Abstract: A combustor of a gas turbine engine including a first forward liner panel. The forward liner panel including a liner panel forward edge, a liner panel aft edge opposite of the liner panel forward edge, and a first liner panel side edge between the liner panel aft edge and the liner panel forward edge. The first liner panel side edge is non-perpendicular to the liner panel forward edge and oriented at an angle about perpendicular to a swirl flow in the combustor.

Abstract: An igniter seal arrangement for a combustion chamber. The combustion chamber has a wall having first surface and second surfaces. A boss projects from the first surface, and has a platform on its remote end. The platform has an inner surface spaced from the first surface of the wall to define a chamber between the first surface of the wall and the inner surface of the platform. The platform has an outer surface facing away from the first surface of the wall and an aperture extends through the wall from the outer surface of the platform of the boss to the second surface. First and second L-shape rails extend from the platform and a sealing member has a first edge and a second edge locatable between the outer surface of the platform and the first and second L-shape rails and the sealing member has an aperture to receive an igniter.

Abstract: A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A plurality of vane cooling passages extending through a forward flange of the first vane stage to the radial inner surface for communicating cooling airflow into the core flow path forward of the leading edge. A seal assembly is disposed between the combustor and the forward flange of the vane stage. The seal includes a plurality of circumferentially spaced apart slots and a plurality of seal cooling passages that extend from the radially outer surface to the radially inner surface at each of the plurality of circumferentially spaced apart slots.

Abstract: A method of manufacturing a heat shield panel for a gas turbine engine comprising: determining a plurality of layers to compose a heat shield panel having one or more cooling apertures with complex geometries; forming each of the plurality of layers into sheets; forming each of the sheets into a three-dimensional slice of the heat shield panel; stacking each of the sheets to form a three-dimensional heat shield panel; and joining the sheets together.

Abstract: A combustor liner assembly is provided including an inlet wall, an inner wall, an outer wall, and a plurality of fasteners. The inlet wall has an opening into a combustion chamber. The inner wall is coupled to the inlet wall at a first end of the combustor liner. The inner wall defines a radially inner end of the combustion chamber. The outer wall includes a plurality of segments. The plurality of segments define a radially outer end of the combustion chamber. A first portion of the plurality of fasteners couples each of the plurality of segments to another of the plurality of segments. A second portion of the plurality of fasteners couples at least one of the plurality of segments to the inlet wall at the first end of the combustor liner.

Abstract: A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a first plurality of openings and the first vane stage includes a second plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A first vane stage assembly and a method are also disclosed.

Abstract: A combustion section of a gas turbine engine may include a cassette configured to couple to an annular combustor dome arranged around a flow path for the gas turbine engine. The annular combustor dome may include a dome wall comprising a plurality of inlets configured to receive compressed air for a combustion chamber located downstream, relative to the flow path, from the annular combustor dome. The cassette may include a combustor wall extended away from the dome wall in a downstream direction. The cassette may further include a cowl integral to the combustor wall. At least a portion of the cowl may extend away from the dome wall in an upstream direction relative to the flow path. The cowl may receive compressed air from a diffusor and guide the compressed air to a space upstream from the inlets, relative to the flow path.

Abstract: A combustor cap assembly includes a cap plate and a cover panel. The cap plate defines an outer perimeter and fuel nozzle openings disposed through the cap plate. The cap plate has a hot side, and the cover panel has a cold side, which is attached to the hot side of the cap plate. The hot side of the cap plate and/or the cold side of the cover panel define a plurality of cooling microchannels extending in a transverse direction relative to a centerline of the combustor cap assembly. A first microchannel has a first channel outlet disposed along the outer perimeter of the cap plate, and/or a second microchannel has a second channel outlet disposed within one of the plurality of fuel nozzle openings. The cover panel may be coated with a bond coat and a thermal barrier coating.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a combustor wall. The combustor wall includes a shell, a heat shield and an annular body. The body extends laterally between an inner surface and an outer surface. The inner surface defines an igniter aperture in the combustor wall. The outer surface is vertically between the heat shield and the shell. The shell defines a first cooling aperture through which air is directed to impinge against the outer surface.

