Abstract: One exemplary embodiment of this disclosure relates to a component for a gas turbine engine. The component includes a baffle provided in an internal cavity of the component. The baffle includes a wall having an orifice therethrough, and the baffle further includes a lobe extending from the wall and at least partially covering the orifice.

Abstract: A thermal mechanical dimple array for combustor liners of gas turbine engines are located at known hot spots along liner to reduce stress and minimize cracking. Such arrays have indentations in the liners at the hot spots that act to thermal mechanically flex the liner at the hot spot away from the heat source and distribute stresses such that cracking of the liner is reduced and durability is improved.

Abstract: A combustor wall is provided for a turbine engine. The combustor wall includes a combustor shell and a combustor heat shield that is attached to the shell. The heat shield includes a first panel and a second panel that sealingly engages the first panel in an overlap joint. A cooling cavity extends between the shell and the heat shield and fluidly couples a plurality of apertures in the shell with a plurality of apertures in the heat shield.

Abstract: A combustor of a gas turbine engine including a combustor shell having an interior surface, a first panel mounted to the interior surface at a first position and a second panel mounted to the interior surface at a second position. The first panel has a first end, a first combustion chamber surface parallel with the interior surface, a first rail extending from the first combustion chamber surface toward the interior surface of the combustor shell, and a first extension extending axially from the first rail to the end of the first panel. The second panel has a second end, a second combustion chamber surface, and a second rail extending from the second combustion chamber surface toward the interior surface of the combustor shell. The first end and the second end are proximal to each other and define a circumferentially extending gap there between.

Abstract: The invention refers to a sequential combustor arrangement having a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the hot gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. The mixer includes of injection tubes pointing inwards from the side walls of the mixer for admixing the dilution gas to cool the hot flue gases leaving the first combustion chamber. A flow guide is arranged in the connecting duct and/or the injection tubes for guiding the dilution gas into the injection tubes. The invention further refers to a gas turbine and a method for operating a gas turbine with such a sequential combustor arrangement.

Abstract: A wall assembly is provided for a gas turbine engine. This wall assembly includes a support shell with a contoured region; and a multiple of liner panels mounted to the support shell. At least one of the multiple of liner panels includes an end rail. The contoured region is deformable to selectively contact at least a portion of the end rail. A method of assembling a wall assembly within a gas turbine engine is also provided. This method includes locating a stud that extends from a cold side of a liner panel through a support shell; and attaching a fastener onto the stud to at least partially close a gap defined between the panel and shell.

Abstract: A gas turbine combustor achieves improved product reliability and a reduced increase in pressure loss through improvements made on a cooling characteristic and structural intensity. The structure of the gas turbine combustor includes a plurality of circularity recesses formed on a side of an annular passage on a partial area of a combustion liner that requires cooling. Each of the circularity recesses has a rectangular surface forming a convex at a right angle with respect to a flowing direction of combustion air. The circularity recesses form a rectangular triangle having an oblique surface facing upstream of the flowing direction of the combustion air.

Abstract: The present invention relates to a combustion chamber tile of a gas turbine, where the combustion chamber tile has a plurality of effusion cooling holes and a mixing air hole provided as a flow-guiding element (chute) and projecting over the surface of the combustion chamber tile, characterized in that on the side of the mixing air hole facing away from a flow along the combustion chamber tile at least one supporting element is arranged on the combustion chamber tile and on the mixing air hole. It also relates to a method for its manufacture.

Abstract: A method includes providing an exhaust liner as a first duct, providing a second duct having an opening therein, and spacing the first duct radially inward of the second duct. A spring connects the first and second ducts together and has first and second ends. A first mount comprises a first bracket portion associated with the spring and a second bracket portion directly fixed to the first duct. The first bracket portion is directly fixed to the first end of the spring and the second end of the spring is fixed to a second mount that comprises a casing plate having an area greater than the opening. The first end of the spring and the first bracket portion are inserted through the opening and the first bracket portion is attached to the second bracket portion. The casing plate is directly fixed to a radially outward facing surface of the second duct.

