Abstract: A combustor component of a gas turbine engine includes a substrate with a cold side, a central core section and a hot side, the cold side and the hot side have a porosity different than the central core section.

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 gas turbine combustor is provided with a combustor basket where combustion gas flows, the combustion gas being produced by combustion of fuel injected from a nozzle, and a first resonance device and a second resonance device mounted on an outer surface of the combustor basket. The second resonance device is disposed on a downstream side from the first resonance device in a flow of the combustion gas and damps combustion oscillation of a frequency higher than the first resonance device. The first and second resonance devices are acoustic liners each having a housing mounted to the outer surface of the combustor basket. A resonance space surrounded by the housing and the outer surface of the combustor basket communicates with an interior space of the combustor basket via a plurality of acoustic holes formed in the combustor basket.

Abstract: A design for an effectively cooling a liner of a gas turbine combustor by means of convective cooling is disclosed.

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: In one embodiment, a gas turbine engine component includes a foam based core and a composite skin member. Both the foam based core and the composite skin member can be used to structurally support the gas turbine engine component. The composite skin member can be a CMC material and is used to partially encapsulate the foam core. The gas turbine engine component can take the form of an airfoil member such as a blade or a vane, a combustor liner, etc. A first portion of the composite skin member includes a first surface extending past an edge of the component creating a step approximate an edge section. In another embodiment, composite skin members can be used to form a continuous shape for the edge section such that the foam core forms part of a gas path surface.

Abstract: In a gas turbine combustor and a gas turbine, an air passage (54) that supplies combustion high-pressure air, a pilot nozzle (44), a main fuel nozzle (45), and a top hat nozzle (47) that supply fuel, and openings (64) that supply film air (cooling air) are provided with respect to a combustor inner cylinder (42), and contraction members (71, 72, and 73) are arranged along a circumferential direction on an inner wall surface in a downstream of a flow direction of combustion gas in the inner cylinder (42). The contraction members (71, 72, and 73) are provided in a predetermined area in the circumferential direction excluding penetrating portions (74, 75, and 76), which do not disturb the flow of cooling air, thereby enabling to suppress generation of carbon monoxide and the like and to suppress the occurrence of unstable combustion.

Abstract: In one embodiment, a system includes a turbine combustor having a combustor liner disposed about a combustion chamber, a head end upstream of the combustion chamber relative to a downstream direction of a flow of combustion gases through the combustion chamber, a flow sleeve disposed at an offset about the combustor liner to define a passage, and a barrier within the passage. The head end is configured to direct an oxidant flow and a first fuel flow toward the combustion chamber. The passage is configured to direct a gas flow toward the head end and to direct a portion of the oxidant flow toward a turbine end of the turbine combustor. The gas flow includes a substantially inert gas. The barrier is configured to block the portion of the oxidant flow toward the turbine end and to block the gas flow toward the head end within the passage.

Abstract: A heat-shield element (14), for lining a combustion-chamber wall (13), the chamber wall has a first wall (17) with a hot side (18) which can be loaded with a hot medium, a cold side (19) which lies opposite the hot side (18), and a circumferential edge (24) which extends on a first (20), a second (21) and a third narrow side (22) of the first wall (17) to a first height (25) beyond the cold side (19), the circumferential edge (24) extends on a fourth narrow side (23) to a second shorter height (26), substantially at the second height (26), a second wall (27) lies opposite the cold side (19) and extends over the width of the fourth narrow side (23) from the fourth narrow side (23) and over a part of the length of the narrow sides (20, 22), which adjoin the fourth narrow side; (23), the second wall (27) has an edge (29) at the end (28) thereof which faces away from the fourth narrow side (23), which edge (29) extends to the first height (25). Furthermore, a combustion chamber and a gas turbine are disclosed.

