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: Described herein are embodiments of systems and methods for the removal of a direct drive unit (DDU) housed in an enclosure, such as a direct drive turbine (DDT) connected to a gearbox for driving a driveshaft connected to a pump for use in a hydraulic fracturing operations.

Abstract: A cooling system for an aircraft has at least one moveable member configured to cover an opening formed within an aircraft outer skin. An actuator moves the at least one moveable member between a fully open position where external atmosphere air can be directed through the opening to an internal passage enclosed by the aircraft outer skin and a fully closed position where the opening is covered. A controller selectively controls the actuator to move the at least one moveable member between the fully open and fully closed positions. An aircraft engine and a method of cooling an aircraft engine in an aircraft are also disclosed.

Abstract: Described herein are embodiments of systems and methods for the removal of a direct drive unit (DDU) housed in an enclosure, such as a direct drive turbine (DDT) connected to a gearbox for driving a driveshaft connected to a pump for use in a hydraulic fracturing operations.

Abstract: A method of operating a turbine unit, wherein recirculated exhaust gas is contacted with a cooling and absorption liquid in a packed bed. An exhaust gas treatment system for a turbine unit, wherein an exhaust gas recirculation line comprises a gas cooling and cleaning device having a packed bed for contacting the exhaust gas with a cooling and absorption liquid. A combined cycle power generating system, wherein an exhaust gas recirculation line comprises a gas cooling and cleaning device having a packed bed for contacting the exhaust gas from a gas turbine with a cooling and absorption liquid and wherein water utilized as a cooling medium for condensation of steam originating from a steam turbine, and the cooling and absorption liquid, are passed to a cooling tower.

Abstract: A gas turbine engine comprises a fan delivering air into a bypass duct as bypass air and into a core engine. The core engine includes a high pressure compressor, a combustor, and a turbine section including a fan drive turbine driving a fan rotor through a gear reduction. The high pressure compressor having a downstream most location and air in a chamber radially outward of the combustor being air downstream of the downstream most location in the high pressure compressor. A tap taps compressed air and moves compressed air through a heat exchanger, then returns the compressed air back into a core engine housing and passes the returned air through a conduit radially outwardly of the combustor. The air is passed from the conduit radially inwardly to cool the turbine section.

Abstract: A combustion section for gas turbine engine includes an inner liner, an outer liner, and a dome attached to the inner liner and to the outer liner. A heat shield is attached to the dome, with the dome, the heat shield, or both define an opening. The combustion section also includes a fuel nozzle extending at least partially into the opening, the fuel nozzle defining a fuel nozzle axis and a radial direction relative to the fuel nozzle axis. The fuel nozzle defines an aft end, the aft end of the fuel nozzle positioned forward of the aft surface of the heat shield along the fuel nozzle axis such that the aft end of the fuel nozzle defines a minimum separation from the aft surface of the heat shield along the fuel nozzle axis of at least about 0.15 inches.

Abstract: The present disclosure is directed to a coupling assembly for a turbomachine. The coupling assembly includes a liner and a sleeve at least partially positioned circumferentially around the liner. A frame is positioned between the liner and the sleeve. The frame includes a first side wall, a second side wall, and a base wall extending from the first side wall to the second side wall. A block is positioned between the liner and the sleeve and is movable relative to the frame to permit the block to move toward and away from the base wall. Moving the block away from the base wall causes the block to exert an inward force on the liner and the base wall to exert an outward force on the sleeve.

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: An aft section structure of a jet engine with an axis is comprised of a casing defining a duct around the axis and opened axially fore and aft; a cone tapering aftward at a first angle with the axis and having a pointed end; guide vanes, each of the vanes radially extending from the cone to the casing and comprising a pressure side at a second angle with a plane containing the axis; spray bars, each of the spray bars extending radially within the duct and comprising trailing sides, each of the trailing sides being directed aftward at a third angle with a plane containing the axis; and flame holders, each of the flame holders extending radially within the duct and comprising one or more interior sides, each of the interior sides being directed aftward at a fourth angle with a plane containing the axis.

Abstract: A compressor is provided including a casing, a hub, a flow path, a plurality of blades, and an end-wall treatment formed in at least one of the casing and the hub, and facing a tip of each blade. The flow path is formed between the casing and the hub, and the plurality of blades is positioned in the flow path. The tip of each blade and the end-wall treatment are configured to move relative to each other. Such end-wall treatment includes a first recess portion extending along a first axis to maintain a fluid flow substantially straight through the first recess portion. The end-wall treatment further includes a plurality of second recess portions spaced apart from each other and extending from the first recess portion along a second axis different than the first axis to maintain the fluid flow substantially straight through the plurality of second recess portions.

