Abstract: A combustor liner for a gas turbine includes a liner that at least partially defines a combustion chamber, and that a plurality of dilution openings therethrough. Each dilution opening includes an outer wall defining an outer perimeter of the dilution opening and defining a dilution opening centerline axis through the dilution opening. A plurality of swirl vanes extend from the outer wall into a dilution airflow passage that extends through the dilution opening. Each of the plurality of swirl vanes extends from the outer wall into the dilution airflow passage at a respective swirl vane angle with respect to the outer wall. The plurality of swirl vanes are arranged in a successive arrangement about the outer wall, and successive respective ones of the plurality of swirl vanes extend from the outer wall at a different swirl vane angle.

Abstract: A compressor diffuser for a gas turbine engine includes diffuser pipes having a tubular body including a generally radial portion, a bend portion and a generally axial portion. The generally axial portion has an exit segment extending parallel to the center axis and terminating at a pipe outlet. The bend portion is disposed radially further from the center axis than the exit segment.

Abstract: A gas turbine engine including a trapped vortex combustor assembly is generally provided. The combustor assembly includes an inner liner wall extended annularly around a combustor centerline, and an outer liner wall extended annularly around the combustor centerline. The inner liner wall and the outer liner wall together define an involute wall extended at least partially as a spiral curve from a circumferential reference line around the combustor centerline. The involute wall defines an involute combustion chamber. One or more of the inner liner wall and the outer liner wall each define a first dilution opening and a shaped dilution opening.

Abstract: A grommet may define a dilution hole in a combustor panel. The grommet may comprise a ridge having a stepped geometry formed about an inner diameter of the grommet, the ridge comprising a passage. The passage may comprise an outlet. The ridge may further comprise a fillet about the inner diameter of the grommet, wherein the outlet is configured to direct a cooling flow circumferentially along the fillet and fill the ridge with the cooling flow.

Abstract: An air-collecting structure effectively cools a transition piece of a duct assembly in a gas turbine combustor. The structure includes a flow sleeve having a plurality of cooling holes and surrounding the transition piece, the cooling holes formed in a lateral side of the flow sleeve to receive a compressed cooling air and arranged in rows running parallel to each other in a longitudinal direction of the flow sleeve, the rows progressing up the lateral side from a lower row to a higher row; and a plurality of scoops arranged in correspondence with predetermined cooling holes among the plurality of cooling holes and configured to collect an amount of air according to row. Each scoop includes an inlet having a predetermined radius for collecting the compressed cooling air. The radius is constant for the scoops of any one row and increases from the lower row to the higher row.

Abstract: A combustor component defining a dilution hole configuration for a combustor wall assembly of a gas turbine engine has first and second dilution holes that communicate radially with one-another and with respect to respective centerlines of each dilution hole. In operation, each dilution hole forms a respective jet stream along respective centerlines for flowing cooling air into a combustion chamber.

Abstract: A combustor wall assembly has a heat shield and a supporting shell with a cooling cavity defined therebetween. A grommet generally includes a wall defining a dilution hole isolated from the cooling cavity, and a flange projecting radially outward from the wall and into the cooling cavity. The flange is space from the heat shield and a cooling channel is defined between the wall and the heat shield that communicates with the cavity for cooling the wall proximate to a combustion chamber.

Abstract: The present disclosure generally relates to a fluid mixer, a system for mixing fluids utilizing the fluid mixer, and a method of mixing fluids using the fluid mixer or the system for mixing fluids, and more specifically, to a compact static mixing device with no moving parts and capable of mixing any fluid, such as air, nitrogen gas, water, oil, polluted water, and the like. A first pressurized, incoming fluid is accelerated locally by a section reduction, is split into streams, and then is released into a second fluid found in a closed volume or an open volume after a period of stabilization. The directed and controlled first fluid slides along an insert up to directional and angled fins at a vortex creator where suction forces from a self-initiating vortex in an internal cavity draws in at least part of the first fluid to fuel the vortex.

Abstract: A combustor heat shield has lips with fins distributed on the lips. The lip-fins have an extended end portion projecting rearwardly from the back face of the heat shield. Impingement jets may be directed against the rearwardly extended end portions of the lip-fins to enhance cooling. The heat shield may define a fuel nozzle opening surrounded by a rail on the back side of the heat shield. Impingement holes or slots may be defined in the rail for allowing cooling air passing therethrough to impinge upon the lip-fins.