Abstract: A gas turbine engine variable porosity combustor liner has a laminated alloy structure. The laminated alloy structure has combustion chamber facing holes on one side and cooling plenum facing holes on a radially opposite side. The combustion chamber facing holes are in fluid communication with the cooling plenum facing holes via axially and circumferentially extending flow passages sandwiched between metal alloy sheets of the laminated alloy structure. Porous zones having respective different cooling flow amounts are formed in the laminated allow structure based on at least one of an arrangement of the combustion chamber facing holes, an arrangement of the cooling plenum facing holes, and an arrangement of the flow passages.

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: An annular combustion chamber (1) for a gas turbine, the chamber having an outer shell (12) which has at least one inlet opening (4) for a burner (3) and an outlet (7) which opens into a turbine chamber (8), wherein ducts (14) which can be closed, which are oriented substantially parallel to the symmetry axis (2) of the annular combustion chamber (1), and through which final compressor air is guided into the annular combustion chamber (1) are provided in the outer shell (12) in the region of the outlet (7). Ring segments around the outer shell have projections selectively movable to at least partially block and open the ducts. A gas turbine is also disclosed.

Abstract: Described is an abradable component for a gas turbine engine, comprising: a base having an outboard side which receives a supply of cooling air in use and a plurality of walls on an inboard side thereof, the walls adjoining one another to provide an abradable network of open faced cells at a gas washed surface thereof; wherein at least one wall includes one or more through-holes for providing a flow of cooling air from the outboard side to the gas washed surface of the abradable network of open faced cells, when in use.

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 gas turbine engine component includes a cooling hole. The component includes a first wall having an inlet, a second wall having an outlet and a metering section extending downstream from the inlet and having a substantially convex first surface and a substantially concave second surface. The component also includes a diffusing section extending from the metering section to the outlet. A gas turbine engine wall includes first and second surfaces and a cooling hole extending between an inlet at the first surface and an outlet at the second surface. The cooling hole includes a metering section commencing at the inlet and a diffusing section in communication with the metering section and terminating at the outlet. The metering section includes a top portion having a first arcuate surface and a bottom portion having a second arcuate surface. The first and second arcuate surfaces have arcs extending in substantially similar directions.

Abstract: In a gas turbine engine (1) that has a plurality shrouds (8c and 9c), in which each one of the shrouds (8c and 9c) are provided with a groove portion (20) that is provided in a surface facing a turbine rotor blade (8a and 9a) and a plurality of film cooling holes (21) that open to a bottom portion of the groove portion (20), blocking of the film cooling holes (21 and 22) is prevented.

Abstract: A combustor for a gas turbine engine is provided. The combustor includes a forward bulkhead, an inner radial combustor wall, and an outer radial combustor wall. The bulkhead includes a plurality of circumferentially disposed injector apertures. The inner radial combustor wall includes a plurality of inner quench aperture sets. Each inner quench aperture set includes a first inner quench aperture and a second inner quench aperture separated from each other by an inner interset distance. Each inner quench aperture set is separated from an adjacent inner quench aperture set by an inner intraset distance. The inner intraset distance is different than the inner interset distance. The outer radial combustor wall includes a plurality of circumferentially disposed outer quench apertures. The outer radial combustor wall is disposed radially outside of the inner radial combustor wall, thereby defining an annular combustion region therebetween.

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: The present application and the resultant patent provide a combustor for mixing a flow of air and a flow of fuel. The combustor may include an air path for the flow of air, a number of fuel injectors positioned in the air path for the flow of fuel, and a crossfire tube positioned within the air path upstream of the fuel injectors. The crossfire tube may include a number of purge holes positioned on a downstream side thereof to reduce a wake in the flow of air caused by the crossfire tube in the air path.

Abstract: A combustor is provided for a turbine engine. The combustor includes a first liner having a first side and a second side and a second liner having a first side and a second side. The second side of the second liner forms a combustion chamber with the second side of the first liner, and the combustion chamber is configured to receive an air-fuel mixture for combustion therein. The first liner defines a plurality of effusion cooling holes configured to form a film of cooling air on the second side of the first liner. The plurality of effusion cooling holes including a first effusion cooling hole extending from the first side to the second side with a non-linear line of sight.