Abstract: A combustor reduces combustion vibration and the generation of nitrogen oxides and includes a burner to mix fuel with compressed air and to eject the mixture; a liner that is coupled to one end of the burner and forms a combustion chamber in which the fuel-air mixture is combusted to produce high-pressure gas; a transition piece to receive a flow of the high-pressure gas from the liner and to discharge the high-pressure gas to a turbine; a plurality of resonance modules disposed on an outer circumferential surface of the liner at positions spaced apart from each other along a circumferential direction of the liner; and a fuel supply unit coupled to each of the plurality of resonance modules and configured to supply fuel to each of the plurality of resonance modules. The resonance modules may be disposed on an outer circumferential surface of the transition piece instead of the liner.

Abstract: An electronic device is disclosed. The electronic device may include a first part coupled to a second part by a fastener. The first part may include a through hole and several spring elements that define the first through hole. The second part may include a through hole and several extensions that define the through hole. The electronic device may include a fastener used to secure the first part with the second part. In this manner, the fastener passes through respective through holes of the first part and the second part. Further, the fastener can engage and deform the spring elements such that the spring elements extend into the through hole of the second part. As a result, the spring elements are displaced and the fastener engages not only the spring elements but also the extensions. The parts may include circuit boards or sheet metal, as non-limiting examples.

Abstract: A method of manufacturing a heat shield panel assembly is provided. The method including: forming a heat shield panel, wherein the heat shield panel includes one or more orifices; and forming each of one or more threaded studs through operations including: injecting melted wax into a negative cavity of a threaded stud; allowing the wax to solidify to form a positive pattern of the threaded stud; removing the positive pattern of the threaded stud from the negative cavity of the threaded stud; coating the positive pattern of the threaded stud with a ceramic; melting the positive pattern of the threaded stud away from the ceramic, the ceramic having a second cavity forming a second negative cavity of the threaded stud; pouring melted metal into the second cavity; allowing metal in the second cavity to solidify to form the threaded stud; and removing the ceramic from the threaded stud.

Abstract: A combustor for a gas turbine engine including a liner panel mounted to a support shell via a multiple of studs, the liner panel including a forward section and an aft section that defines an arcuate surface section therebetween in the axial profile between the forward section and the aft section.

Abstract: A combustor for a gas turbine engine including a liner panel mounted to a support shell via a multiple of studs, the liner panel including a forward section and an aft section that defines an angle therebetween in the axial profile between the forward section and the aft section.

Abstract: The described reverse flow combustor of a gas turbine engine includes inner and outer combustor liners defining a combustor chamber therewithin. A large exit duct and a small exit duct are disposed at downstream ends of the outer and inner liner respectively. The small exit duct includes an annular ring removably mounted to a support element of the gas turbine engine by one or more fastening elements that are integrally formed with the annular ring. The fastening elements are mountable with corresponding features of the support element and removably fastened thereto.

Abstract: A gas turbine engine includes a high pressure compressor disposed about a central longitudinal axis, a combustor, and a diffuser case. The diffuser case includes an outer ring providing a first platform and an inner ring providing a second platform. The first platform and the second platform are axially between the high pressure compressor and the combustor with respect to the central longitudinal axis. A plurality of circumferentially spaced struts extend radially from the first platform to the second platform and each include first and second circumferential sides. A plurality of heat shields are disposed on a leading edge defined at a forward end of a respective one of the plurality of struts.