Abstract: In a gas turbine combustor having an inner and outer liner defining an annular combustion chamber, at least an annular scoop ring provided on each inner and outer combustor liner. The annular scoop ring includes a solid radial inner base provided with bores defined therein and communicating with the combustion chamber to form air dilution inlets. The scoop ring has a radial outer portion in the form of a C-shaped scoop open to receive high velocity annular air flow. The bores of the inlets communicating with the scoop portion to direct the air flow into the combustion chamber whereby the bores of the inlets form jet nozzles to generate air jet penetration and direction within the combustion chamber.

Abstract: Methods and systems are provided for transferring heat from a transition nozzle. The transition nozzle includes a transition portion, a nozzle portion integrally formed with the transition portion, and at least one surface feature configured to transfer heat away from the transition portion and/or the nozzle portion. The transition portion is oriented to channel the combustion gases towards the nozzle portion.

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 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: 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: 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 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 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: 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 gas turbine combustor liner has at least one circumferential row of air holes adapted to supply air in a radial direction into a combustion chamber within the liner. One or more of the air holes have a thimble fixed therein, the thimble having a substantially circular body and a pair of lips extending from an interior end of the thimble in diametrically opposed upstream and downstream directions.

Abstract: A combustor for a gas turbine includes a plurality of nozzles provided in an array; a baffle plate configured to provide a desired air flow distribution to the array of nozzles; and a casing comprising a plurality of holes in an outer surface. The casing extends from a headend of the combustor to the baffle plate. A method of distributing an air flow in a combustor of a gas turbine includes providing an air flow to the outer surface of the casing; directing the air flow around the baffle plate; and distributing the air flow through the baffle plate to the array of nozzles.

Abstract: A method for assembling a gas turbine engine includes coupling a transition piece between a combustor liner and a nozzle assembly. The method also includes extending a first portion of a flow sleeve from the transition piece about at least a portion of the combustor liner. The method further includes coupling a second portion of the flow sleeve to the first portion of the flow sleeve such that the flow sleeve second portion extends from the flow sleeve first portion and at least partially about at least a portion of the transition piece. The flow sleeve second portion includes a scoop that cooperates with the transition piece to at least partially define a unitary cooling air passage that includes a unitary scoop-shaped opening. The scoop is oriented to introduce a substantially uniform cooling air flow to the transition piece.

Abstract: A gas turbine component in which combustion occurs. The gas turbine component includes a liner, including a first surface facing a first space and a second surface facing a second space, the liner being interposed between the first and second spaces and having a through-hole defined therein extending from the first to the second surface by which incoming flows proceed from the first space and to the second space. At least the first surface is formed to flow condition the incoming flows to resist separating from sidewalls of the through-hole.

Abstract: The present specification generally relates to improvements to combustors such as burners and engines. In one aspect, the specification presents an acoustically enhanced ejector system which can be used as part of an intake system for a combustor. In another aspect, the specification teaches the use of a combustor combustion chamber as an oscillator to magnify a harmonic frequency of a pulsating frequency of the combustor. In still other aspects, the specification presents a combustion chamber having an inlet with a plurality of tangentially spaced apertures, and an in-line intake system connected to the apertures.

Abstract: A scoop (54) over a coolant inlet hole (48) in an outer wall (40B) of a double-walled tubular structure (40A, 40B) of a gas turbine engine component (26, 28). The scoop redirects a coolant flow (37) into the hole. The leading edge (56, 58) of the scoop has a central projection (56) or tongue that overhangs the coolant inlet hole, and a curved undercut (58) on each side of the tongue between the tongue and a generally C-shaped or generally U-shaped attachment base (53) of the scoop. A partial scoop (62) may be cooperatively positioned with the scoop (54).

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 jet impingement array design and method are disclosed for efficiently cooling the liner of a gas turbine combustion chamber, while eliminating almost all effects of coolant gas crossflow. The design includes an array of extended jet ports for which the distance between the ends of the jet ports and the surface to be cooled progressively decreases from upstream to downstream. Spent air from upstream jets is directed away from downstream jets, thereby reducing the detrimental effects of crossflow, and optimizing heat transfer.