Abstract: A combustor assembly includes a hot skin of a combustion chamber wall having an inner face exposed to the combustion chamber and an opposite outer face, a receiving skin having a securing portion affixed to the hot skin outer face in an air-tight manner and a receiving flange, extending from the securing portion, that is offset from the hot skin outer face to form a female recess, a cold skin having a cold wall portion spaced from the hot skin and forming a cooling cavity therebetween, a securing portion extending from a first end of the cold wall portion affixed to the hot skin outer face in an air-tight manner and a male flange extending from a second end of the cold wall portion opposite the first end, the male flange snugly received in the female recess and forming a sliding engagement therebetween.

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: The invention relates to a combustion chamber for burning an air-fuel mixture, a gas mixture or a gas-liquid mixture, including an outer casing; a flame tube located inside the outer casing and provided with an inner space, whereby the outer casing and the flame tube form an annular channel therebetween, the flame tube including flow apertures which penetrate through its casing and interconnect the annular channel and the inner space of the flame tube; a cover part arranged in inlet ends of the outer casing and the flame tube and provided with flow apertures opening up into the annular channel; a combustion air inlet part which is arranged in an outer surface of the cover part and which is in connection with the flow apertures of the cover part; a fuel jet located in the cover part and extending inside the flame tube for feeding fuel into the flame tube.

Abstract: A process is provided for forming shaped air holes, such as for use in turbine blades. Aspects of the disclosure relate to forming shaped portions of air holes using a short pulse laser, forming a metered hole corresponding to each shaped portion, and separately finishing the shaped portion using a short-pulse laser. In other embodiments, the order of these operations may be varied, such as to form the shaped portions and to finish the shaped portions using the short-pulse laser prior to forming the corresponding metered holes.

Abstract: A method for the arrangement of effusion holes and impingement cooling holes in a combustion chamber wall including: distribution of the effusion holes in the surface to be cooled in accordance with pattern, diameter and dimension selections made; distribution of the impingement cooling holes in accordance with pattern, diameter and dimension selections made; checking of the number of matches and their spacing from one another, taking into account the component and assembly tolerances; comparison with the permitted number and their minimum spacing; and, if quality requirements are not met, taking corrective actions, including selecting alternative diameters and patterns.

Abstract: A low calorific value fuel-fired can combustor for a gas turbine include a generally cylindrical housing, and a generally cylindrical liner disposed coaxially within the housing to define with the housing a radial outer flow passage for combustion air, the liner also defining inner combustion and a dilution zone, the dilution zone being axially distant a closed housing end relative to the combustion zone. A nozzle assembly disposed at the closed housing end includes an air blast nozzle and surrounding swirl vanes. An impingement cooling sleeve coaxially disposed in the combustion air passage between the housing and the liner impingement cools the portion of the liner defining the combustion zone. The combustion liner has an L/D ratio of in the range 1?L/D?4, and a ratio of the combustion zone volume (m3) to heat energy flow rate Q (MJ/sec) in the range 0.0026?V/Q?0.018.

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 gas turbine engine variable porosity combustor liner has a laminated alloy structure. The laminated alloy structure has combustion chamber facing holes on one side and cooling plenum facing holes on a radially opposite side. The combustion chamber facing holes are in fluid communication with the cooling plenum facing holes via axially and circumferentially extending flow passages sandwiched between metal alloy sheets of the laminated alloy structure. Porous zones having respective different cooling flow amounts are formed in the laminated alloy structure based on at least one of an arrangement of the combustion chamber facing holes, an arrangement of the cooling plenum facing holes, and an arrangement of the flow passages.