Abstract: An improved system, apparatus and method may comprise an inner liner and an outer liner extending circumferentially around an engine centerline axis thereby forming a combustion chamber therebetween. The system, apparatus, and method may include a front end extending between the inner and outer liners and having a plurality of fuel nozzle ports configured to each receive a fuel nozzle. First and second outer plunged inlet holes may be formed in the outer liner and extending in a first radial direction toward the engine centerline axis and into at least a portion of the combustion chamber. First and second inner plunged inlet holes may be formed in the inner liner and extending in a second radial direction away from the engine centerline axis and into at least a portion of the combustion chamber.

Abstract: A gas turbine engine combustor has a shell defining a combustion chamber having a primary zone and a dilution zone downstream of the primary zone. The shell has an outer skin and an inner skin defining an air gap therebetween. The inner skin in the upstream portion of the primary zone is free from effusion holes and has heat transfer augmenters projecting from a back side thereof into the air gap. Effusion holes in the inner skin are only disposed downstream towards the tail end of the primary zone in order to delay the injection of any disruptive cooling air that could potentially interfere with the combustion process.

Abstract: A combustion chamber comprises an upstream end wall structure and inner and outer annular wall structures. The upstream end wall structure comprises an upstream wall and a plurality of circumferentially arranged heat shields secured to the upstream wall. The upstream wall has a plurality of circumferentially spaced fuel injector apertures. Each heat shield has radially outer and radially inner ends and a fuel injector aperture aligned with a corresponding fuel injector aperture in the upstream wall. The radially outer and inner ends of each heat shield have outer and inner rails spacing the heat shield from the upstream wall. The radially outer and inner ends of each heat shield have first and second pluralities of circumferentially spaced apertures extending there-through and through the associated outer and inner rails to direct coolant over the surface of the outer and inner annular wall structures to form respective films of coolant.

Abstract: A gas turbine engine includes a cover, a support, and a sealing member. The cover has a cover housing and a cover arm extending from the cover housing between a first arm end connected to the cover housing and a second arm end having a cover end surface. The support has a support body, a first support arm extending between a first support arm first end connected to the support body and a first support arm second end. The cover arm extends into the first support arm. The sealing member at least partially extends across a gap defined between the cover end surface and the first support arm.

Abstract: A combustor for a gas turbine engine includes an inner combustor wall and an outer combustor wall radially outboard of the inner combustor wall. The inner and outer combustor walls define a combustion space therebetween with an upstream inlet and a downstream outlet. A combustor dome connects between inner and outer combustor walls at the upstream inlet of the combustion space. The combustor dome includes a plurality of circumferentially spaced apart nozzles and a plurality of tiles mounted to the combustor dome to fluidly separate a downstream side of the combustor dome from an upstream side of the combustor dome.

Abstract: A casing for a combustor includes an upstream segment, a downstream segment and a joint. The downstream segment is constructed and arranged to move telescopically with respect to the upstream segment and along a centerline. The joint includes a flange and bonding material, which attaches the flange to the upstream segment or the downstream segment. A melting point of the bonding material is less than a melting point of the flange and/or a melting point of the segments.

Abstract: A combustor liner, a gas turbine engine including a combustor having the combustor liner, and a method for regulating jet wakes in the combustor are disclosed. The combustor liner includes a panel, at least one first through-hole, and a plurality of second through-holes including first and second set of through-holes. The at least one first through-hole is disposed on a section of the panel in a first row. The plurality of second through-holes is disposed on the section along axial and circumferential directions and arranged adjacent to the at least one first through-hole. The first set and the second set of through-holes are arranged in a second row and a third row respectively. The first, second, and third rows extend along the circumferential direction. The at least one first through-hole and the plurality of second through-holes collectively cover circumferential plane of the section along the first, second, and third rows.

Abstract: The present invention is a combustor provided with a combustor transition pipe connected to a turbine while interposing a transition pipe seal in between, including a flange portion provided at an end portion on a downstream side in a fluid flow direction of the combustor transition pipe, the flange portion projecting to radially inside and extending in a circumferential direction. The flange portion includes a pin hole into which a pin to position the transition pipe seal is inserted, a circumferential slit portion either extending within a range in a radial direction where the pin hole is formed or being located on radially outside of the pin hole and extending in the circumferential direction, and a hole portion on which part of the circumferential slit portion abuts.

