Abstract: A combustor system for a gas turbine engine includes a combustor case, a combustor liner disposed within the combustor case and defining a main combustion chamber, a dome defining an upstream end of the main combustion chamber, and at least one torch injector attached to the dome of the main combustion chamber configured to inject combustion gases into the main combustion chamber. Each torch injector includes a torch injector housing defining and surrounding a torch combustion chamber configured to house a combustion reaction, an outlet passage defined by the torch injector housing, an electrothermal ignition source extending at least partially into the torch combustion chamber, and a fuel injector configured to inject fuel into the torch combustion chamber to at least partially impinge on the electrothermal ignition source and generate the combustion gases. The outlet passage directly fluidly connects the torch combustion chamber to the main combustion chamber.

Abstract: A gas turbine engine that includes a stage of guide vanes, a compressor downstream from the stage of guide vanes, and a combustor downstream from the compressor. The combustor includes a primary combustion zone and a secondary combustion zone downstream from the primary combustion zone. The primary combustion zone includes an exit configured to channel combustion gases towards the secondary combustion zone. A controller is communicatively coupled with the stage of guide vanes, the controller configured to monitor a temperature at the exit of the primary combustion zone, and selectively open and close the guide vanes to maintain the temperature within a predefined temperature range.

Abstract: A flow device for a combustor assembly defining a longitudinal axis around which the flow device is positioned is provided. The flow device includes a body extended around the longitudinal axis. The body includes a first wall and a second wall separated longitudinally from one another along the longitudinal axis. A fuel injector passage is defined through the body coaxial to the longitudinal axis. A plurality of vane walls is extended radially between the first wall and the second wall relative to the longitudinal axis. A radial flow passage is extended between the plurality of vane walls and the fuel injector passage.

Abstract: A fuel spray nozzle including a primary atomizer to discharge a flow of swirled atomised fuel along and around a fuel spray nozzle axis. The primary atomiser includes outer air swirler disposed radially outwardly of a fuel pre-filmer channel. A secondary atomiser disposed around the primary atomiser includes secondary inner air swirler to swirl flow along an inner air channel. The secondary inner air swirler disposed radially inwardly of a secondary fuel pre-filmer channel of the secondary atomiser. A primary outer air channel defined between the primary outer swirler and the secondary inner swirler. The secondary inner air swirler include splitter wall to separate swirling flow in the secondary inner channel from the primary flow of atomised fuel. The secondary inner air swirler includes primary cap wall integral with and extending radially inwardly from the splitter wall to direct flow from the primary outer channel inwardly towards the fuel spray.

Abstract: A gas turbine includes a plurality of first-stage stator vanes arranged in a circumferential direction, the first-stage stator vanes each including a vane surface having a pressure surface and a suction surface, a first combustor disposed on a suction surface side of one of the first-stage stator vanes, the first combustor having a first combustor outlet which includes a first side wall portion extending along a radial direction, and a second combustor disposed on a pressure surface side of the one of the first-stage stator vanes and adjacent to the first combustor in the circumferential direction. The one of the first-stage stator vanes satisfies 0.05??y/P?0.25, where ?y is a protruding amount of the suction surface from the inner wall surface of the first side wall portion to the circumferential direction, and P is an arrangement pitch of the first-stage stator vanes in the circumferential direction.

Abstract: The present invention is a downstream plate which is provided on a downstream side of an upstream plate includes an air inlet into which cooling air is introduced from a cooling air introduction pipe and a cooling flow path which extends in a direction along a surface of the downstream plate from the air inlet so as to avoid a premix tube and through which cooling air flows.

Abstract: A rotating detonation combustion (RDC) system is provided. The RDC includes a first outer wall and a second outer wall each extended around a centerline axis, and a detonation chamber formed radially inward of the second outer wall. A fuel passage extended between the first outer wall and the second outer wall, the fuel passage including a first inlet opening proximate to the aft end through which a flow of fuel is received into the fuel passage. The flow of fuel is provided through the fuel passage from the aft end to the forward end of the RDC system and to the detonation chamber.

Abstract: A core engine article includes a combustor liner defining a combustion chamber therein and a turbine nozzle. The combustor liner includes a plurality of injector ports, and the plurality of injector ports have a shape that tapers to a corner on a forward side of the injector ports. The turbine nozzle includes a plurality of airfoils. The combustor liner and turbine nozzle are integral with one another. A method of making a core engine article is also disclosed.

Abstract: Disclosed is a linkage assembly for a gas turbine engine having a link having a first end, a second end, and a rod extending therebetween, the first end having a first sliding bearing disposed within a first sliding bearing housing, a fastener comprising a first flange and a second flange, a pin extending between the first flange and the second flange, wherein the first sliding bearing is pivotally connected to the pin; and a biasing member secured between the first flange and the sliding bearing housing, the biasing member contacting the sliding bearing housing and biasing the link against rotation about a center axis for the rod of the link.

Abstract: A combustor and a gas turbine including the same which can reduce a loss of pressure and enhance a cooling efficiency of a liner and transition piece are provided.