Abstract: The present disclosure provides a nozzle for an aircraft turboprop engine with an unducted fan, including: an inner wall, an outer wall radially spaced apart from the inner wall and concentric with the inner wall, a junction area of the inner and outer walls including an opening contained in a plane transverse to a longitudinal axis of the nozzle. In particular, the junction area of the inner and outer walls includes two connecting plates and a member to secure the two connecting plates together, or in another form, the junction area includes a pad secured to the inner wall, and a pad secured to the outer wall, facing the pad of the inner wall of the nozzle.

Abstract: The present disclosure relates to combustor configurations, panels and components for a gas turbine engine. In one embodiment, a combustor for a gas turbine engine includes a support structure including a plurality of openings and a plurality of panels mounted to the structure. The plurality of panels define a combustion cavity of the combustor. Each panel includes a first wall configured to receive cooling air and a second wall configured to provide air flow for the cavity. The first and second walls form a cavity and include one or more elements for controlling the cooling effectiveness of each panel. Another embodiment is directed to a combustor panel including one or more elements for controlling cooling effectiveness. Another embodiment is directed to a support structure for a combustor of a gas turbine engine. Another embodiment is directed to configurations of panels including single walled portions or single walled panels.

Abstract: A combustor for a turbine engine is provided. A first liner has a first surface and a second surface. A second liner forms a combustion chamber with the second side of the first liner, and the combustion chamber 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 surface of the first liner. The plurality of effusion cooling holes includes a first effusion cooling hole extending from the first surface to the second surface and including an inlet portion extending from the first surface, a metering portion fluidly coupled to the inlet portion, and an outlet portion fluidly coupled to the metering portion and extending to the second surface. The inlet portion is larger than the metering portion.

Abstract: The present disclosure refers to an impingement cooling arrangement for cooling a duct wall of a duct guiding a hot gas flow. The impingement cooling arrangement includes an impingement sleeve which is at least partly disposed in a compressed air plenum, and spaced at a distance to the duct wall to form a cooling flow path between the duct wall and the impingement sleeve such that cooling air injected from the compressed air plenum through apertures in the sleeve impinges on the duct wall. At least one flow diverter is arranged in the cooling flow path to divert the cross flow away from at least one aperture. Besides the impingement cooling arrangement a gas turbine with such an arrangement as well as a method for cooling a duct wall are provided.

Abstract: A component within a gas turbine engine includes a surface with one or more upstream-directed cooling film holes therethrough.

Abstract: A gas turbine combustor part of a gas turbine includes a wall, containing a plurality of near wall cooling channels extending essentially parallel to each other in a first direction within the wall in close vicinity to the hot side and being arranged in at least one row extending in a second direction. The near wall cooling channels are each provided at one end with an inlet for the supply of cooling air, and on the other end with an outlet for the discharge of cooling air. The inlets open into a common feeding channel for cooling air supply, and the outlets open into a common discharge channel for cooling air discharge. The feeding channel and the discharge channel extend in the second direction.

Abstract: The invention refers to a sequential combustor arrangement comprising a first burner, a first combustion chamber, a mixer for admixing a dilution gas via a dilution gas inlet to the hot gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. The sequential combustor arrangement further includes four cooling zones with a cooling channel. During operation a cooling gas flows through the cooling channels. The disclosure further refers to a method for operating a gas turbine with such a sequential combustor arrangement.

Abstract: A silo combustion chamber for a gas turbine having a flame tube having an inner wall and flame tube base delimiting a combustion chamber; an outer wall surrounding the inner wall, forming a cavity; an annular chamber at least partly surrounding the outer wall and a number of supply lines fluidically coupled to the hot side of a heat exchanger during operation of the turbine; a number of burners, each of which opens into the combustion chamber on the outlet side through an opening in the flame tube base. The oxygen supply is fluidically connected to the annular chamber. A collecting chamber is arranged over the flame tube base. The annular chamber is connected to the collecting chamber via each connecting piece, each burner is fluidically connected to the collecting chamber to supply oxygen. The cavity between the inner and outer wall is locally fluidically separated from the collecting chamber.

Abstract: A combustor for use in a gas turbine engine includes a liner with a plurality of tiles coupled to a shell via a plurality of fasteners. Each of the plurality of fasteners extends through at least one of the tiles.