Abstract: A liner panel within a gas turbine engine includes a perimeter rail that defines a first height from a cold side. The liner panel also includes an intermediate rail that defines a second height from the cold side, where the second height is less than the first height.

Abstract: A combustor for use in a gas turbine engine has a combustor outer shell. A panel has an inner face which will face hot products of combustion, and a boss surrounding a feature, with the boss extending to an outer end. A spacing surface is spaced from the boss, and is at an outer position that is inward of the outer end of the boss. The spacing surface spaces the panel from the outer shell. A trough is intermediate the boss and the spacing surface. The trough extends to an outer end which is inward of the outer position of the spacing surface. A gas turbine engine is also disclosed.

Abstract: A heat shield panel for a gas turbine engine combustor may comprise a perimetrical rail extending at least partially around a perimeter of the heat shield panel and defining a main backside cavity, a plurality of air admittance holes, each air admittance hole extending along a hole axis, and a midrail dividing the main backside cavity into a first backside cavity and a second backside cavity, wherein at least a portion of the midrail is axially offset from at least one of the hole axes.

Abstract: Combustors of gas turbine engines having a combustor shell having a first end and a second end opposite the first end, a first securing element positioned at the first end of the combustor shell, a second securing element positioned at the second end of the combustor shell, a plurality of high temperature material panels fixedly secured by the first securing element at the first end and the second securing element at the second, wherein a panel gap is formed between edges of adjacent high temperature material panels of the plurality of high temperature material panels, and a seal divider extending from the first end to the second end and positioned on the combustor shell and arranged to seal the panel gap between adjacent first and second high temperature material panels.

Abstract: A cooling mechanism includes a bottom wall (22) in contact with a combustion chamber, an upper wall (30), and a cooling passage (40) arranged between the bottom wall (22) and the upper wall (30). The cooling passage (40) includes a first passage (50) extending to a first direction, a second passage (60) extending to the first direction, and a connection section (70) connected with the first passage (50) and the second passage (60). The second passage (60) is arranged to have an offset to the first passage (50) in a second direction perpendicular to the first direction and along the bottom wall (22).

Abstract: A casing for use in a turbofan engine is provided. The casing includes an inner hub, an intermediate casing positioned radially outward from the inner hub, an outer casing positioned radially outward from the intermediate casing, and a plurality of struts spaced circumferentially about the inner hub and extending between the inner hub and the outer casing. At least one strut of the plurality of struts includes a flow channel extending therethrough. The casing also includes a fluid scoop defined within the inner hub, and a collector that couples the fluid scoop in flow communication with the flow channel.

Abstract: A hot gas path (HGP) component of an industrial machine includes primary and secondary cooling pathways. A body includes an internal cooling circuit carrying a cooling medium. A primary cooling pathway is spaced internally in the body and carries a primary flow of a cooling medium from an internal cooling circuit. A secondary cooling pathway is in the body and in fluid communication with an internal cooling circuit. The secondary cooling pathway is fluidly incommunicative and spaced internally from the primary cooling pathway. In response to an overheating event occurring, the secondary cooling pathway opens to allow a secondary flow of cooling medium through to the outer surface of the body and/or the primary cooling pathway. The primary flow flows in the primary cooling pathway prior to the overheating event, and the secondary flow of cooling medium does not flow until after an opening of the secondary cooling pathway.

Abstract: A combustor includes a combustor shell, an inner liner disposed inside the combustor shell and having an inner surface defining a cavity configured to receive hot combustion gases from a combustion chamber of the combustor, and an outer surface, the combustor shell and the inner liner defining an annular flow channel therebetween, and a segment carrier operatively connected to the inner liner and operative to receive an upper portion of the inner liner, the segment carrier and the inner liner defining a purging cavity therebetween. The inner liner includes a plurality of impingement jet holes configured to direct a flow of cooling air from the annular flow channel to the purging cavity.