Abstract: An apparatus and method for mounting a combustion liner within a flow sleeve of a gas turbine combustion system is disclosed. A mounting system comprises a plurality of low-profile mounting tabs secured to a combustion liner where each of the mounting tabs are placed within slots of flow sleeve pegs when the combustion liner is installed in a flow sleeve. A plurality of liner stop brackets are removably secured to a flange of the flow sleeve and have an arm extending to be adjacent to a top contact surface of the mounting tabs. The mounting system reduces blockage to the surrounding airflow.

Abstract: A method is disclosed for controlling fuel injection in a reheat combustor of a gas turbine combustor assembly including a combustor casing defining a gas flow channel and a plurality of injection nozzles distributed in or around the gas flow channel; the method includes the step of distributing fuel among the injection nozzles according to a non-uniform distribution pattern.

Abstract: A combustion chamber for a turbomachine having a bottom wall, at least one mixing bowl for promoting the mixing of air and fuel, mounted in an opening in the bottom wall, at least one annular baffle mounted axially downstream of the bottom wall, with respect to the direction of the gas flow within the combustion chamber, around the opening. The baffle is produced in one piece with the mixing bowl so as to form a one-piece assembly that has at least one channel for the flow of cooling air. The channel has an air inlet located upstream of the bottom wall and an air outlet located downstream of the bottom wall. The air outlet is located radially opposite the baffle.

Abstract: A fuel supply system includes a first component configured to direct a fuel flow to a combustor of an engine system. The first component includes a first inner surface and a first outer surface. The fuel supply system also includes an outer coating disposed on the first outer surface of the first component. The outer coating is configured to thermally insulate a first interior of the first component to reduce non-catalytic coke formation in the first interior. Additionally, the fuel supply system includes an inner coating disposed on the first inner surface of the first component. The inner coating is configured to reduce the fuel flow from contacting a base material of the first inner surface of the first component to reduce catalytic coke formation in the first interior.

Abstract: A cable conduit for a bypass flow duct may comprise a head end, a sleeve, a foot end, a boot, and a split grommet wherein the head end is coupled to the sleeve opposite the foot end, wherein the foot end comprises a flared portion, wherein the boot comprises a first flange and a neck, wherein the boot is configured to couple at the neck to the flared portion of the foot end, wherein the head end comprises a second flange and a cutout penetrating into an interior volume of the sleeve, and wherein the split grommet is coupled within the interior volume of the sleeve.

Abstract: A gas turbine engine includes a low-pressure spool, a high-pressure spool, and an electric motor. The low-pressure spool includes a low-pressure compressor in rotational communication with a low-pressure turbine and a fan via a first shaft. The high-pressure spool includes a high-pressure compressor in rotational communication with a high-pressure turbine via a second shaft. The electric motor is operably connected to the second shaft. The electric motor is configured to apply a rotational force to the second shaft.

Abstract: A gas turbine engine component having a first surface in communication with a core airflow. The gas turbine engine component further includes a second surface, different than the first surface, for cooling the first surface. The gas turbine engine component further includes a heat transfer augmentor extending from the second surface. The gas turbine engine component further includes a heat transfer augmentor effusion hole extending through the gas turbine engine component from a sidewall of the heat transfer augmentor to the first surface.

Abstract: A method of operating a rotary machine below a minimum emissions compliance load in a response mode includes reducing a fuel split to zero. The fuel split apportions a total flow of fuel to the combustor between a first combustion zone and a second combustion zone. The method also includes determining a current operating temperature of the first combustion zone using a digital simulation of the rotary machine. The method further includes determining a target operating temperature of the first combustion zone. The target operating temperature enables the rotary machine to operate below a traditional Minimum Emissions Compliance Load (MECL) while still in compliance with emissions standards. The method also includes channeling a first flow of fuel to the first combustion zone. The first flow of fuel decreases the temperature of the first combustion zone to the target operating temperature.

Abstract: Methods and systems for increasing efficiency of combustion in a turbine, methods including expanding in an expansion unit a pressurized fluid stream to form an expanded, cooled fluid stream; exchanging heat between an oxygen containing stream and the expanded, cooled fluid stream to reduce temperature of the oxygen containing stream to a reduced temperature and create a reduced temperature turbine compressor inlet oxygen containing stream; and compressing the reduced temperature turbine compressor inlet oxygen containing stream to an operating pressure of the turbine, where the step of compressing the reduced temperature turbine compressor inlet oxygen containing stream is more efficient than compressing the oxygen containing stream.

Abstract: A fuel nozzle for a turbomachine includes a centerbody that extends axially with respect to a centerline of the fuel nozzle. A confining tube is positioned radially outward of the centerbody. A plurality of swirler vanes is disposed between the centerbody and the confining tube. Each of the plurality of swirler vanes includes a radially inner base and a radially outer tip. Each of the swirler vanes further includes an upstream portion that extends generally axially from a leading edge. A downstream portion extends from the upstream portion to a trailing edge. The downstream portion defines a bend length between the upstream portion and the trailing edge. The bend length at the radially outer tip is greater than the bend length at the radially inner base.