Abstract: A transition piece of a combustor that sends high temperature combustion gas to a turbine includes a cylindrical body wall and cooling air passages. The passages are formed in the body wall so as to extend in an axial direction thereof, and each of the passages has cooling air inlet ports opened at an outer circumferential surface of the transition piece and cooling air outlet ports opened at an inner circumferential surface of the transition piece. The cooling air outlet ports form a plurality of lines in a direction oblique with respect to the axial direction of the body wall. A first distance between a first line of the cooling air outlet ports and a second line of the cooling air outlet ports adjacent to the first line is larger than a second distance between the cooling air outlet ports adjacent to each other.

Abstract: A gas turbine engine includes a combustor having a dual-wall impingement convention effusion combustor tile assembly. The dual-wall tile assembly provides a cooling air flow channel and attachments for securing the tile to the cold skin liner of the combustor. Cooling is more efficient in part due to the dual wall construction and in part due to reduced parasitic leakage, and the design is less sensitive to attachment features which operate at lower temperatures.

Abstract: A combustion arrangement includes a combustion section. Also included is an air discharge section downstream of the combustion section. Further included is a transition region disposed between the combustion section and the air discharge section. Yet further included is a transition piece defining the combustion section and the transition region, wherein the transition piece is configured to carry a combusted gas flow from the combustion section to the air discharge section. Also included is a damping device operatively coupled to the transition piece proximate the air discharge section.

Abstract: The invention is a combustor liner (12) of a dual wall cooling structure including an inner wall section (30) configured to surround a combustion region (13) and in which a plurality of effusion cooling holes (31) are formed, and an outer wall section (20) formed to be spaced apart from the inner wall section (30) and in which a plurality of impingement cooling holes (21) are formed, wherein the inner wall section (30) is constituted by a plurality of plate-shaped members (40), and a support guide member (50) is provided which is configured to guide the plurality of plate-shaped members (40) to enable free insertion and extraction and support the plurality of plate-shaped members (40) at intervals such that deformation by thermal expansion is able to be absorbed.

Abstract: The present invention relates to a gas-turbine combustion chamber having a combustion chamber wall including a tile carrier, on which wall tiles are mounted at a distance to form an impingement cooling gap, where the tile carrier has impingement cooling holes and the tile is provided with effusion holes, where the tile has on its side facing the tile carrier a surface structure which raises from the surface of the tile and extends in the direction of the tile carrier.

Abstract: A combustor in a gas turbine includes a liner having an interior volume defining a main combustion zone, a fuel injection system for delivering fuel into the main combustion zone, and a flow sleeve that defines, with the liner, a passageway for air to flow on its way to be mixed with fuel from the fuel injection system, wherein the mixture is burned in the main combustion zone to create hot combustion gases. The combustor further includes a flow conditioner including at least one panel having a configuration such that air is able to pass through the panel(s) on its way to the passageway, wherein at least a substantial portion of the air that enters the passageway for being burned in the main combustion zone passes through the panel(s).

Abstract: A method of assembling a combustor assembly for use in a turbine engine is provided. A combustor liner that defines a combustion chamber therein is provided. The liner includes a forward end and an aft end, wherein the aft end includes at least one channel extending therethrough. The channel is aligned obliquely with respect to a centerline extending through the aft end. A plurality of fuel nozzles are coupled to the forward end such that the fuel nozzles extend through the forward end. An annular sleeve that includes at least one opening extending radially therethrough is coupled to the aft end, wherein the sleeve substantially circumscribes the aft end such that a cavity is defined therebetween and such that a fluid channeled through the at least one opening impinges against a surface of the aft end prior to the fluid being channeled into the combustion chamber.

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: Combustor liners made using additive manufacturing techniques can employ cooling hole patterns which are not possible, or at least time consuming or expensive, to make using traditional subtractive manufacturing techniques. By additively manufacturing floatwall panels, cooling holes may be placed along axes that transect features on the floatwall panel, such as mounting studs, spacers, cooling pedestals, and rails.

Abstract: A heat shield for a gas turbine engine includes a hot side with one or more raised features that extend therefrom.