Abstract: A structure is provided for a turbine engine. The structure includes a shell with a first surface, and a heat shield with a textured second surface and a textured third surface. The texture of a portion of the second surface is different than the texture of a portion of the third surface. The first surface and the second surface define a first cooling cavity between the shell and the heat shield. The first surface and the third surface define a second cooling cavity between the shell and the heat shield.

Abstract: A combustor for a gas turbine engine including a combustor liner support shell with a furrow formed therein; a forward liner panel mounted to the support shell via a multiple of studs, the forward liner panel including a forward liner panel rail the extends into the furrow; and an aft liner panel mounted to the support shell via a multiple of studs downstream of the forward liner panel, the aft liner panel including an aft liner panel rail that extends into the furrow.

Abstract: A combustor of a gas turbine engine includes a multiple of liner panels mounted to the support shell, at least one of the multiple of liner panels includes a first impingement cavity that operates at a first pressure and a second impingement cavity that operates at a second pressure different than the first pressure. A method of cooling a wall assembly within a combustor of a gas turbine engine includes directing air through a support shell and a liner panel that defines a first impingement cavity and a second impingement cavity. The first impingement cavity operates at a first pressure and the second impingement cavity operates at a second pressure that is different than the first pressure.

Abstract: Components for gas turbine engines having a hot side surface, a cold side surface, the cold side surface receiving cooling impingement at one or more cold locations, and at least one thermal transfer feature located between the hot side surface and the cold side surface within the component and arranged such that a condenser section of the thermal transfer feature is located proximate at least one of the cold locations and an evaporator section of the thermal transfer feature is located away from the cold location.

Abstract: A gas turbine includes a plurality of combustors and a turbine with a sealing arrangement between aft frames of the combustors and a first stage nozzle of the turbine. The sealing arrangement includes first and second circumferential outer seals with a radial side seal therebetween, the outer seals include circumferential extensions with an installation clearance, and the side seal includes a leaf that obstructs the installation clearance. The sealing arrangement also includes a notch on the side seal that forms a sliding joint with extensions of the first and second outer seals.

Abstract: A case is provided for a gas turbine engine. The case includes a wall defining a through-hole. The case also includes first and second pockets adjacent to, and on opposite sides, of the through hole. A method of reducing stress in a case of a gas turbine engine is also provided that includes reducing stress about a through-hole by providing a concavity on each side of the through-hole.

Abstract: A combustor wall is provided for a turbine engine. The combustor wall includes a shell and a heat shield that is attacked to the shell. The heat shield includes a rail and a cooling element connected to the rail in a cavity. The cavity extends in a vertical direction between the shell and the heat shield. The cavity fluidly couples a plurality of apertures in the shell with a plurality of apertures in the heat shield.

Abstract: An annular grommet is provided for a wall assembly of a combustor section of a gas turbine engine. The annular grommet includes a wall that at least partially defines a chamber. A wall assembly within a gas turbine engine includes a liner panel with a hot side and a cold side. The wall assembly also includes an annular grommet with a passage wall and a flange wall transverse to the passage wall. The annular grommet includes a chamber therein. A method of cooling a wall assembly within a gas turbine engine includes injecting air through a chamber in an annular grommet.

Abstract: A combustor liner for a gas turbine is provided. The combustor liner comprises a wall and a plurality of airflow injection holes in the wall arranged in a circumferentially-extending row, the plurality of airflow injection holes including a plurality of circular first airflow injection holes and at least one non-circular second airflow injection hole.