Abstract: A combustion duct for a gas turbine engine has a combustion liner to receive products of combustion, and deliver the products of combustion downstream toward a turbine rotor. An outer housing is positioned radially outwardly of the combustion liner. A flow sleeve is positioned radially intermediate the outer housing and the combustion liner. A chamber radially outwardly of the flow sleeve receives cooling air. A plurality of holes through the flow sleeve deliver cooling air from the chamber against an outer periphery of the combustion liner. A plurality of tabs are associated with at least some of the holes in the flow sleeve. The tabs are positioned to extend radially inwardly on a downstream side of the holes.

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 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 transition (200) for a gas turbine engine (100) comprises a transition wall (201) comprising cooling channels (213L, 213R, 223L, 223R) that are adapted to pass a cooling fluid, such as compressed air from a compressor (102) during operation of the gas turbine engine (100) so as to cool combusted hot gases passing through the transition (200). For each cooling channel (213L, 213R, 223L, 223R), respective entry ports (212L, 212L, 222L, 222R) and exit ports (214L, 214R, 224L, 224R) are arranged so as to obtain a performance improvement based upon pressure differentials between the respective entry and exit ports. In various embodiments, a scoop (220) is associated with an entry port (222L, 222R) so as to establish a more elevated pressure differential in the respective cooling channel (223L, 223R). An entry port (330a-h) may be positioned offset relative to a lower-positioned exit port (340a-h) so as to so minimize or eliminate intake of heated airflow from a respective nearby exit port (340a-h).

Abstract: A combustor assembly comprising a combustor liner having an opening and a combustor chute, which is formed of a sheet material which is bent to form the desired chute. The chute allows the passage of cooling and dilution air into the combustor.

Abstract: A system and method for reintroducing gas turbine combustion bypass flow. The system may include a combustor body, wherein the combustor body includes a reaction zone for primary combustion of fuel and air, and a casing enclosing the combustor body and defining an annular passageway for carrying compressor discharge air into the combustor body at one end. The system further may include a reintroduction manifold for receiving combustor bypass air extracted from the compressor discharge air in the annular passageway, and one or more reintroduction slots in communication with the reintroduction manifold for injecting the combustor bypass air into the combustor body downstream of the reaction zone. The method may include extracting combustor bypass air from the annular passageway, transporting the combustor bypass air to a reintroduction manifold, and reintroducing the combustor bypass air into the combustor body through one or more reintroduction slots in communication with the reintroduction manifold.

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 method and apparatus for cooling a combustor liner and transitions piece of a gas turbine include a combustor liner with a plurality of turbulators arranged in an array axially along a length defining a length of the combustor liner and located on an outer surface thereof; a first flow sleeve surrounding the combustor liner with a first flow annulus therebetween, the first flow sleeve having a plurality of rows of cooling holes formed about a circumference of the first flow sleeve for directing cooling air from the compressor discharge into the first flow annulus; a transition piece connected to the combustor liner and adapted to carry hot combustion gases to a stage of the turbine; a second flow sleeve surrounding the transition piece a second plurality of rows of cooling apertures for directing cooling air into a second flow annulus between the second flow sleeve and the transition piece; wherein the first plurality of cooling holes and second plurality of cooling apertures are each configured with an eff

Abstract: A combustion chamber for use in an engine, such as a rocket engine, has an inner wall with a corrugated configuration. The inner wall is formed with a plurality of tubes for carrying a coolant. As a result of the tubular construction, the inner wall has a plurality of spaced apart crowns and a plurality of valleys intermediate with the spaced apart crowns. To enhance the rate of heat flux or transfer in the valleys, a device is installed in each valley to cause local intensification of the turbulence in the hot gas flow in an effective way. The device may take the shape of a chevron or a bump.

Abstract: A method for providing cooling air to the venturi and the combustion chamber in a low NOx emission combustor as used in a gas turbine engine that includes the steps of providing an annular air passage surrounding said combustion chamber and venturi where said cooling air under pressure enters the combustion chamber/venturi near the aft portion of the combustion chamber, passing the air along the combustion chamber, past the venturi where the air exits near the front portion of the convergent area of the venturi. The method prevents any channel/passage cooling air from being received into the combustion chamber, and at the same time, introduces the outlet of the cooling air, after the air has passed over the combustion chamber of the venturi and has been heated, back into the premix chamber thereby improving the efficiency of the combustor while reducing and lowering NOx emission in the combustion process.