Abstract: A combustor nozzle includes a downstream surface having an axial centerline. A plurality of passages extend through the downstream surface and provide fluid communication through the downstream surface. A plurality of slits are included in the downstream surface, and each slit connects to at least two passages. A method for modifying a combustor nozzle includes machining a plurality of slits in a downstream side of a body. The method further includes connecting each slit to at least two passages that pass through the body.

Abstract: A cooling arrangement for a surface of a wall in a gas turbine engine, the wall having a plurality of effusion holes each with an outlet onto the surface for supplying an effusion flow to the surface and an inlet, the inlets of the effusion holes being arranged at the peripheries of groups tessellated on an opposing surface of the wall, each inlet being located on the peripheries of three groups. The arrangement comprises a second wall spaced apart from the opposing surface having impingement orifices each for directing a flow of air in use to a respective impingement location on the opposing surface, each group having a centrally positioned impingement location.

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: An annular wall for a combustion chamber of a turbomachine. The wall presents a hot side and a cold side and includes at least one primary hole for enabling a first flow of air flowing on the cold side of the wall to penetrate to the hot side of the wall to feed combustion of fuel inside the combustion chamber, and together with a plurality of cooling holes, each having a diameter no greater than 1 mm, to enable a second flow of air flowing on the cold side of the wall to penetrate to the hot side of the wall to cool the hot side of the wall. The plurality of cooling holes can also dilute combustion gas resulting from the combustion by using the flow of air penetrating to the hot side of the wall through the cooling holes.

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: A method for filling cooling holes in a component of a gas turbine engine is disclosed. The component may include a plurality of first cooling holes extending through the wall of the component. The method may comprise the steps of exposing the outer surface of the component, filling the plurality of first cooling holes with a polyimide, curing the polyimide to block the passage of cooling fluid through the plurality of first cooling holes, and applying a thermal bather coating over the outer surface of the component. The method may further include the step of installing a second plurality of cooling holes in the wall of the component wherein the plurality of second cooling holes penetrate the thermal barrier coating and the wall of the component and allow cooling fluid to pass therethrough.

Abstract: A gas turbine engine has a combustor module including an annular combustor having a liner assembly that defines an annular combustion chamber and includes a circumferential row of a plurality of relatively large combustion dilution air admission holes and a circumferential row of a plurality of smaller quench air admission holes disposed downstream with respect to the flow of combustion gas products. The plurality of quench air admission holes are arranged with respect to the plurality of relatively large dilution air admission holes disposed upstream thereof such that there is associated with each dilution air admission hole a first quench air admission hole and a second quench air hole, the first quench air hole being offset laterally in a first lateral direction and the second quench air hole being offset laterally in a second lateral direction opposite to the first direction.

Abstract: A transition piece of a combustor that sends high temperature combustion gas to a turbine includes a cylindrical body wall and cooling air passages. The passages are formed in the body wall so as to extend in an axial direction thereof, and each of the passages has cooling air inlet ports opened at an outer circumferential surface of the transition piece and cooling air outlet ports opened at an inner circumferential surface of the transition piece. The cooling air outlet ports form a plurality of lines in a direction oblique with respect to the axial direction of the body wall. A first distance between a first line of the cooling air outlet ports and a second line of the cooling air outlet ports adjacent to the first line is larger than a second distance between the cooling air outlet ports adjacent to each other.

Abstract: A system for cooling a hot gas path component for a combustor generally includes an impingement sleeve that circumferentially surrounds an outer surface of the hot gas path component. A first cooling chamber is defined between the impingement sleeve and a first portion of the outer surface of the hot gas path component. A second cooling chamber is disposed downstream from the first cooling chamber. The second cooling chamber is defined between the impingement sleeve and a second portion of the outer surface of hot gas path component. An inlet extends through the impingement sleeve so as to define a first flow path into the first cooling chamber. An outlet defines a second flow path between the first cooling chamber and the second cooling chamber.

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.