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 primary and intermediate regions of a combustion zone 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. A portion of the combustor air is introduced directly into the intermediate region.

Abstract: A combustor of a gas turbine engine includes a heat shield panel mounted to a support shell and an insert at least partially between opposed surfaces of the support shell and the heat shield panel, the insert exposed to a combustion chamber, and the insert being flush with a hot side of the heat shield.

Abstract: A combustion chamber comprises an outer wall and an inner wall spaced from the outer wall. The outer wall has at least one mounting aperture extending there-through and the inner wall has threaded studs extending there-from. The threaded studs extend through the mounting apertures in the outer wall. Cooperating nuts locate on the studs and washers are positioned between the outer wall and the cooperating nuts. Each washer has a rim and a bore. The washers have one or more passages extending there-through from the rim to the bore of the washer to provide a flow of coolant through the passages in the washers, the mounting apertures in the outer wall and around the threaded studs to cool the threaded studs to increase the working life of the inner wall. Other cooling arrangements for the threaded studs are disclosed.

Abstract: A combustion chamber louver assembly includes an aft louver having a forward panel 24 that extends axially from a louver leading edge 26 to a corner 28, and a forward louver joined to the forward panel of the aft louver. A lip 48 defined by a portion of the forward louver that extends axially past the corner to a louver trailing edge 36 includes circumferentially distributed trailing edge slots 60 extending forwardly from the trailing edge a nominal distance equal to about 88% to 95% of the length L of the lip.

Abstract: An object of the present invention is to provide a gas turbine combustor that can suppress an increase in pressure loss while improving product reliability. The gas turbine combustor includes a combustor liner, an air transfer casing installed on the outer circumference of the combustor liner, the combustor liner and the air transfer casing defining an annular passage therebetween adapted to allow a heat-transfer medium to flow therethrough, and a plurality of vortex generating devices disposed on an inside surface of the air transfer casing, the vortex generating devices generating longitudinal vortices each having a rotational axis extending in a flow direction of a heat-transfer medium. The plurality of vortex generating devices are arranged in paired manner, with each pair of devices generating vortices having rotational directions opposed to each other.

Abstract: Disclosed in various exemplary embodiments are turbine engine combustors with effusion and impingement cooling and methods for manufacturing the same. In one exemplary embodiment, disclosed is a combustor for a turbine engine that includes an annular liner portion including a first metering hole positioned on a cold side annular surface of the annular liner portion and an impingement chamber positioned in the annular liner. The impingement chamber connects to an entry hole on the cold side annular surface and includes a cooling air outlet passageway that is angled with respect to a hot side annular surface of the annular liner portion and that connects to an exit hole positioned on the hot side annular surface of the annular liner portion. The first metering hole is connected to the impingement chamber. The cooling air outlet passageway directs the air onto the hot side annular surface and spreads the airflow axially and laterally parallel to the hot side annular surface.

Abstract: A gas turbine combustion chamber has a housing (1), a flame tube (2) received therein, and a perforated plate (3) which at least partially surrounds the latter and which is fastened at both of its end faces (3.1, 3.2) to the housing.

Abstract: A turbomachine combustion chamber shell ring in which dilution holes in the turbomachine combustion chamber shell ring are covered with inserts defining chambers around same on an inner face of the shell ring. Ventilation holes, through the insert, induce ventilation of portions of the shell ring surrounding the dilution holes, cool the portions, and prevent crack formation.

Abstract: The present invention relates to a combustion chamber of a gas turbine having an outer combustion chamber wall and an inner combustion chamber wall, where the inner combustion chamber wall, at its front end area relative to the direction of flow through the combustion chamber, is fixed to the outer combustion chamber wall and, at its rear end area, is held longitudinally movable at the outer combustion chamber wall.

Abstract: A combustor liner is arcuate in shape and defines an axis and a circumferential direction. The combustor liner includes a first row of dilution openings and a second row of dilution openings. The first row runs in the circumferential direction. The second row runs parallel to the first row and is axially spaced from the first row. Each dilution opening of the second row overlaps in an axial direction a portion of each of two adjacent dilution openings of the first row.

