustion outlet, and an air circulation path extending between and along the inner and outer casings. The counterflow injection mechanism includes a fuel-air injection mechanism having fuel and air passages extending through the air circulation path and leading to fuel and air injection openings disposed at an off-center position. The method includes injecting fuel and air from the injection mechanism toward the closed rear end of the inner casing in a direction counterflow to a generally lengthwise downstream flow of combustion products in a gas turbine combustor.

Abstract: A combustor includes: a combustor basket configured to surround a fuel nozzle from an outer circumferential side, and a plurality of connecting members formed in a circumferential direction at intervals to connect a rear end of the combustor basket and a casing and configured to define a flow path through which compressed air introduced into the combustor basket flows. A circulation direction of the compressed air flowing through the flow path is configured to be reversed at a rear end of the combustor basket and the compressed air is introduced into the fuel nozzle. The flow path is partially or entirely inclined in the circumferential direction to blow the compressed air.

Abstract: A combustion arrangement for a turbine is provided including: a casing; a combustion chamber within the casing, an inner casing volume may be defined as a volume inside the casing but outside the combustion chamber; a partitioning wall partitioning the inner casing volume into first and second volume portions, the partitioning wall may have at least one aperture to allow fluid communication between the first and second volume portions; and a valve may be arranged at the casing to allow an outgoing fluid flow from the inner casing volume to an outside of the casing depending on a valve operating position. The combustion chamber has a combustion entry port for supplying an oxidant into the combustion chamber, where the combustion entry port may be in fluid communication with the first volume portion. The arrangement may be adapted to adjust the valve operating position for damping an oscillation of the arrangement.

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 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 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 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.

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: 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 system for supplying fuel to a combustor includes a combustion chamber, a liner that circumferentially surrounds at least a portion of the combustion chamber, and a flow sleeve that circumferentially surrounds at least a portion of the liner. A tube provides fluid communication for a working fluid to flow through the flow sleeve and the liner and into the combustion chamber, and the tube includes a tube wall. A plurality of fuel ports through the tube wall provide fluid communication for fuel to flow through the tube wall and into the tube, and the fuel ports are non-circular.

Abstract: A combustor of a turbine engine includes a portion where the aft end of a combustor liner is sealed to a forward end of a transition piece. A cover sleeve surrounds the aft end portion of the combustor liner to form an annular airflow passage between the exterior of the combustor liner and the inner side of the cover sleeve. A plurality of turbulators project radially outward from the combustor liner. One or more circumferential rows of supports also extend radially outward from the combustor liner to support the cover sleeve.

Abstract: A system for supplying a working fluid to a combustor includes a combustion chamber, a liner that circumferentially surrounds at least a portion of the combustion chamber, and a flow sleeve that circumferentially surrounds at least a portion of the liner. A tube provides fluid communication for the working fluid to flow through the flow sleeve and the liner and into the combustion chamber, and the tube spirals between the flow sleeve and the liner.

Abstract: A fuel nozzle for use with a turbine engine is provided. The fuel nozzle includes a housing coupled to a combustor liner defining a combustion chamber. The housing is at least partially positioned within an air plenum and comprises an endwall that at least partially defines the air plenum. The fuel nozzle includes a plurality of mixing tubes extending through the housing for channeling a fuel to the combustion chamber, a cooling fluid plenum at least partially defined within the housing by the housing endwall, and a plurality of apertures defined within the housing endwall for channeling a cooling fluid from the cooling fluid plenum to the air plenum.

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 gas turbine system comprising, a diffuser operative to diffuse an airstream output from a compressor, a fuel nozzle operative to receive fuel and emit the fuel in a combustor, and at least one bleed duct operative to direct bleed air from down stream of the combustor to the fuel nozzle.

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 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.

Abstract: In accordance with certain embodiments, a system includes a counterflow injection mechanism. The counterflow injection mechanism includes a fuel-air injection mechanism having fuel and air passages leading to fuel and air injection openings, wherein the fuel and air injection openings are disposed at an off-center position and a generally counterflow direction relative to a generally lengthwise flow axis of a gas turbine combustor.

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 combustor assembly that includes a casing that includes a plenum, a combustor liner positioned within the plenum and defining a combustion chamber there within, and a transition piece that includes a forward portion that is coupled to the combustion liner and an aft portion that extends from the forward portion. The aft portion includes a transition piece frame. The transition piece frame includes an upstream surface, a downstream surface, and a radially inner surface extending between the upstream and downstream surfaces. The radially inner surface includes a leading edge that extends between the radially inner surface and the upstream surface. A plurality of first cooling passages extend from the upstream surface to the downstream surface. Each of the first cooling passages is oriented obliquely with respect to the leading edge from the upstream surface towards the downstream surface.

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 combustor assembly for a turbine engine includes a combustor liner, a flow sleeve and a baffle ring. The flow sleeve surrounds the combustor liner. An annulus is formed between the flow sleeve and the combustor liner. A plurality of row of cooling holes are formed in the flow sleeve. The baffle ring radially surrounds the combustor liner and is located in the annulus.

Abstract: A resilient annular seal structure is disposed radially between an aft end portion of a combustor liner and a forward end portion of a transition piece, the resilient annular seal structure configured to form a first annular cavity radially between the forward end portion of the transition piece and the aft end portion of said combustor. At least one transfer tube radially extends from the second flow sleeve through the second flow annulus to the transition piece, and is arranged to supply compressor discharge cooling air radially from an area outside the first and second substantially axially extending flow annuli directly to the resilient annular seal structure and to the aft end 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.