Abstract: A combustor liner for a gas turbine includes a substantially cylindrical body having a plurality of raised circular ribs arranged in an array on an outside surface of the combustor liner, each rib defining an enclosed are on the outside surface of the liner, forming a dimple or bowl that is sufficient to form vortices for fluid mixing in order to bring about heat transfer enhancement by both turbulated effect and dimpled effect.

Abstract: A combustor having a combustion liner and at least one combustor orifice assembly, the combustor orifice assembly comprising a boss, an orifice plate that defines an orifice, the orifice plate having a bottom surface that is adapted to be received by the boss, and a retaining ring, whereby the orifice plate is retained between the retaining ring and the boss.

Abstract: A transition piece assembly for a gas turbine includes a transition duct having one end adapted for connection to a gas turbine combustor and an opposite end adapted for connection to a first turbine stage, and a pair of side panels. The assembly also includes an impingement sleeve surrounding the transition duct and establishing a cooling path therebetween. The impingement sleeve is formed with a plurality of rows of cooling holes therein; and a plurality of flow catcher devices on an external surface of the impingement sleeve, each flow catcher device at least partially surrounding one of the cooling holes.

Abstract: An interchangeable combustor chute for directing airflow into a gas turbine engine combustor. In one aspect the combustor chute is secured to the combustor liner by mechanical fastener. The combustor chute is formed of a high temperature resistant material such as ceramics.

Abstract: Dilution sleeves are provided in radially aligned openings in the combustion liner and flow sleeve of a combustor for a gas turbine. The sleeves have a flange for seating on a collar of the flow sleeve about the opening therethrough enabling the dilution sleeve to extend through the aligned openings. A cover is removably mounted on the outer casing and a spring bears between the cover and the dilution sleeve flange to maintain the dilution sleeve in the openings. Dilution sleeves having different cross-sectional flow areas may be provided in lieu of initially installed dilution sleeves upon comparing measured emission levels with desired emission levels and ascertaining the desired increase or decrease in cross-sectional flow area through the dilution sleeves.

Abstract: Described is a plenum, in particular the air plenum of a gas-turbine combustion chamber, having at least two flow inlets, for introducing gas flows into the plenum, in which the gas flows are guided essentially along the inner wall of the plenum and are directed toward one another in such a way that the gas flows, after coinciding, are directed as a free gas flow away from the inner wall.The invention is distinguished by the fact that at least one flow obstacle (11) oriented in interaction with the direction of flow of the gas flows (1, 2) guided on the inner wall (3) is provided on the inner wall (3).

Abstract: A flow-guiding body is designed as a pointed, substantially conical molded shell. The projection of its base surface is formed by a straight line and by a curve that interconnects the ends of the straight line. The curve forms no significant angles. The molded shell faces with its point the fluid flow that hits its outer side and may be used as a mixing element for gaseous fuel and air, as an air sprayer with flame-holder, as a mixing element for admixed air in combustion chambers, as a swirling element or as a shell-shaped air sprayer combined with a fuel film generator or a fuel pressure spraying nozzle.

Abstract: This invention pertains to small air breathing gas turbine engines (1). Engines of the invention employ improved combustion systems, including improved fuel injection systems and improved combustor design. These improvements enable the engine to idle at speeds as low as one-tenth the maximum operating speed of the engine, which gives the engine an increased range of operating speeds. The improved designs generally provide for pre-mix assemblies intimate with the fuel injectors, a typically flame-free primary "A" mixing zone in the combustion chamber, and a primary "B" combustion zone, in addition to the secondary combustion zone and the dilution zone. In preferred embodiments, the bearing assembly is located between the compressor and the turbine wheel. A cooling cavity is provided between the turbine wheel and the bearing assembly, thus protecting the bearings from the turbine heat.

Abstract: A combustor in which fuel is introduced into a combustion zone at a first acute angle with respect to the radially outward direction so as to swirl the fuel about the combustor axis, thereby improving the stability of the combustion process. Combustion air is introduced into the combustion zone at a second acute angle with respect to the radially inward direction so as to swirl the air about the combustor axis in the same circumferential direction that the fuel is swirled, with the first and second acute angles being approximately equal. By swirling the combustion air in a manner that is compatible with the swirling of the fuel, noise associated with the combustion process is reduced.