Abstract: A wall assembly for use in a combustor of a gas turbine engine includes a transverse structure with at least one effusion passage that extends at an angle ? therethrough. The effusion passage includes an inlet in an outer periphery of a wall. A wall assembly within a gas turbine engine include a liner panel generally parallel to a support shell and a transverse structure with at least one effusion passage that extends at an angle ? therethrough. The effusion passage includes an inlet in an outer periphery of a wall that is transverse to the liner panel and the support shell.

Abstract: A combustor including: a transition piece having a cylindrical shape and including an inlet of combustion gas at one end and an outlet of the combustion gas at the other end and configured to allow the combustion gas flowing from the inlet to flow out from the outlet so as to introduce the combustion gas into a turbine; a cooling medium introduction unit introduced with the cooling medium and provided on an outer periphery portion in an outlet side of the transition piece; a cooling medium inlet configured to introduce the cooling medium into the cooling medium introduction unit; and a cooling portion, connected to the cooling medium introduction unit, provided on a portion from the outlet of the transition piece toward the inlet and configured to allow the cooling medium from the cooling medium introduction unit to pass through the outlet toward the inlet.

Abstract: A combustion chamber includes at least one annular wall which includes at least one box like structure and each box like structure includes an inner wall, outer wall, upstream end wall and downstream end wall. The inner wall is spaced radially from the outer wall and the outer wall has a plurality of apertures for the supply of coolant into the box like structure. The inner wall, the outer wall, the upstream end wall and the downstream end wall are integral. The upstream end of the annular wall has features to secure the annular wall to an upstream ring structure and a downstream end of the annular wall has features to mount the annular wall on a downstream ring structure. The inner wall has at least one slot extending through the full thickness of the inner wall to accommodate differential thermal expansion between the inner wall and the outer wall.

Abstract: A combustor cooling panel has a first channel and a second channel through which a cooling medium flows. Relative to an end zone in a longitudinal direction of the first channel, the second channel has: an overlapping zone that overlaps the end zone in an extension direction of the end zone by extending in the extension direction while being separated from the end zone in a circumferential direction intersecting the extension direction; and a non-overlapping zone that does not overlap the end zone in the extension direction. A bent portion that is bent toward the end zone in the circumferential direction is formed in the second channel.

Abstract: A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a first plurality of openings and the first vane stage includes a second plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A first vane stage assembly and a method are also disclosed.

Abstract: The present disclosure relates to a gas turbine part, which can be exposed to high temperatures and centrifugal forces within a gas turbine. The gas turbine part can include plural sliced parts, wherein at least one of said sliced parts is made from a ternary ceramic called MAX phase, having the formula Mn+1AXn, where n=1, 2, or 3, M is an early transition metal such as Ti, V, Cr, Zr, Nb, Mo, Hf, Sc, Ta, and A is an A-group element such as Al, Si, P, S, Ga, Ge, As, Cd, In, Sn, Tl, Pb, and X is C and/or N.

Abstract: A method for mounting a combustion chamber of a gas turbine engine, wherein an annular outer combustion chamber wall and an annular inner combustion chamber wall are brought in position with respect to one another and are connected to a head plate, and wherein subsequently a combustion chamber head is mounted, characterized in that the head plate is connected to the outer combustion chamber wall and the inner combustion chamber wall by means of rivets that are arranged in a circumferentially distributed manner, and that subsequently the combustion chamber head is brought in position and is screwed together by means of threaded bolts and nuts to the arrangement of head plate, outer combustion chamber wall and inner combustion chamber wall, which is pre-mounted by means of the rivet.

Abstract: A combustor liner for a gas turbine is provided. The combustor liner comprises a wall and a plurality of airflow injection holes in the wall arranged in a circumferentially-extending row, the plurality of airflow injection holes including a plurality of circular first airflow injection holes and at least one non-circular second airflow injection hole.

Abstract: A turbine engine includes a diffuser connecting a compressor section and a turbine section. The diffuser includes a diffuser frame and multiple diffuser case sections are connected to the diffuser frame.