Abstract: A shell for a combustor liner includes a cold side, a hot side, a row of cooling holes and a jet wall. The jet wall projects from the hot side for creating a wall shear jet of increased velocity cooling flow in a tangential direction away from the row of cooling holes and along an adjacent heat shield cold side wall.

Abstract: An exemplary burner arrangement and method for operating a burner arrangement are disclosed. During operation of the burner arrangement a hot combustion gas, including combustion air, flows essentially parallel to a burner wall through a mixing chamber, which is delimited by the burner wall, to a combustion chamber. In the mixing chamber the hot combustion gas is mixed with an injected fuel, where cooling air from the outside of the burner wall flows through effusion holes in the burner wall into an interior of the mixing chamber. The cooling air, on the outside of the burner wall, is deflected in a directed manner in its flow direction by means of deflection elements which are in a distributed arrangement.

Abstract: A turbomachinery component includes a surface exposed to hot working fluid flow. The surface has an undulating contour formed from a series of alternating protuberances and troughs. A set of three cooling outlets is associated with each trough.

Abstract: A ducting arrangement (10) for a can annular gas turbine engine, including: a duct (12, 14) disposed between a combustor (16) and a first row of turbine blades and defining a hot gas path (30) therein, the duct (12, 14) having raised geometric features (54) incorporated into an outer surface (80); and a flow sleeve (72) defining a cooling flow path (84) between an inner surface (78) of the flow sleeve (72) and the duct outer surface (80). After a cooling fluid (86) traverses a relatively upstream raised geometric feature (90), the inner surface (78) of the flow sleeve (72) is effective to direct the cooling fluid (86) toward a landing (94) separating the relatively upstream raised geometric feature (90) from a relatively downstream raised geometric feature (94).

Abstract: A combustor is provided for a turbine engine. The combustor includes a first liner having a first hot side and a first cold side; a second liner having a second hot side and a second cold side, the second hot side and the first hot side forming a combustion chamber therebetween. The combustion chamber is configured to receive an air-fuel mixture for combustion therein. The combustor further includes an insert having a body portion extending through the first liner and terminating at a tip, the body portion configured to direct air flow into the combustion chamber. The insert further includes a cooling hole defined in the body portion and configured to direct a first portion of the air flow toward the tip as cooling air.

Abstract: An impingement cooling sleeve includes a sleeve body having an inner surface to face a transition duct and an outer surface facing opposite the inner surface. At least one cooling hole is formed within the sleeve body and is used to direct cooling air toward the transition duct. At least one conduit member is attached to the sleeve body and is associated with the at least one cooling hole. The conduit member has a first opening to define an air inlet and a second opening to define an air outlet. In one example, the first opening is spaced apart from the outer surface of the sleeve body by a distance. In one example, the first opening comprises an annular end face surface that defines a plane that is obliquely orientated relative to the outer surface of the sleeve body.

Abstract: There is provided a gas turbine combustor that achieves improved product reliability and a reduced increase in pressure loss through improvements made on a cooling characteristic and structural intensity. A gas turbine combustor structure includes a plurality of circularity recesses 20 formed on a side of an annular passage 11 on a partial area of a combustion liner 8 that requires cooling. The circularity recesses 20 each have a rectangular surface 25 forming a convex at a right angle with respect to a flowing direction of combustion air 2. The circularity recesses 20 is a rectangular triangle having an oblique surface 26 facing upstream of the flowing direction of the combustion air 2.

Abstract: A gas-turbine combustion chamber has a cooling-air supply device delivering cooling air to a combustion chamber wall 6 to be cooled. The cooling-air supply device includes at least one air-supply duct 8 which, in respect of its flow axis, is arranged vertically to the combustion chamber wall 6, with the air-supply duct 8 being flow-connected to at least one cooling duct 9. The air-supply duct 8 forms a throttle and has a smaller diameter than the cooling duct 9. The cooling duct 9 issues at a shallow angle, relative to its flow axis, to the surface of the combustion chamber wall 6 in a recess 10 of the combustion chamber wall 6, with the cooling duct 9 being designed to impart a swirl to the air flowing through it.

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

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

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

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

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

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

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

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

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

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

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

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

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

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