Abstract: The turbine housing assembly includes a housing surrounding a plurality of turbine shroud segments mounted to support members of the housing. Impingement holes extend through the housing and have outlet openings communicating with a cavity between the shroud segments and the support. A deflector rail protrudes axially away from the support members into the cavity. The deflector rail defines a flow-redirecting surface to redirect the cooling air flow from the impingement holes radially outwardly, away from the turbine shroud segments.

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 gas turbine engine according to the present disclosure includes a first compressor and a first turbine for driving the first compressor. A core section includes a second compressor and a second turbine for driving the second compressor. A third turbine is arranged fluidly downstream of the first turbine and the second turbine and configured to drive a power take-off. A first duct system is arranged fluidly between the low-pressure compressor and the core section. The first duct system is arranged to reverse fluid flow before entry into the core section.

Abstract: A gas turbine engine includes a gearbox including a sun gear, annulus gear, plurality of planet gears and a planet gear carrier. Each planet gear is rotatably mounted in the planet gear carrier by a planet gear bearing. A lubrication system is arranged to supply lubricant to the planet gear bearings and at least one of the sun gear, the planet gears and the annulus gears. The planet gear carrier includes an annular extension. The lubrication system includes an annular member arranged around annular extension to define an annular chamber which is sealed at its ends to the planet gear carrier and which contains a reserve of lubricant. The annular member has an inlet to supply lubricant into the annular chamber and the planet carrier has a first passage to supply lubricant to the sun, planet or annulus gear and a second passage to supply lubricant to planet gear bearings.

Abstract: A gas turbine transition duct includes: an inner tubular body, defining a transition channel and having a first upstream end and a first downstream end for coupling to a can combustor and to a turbine section of a gas turbine assembly, respectively; an outer tubular body, arranged around the inner tubular body and having a second upstream end at the first upstream end of the inner tubular body and a second downstream end at the first downstream end of the inner tubular body; wherein a convective cooling channel is defined between the inner tubular body and the outer tubular body, the convective cooling channel having an inlet between the first downstream end and the second downstream end; and wherein the outer tubular body is continuous between the second upstream end and the second downstream end.

Abstract: A combustion duct assembly for a gas turbine includes a liner, an annular spring seal and a transition piece. The annular spring seal is outwardly convex and coupled to an outer face of the downstream side of the liner. The transition piece has includes an inner wall and an outer wall that surrounds the inner wall. The convex part of annular spring seal is operable to elastically engage the inner wall of the transition piece. An end portion of the downstream side of the annular spring seal is fixed to the line. The annular spring seal includes a plurality of through holes formed in an upstream side of a point where the annular spring seal comes into contact with the inner wall. The liner includes a plurality of vent holes formed in an inner area surrounded by the spring seal.

Abstract: A gas turbine engine has a LP spool including an LP shaft and a HP spool including an HP shaft, an accessory gear box (AGB) drivingly engaged to a plurality of accessories, an engine starter drivingly engaged to the AGB, and a drive shaft drivingly engaged to the AGB. In a first engine mode, the drive shaft is drivingly engaged to the HP shaft and is disengaged from the LP shaft, the engine starter is in driving engagement with the HP shaft via the drive shaft, and the LP shaft is disengaged from the AGB. In a second engine mode, the drive shaft is drivingly engaged to the LP shaft and is disengaged from the HP shaft, the LP shaft is in driving engagement with the AGB via the drive shaft, and the HP shaft is disengaged from the AGB.

Abstract: An assembly comprises a first cooling cavity disposed within one or more of a turbine assembly or a combustion chamber of an engine. The first cooling cavity directs cooling air within the one or more of the turbine assembly or the combustion chamber. The assembly comprises a second cooling cavity also disposed within the one or more of the turbine assembly or the combustion chamber. The second cooling cavity receives at least some of the cooling air from the first cooling cavity. A forward facing step nozzle forms a channel that fluidly couples the first cooling cavity with the second cooling cavity. The step nozzle includes steps having elongated first sides and narrow second sides. The elongated first sides of the steps protrude into the channel such that a cross-sectional area of the channel of the step nozzle at the steps is smaller than a cross-sectional area of the channel of the step nozzle outside of the steps.

Abstract: A fuel injector is provided for a gas turbine engine. The fuel injector includes a fuel conduit and a cooling fluid circuit through a strut. A gas turbine engine is provided that includes a multiple of fuel injectors in communication with a combustor and a cooling system in communication with each of the multiple of fuel injectors.