Abstract: A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel; a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to the first rail to form an interface passage; and at least one heat transfer feature within the interface passage.

Abstract: An electric power generating device for an aircraft, which is driven by an output of an aircraft engine, comprises an input shaft to which a driving force of the engine is transmitted; a transmission arranged with the input shaft; an electric power generator arranged with the input shaft and the transmission and driven by an output of the transmission; a driving force transmission mechanism disposed on a first end side of the input shaft in an axial direction thereof, the driving force transmission mechanism being configured to transmit the output from the transmission to the electric power generator; and a casing including a mounting section on the first end side of the input shaft in the axial direction, wherein the transmission includes an input section provided on a second end side of the input shaft in the axial direction, the input section being configured to receive as an input a rotational driving force from the input shaft, and an output section provided on the first end side of the input shaft in th

Abstract: A convergent-divergent nozzle for a turbofan engine of a supersonic aircraft, wherein the nozzle has an inner wall and forms a flow channel with a nozzle throat surface and a nozzle exit surface. A trim ring that is displaceable in the axial direction between a first position and a second position is provided, wherein the inner wall of the nozzle and the trim ring are embodied and positioned with respect to each other in such a manner that the trim ring extends at a radial distance to the inner wall in the first position, wherein a ring-shaped bypass channel is provided, extending between the trim ring and the inner wall of the nozzle, and the trim ring abuts the inner wall in the second position.

Abstract: A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel; and a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to said first rail, the second rail includes a discontinuous rail end surface.

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: Labyrinth seal system includes stationary component and rotatable component, where one of the stationary and rotatable components includes teeth. The labyrinth seal system further includes abradable component coupled to surface of other of the stationary and rotatable components and disposed facing the teeth. The abradable component includes a plurality of honeycomb cells disposed adjacent to each other along axial direction and circumferential direction. Each honeycomb cell includes a plurality of radial sidewalls, where each radial sidewall includes first portion and second portion. The first portion is coupled to the surface and the second portion extends from the first portion towards clearance defined between the stationary and rotatable components. The second portion is bent relative to radial axis of the labyrinth seal system.

Abstract: A combustor for a gas turbine engine includes a support shell with an angled shell interface; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel adjacent to the angled shell interface; and a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to said first rail, the second rail adjacent to the angled shell interface.

Abstract: A system for improving fuel-air mixing inside an injection system of a turbomachine combustion chamber. An air intake ring has an annular deflection wall, or venturi, having an internal profile provided with a discontinuity inducing an increase in the radius (?) of the internal profile downstream of the discontinuity. A method of atomizing fuel includes separating fuel trickling over the internal profile of the annular deflection wall from the internal profile at the level of the discontinuity so as to form droplets within a flow of air coming from an upstream air circulation space of the air intake ring.

Abstract: A liquid injection apparatus for use with a compressor having a rotor shaft and a casing part is provided. The liquid injection apparatus includes at least one conduit and at least one nozzle coupled to the conduit(s). The liquid injection apparatus is panel-shaped and has a first panel segment, a second panel segment, and at least one of: a passage defined between the first panel segment and the second panel segment, the passage sized to receive at least one of the rotor shaft and the casing part; and a third panel segment positioned between the first panel segment and the second panel segment. The third panel segment is at least in part detachable from the first panel segment and/or the second panel segment for receiving at least one of the rotor shaft and the casing part between the first panel segment and the second panel segment.

Abstract: The present disclosure is directed to a gas turbine engine, wherein the gas turbine engine defines a longitudinal direction, a radial direction, and a circumferential direction, and an axial centerline extended along the longitudinal direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes an annular stationary turbine frame centered around the axial centerline; an engine shaft extended generally along the longitudinal direction; an input shaft extended generally along the longitudinal direction; and a gear assembly including a first gear coupled to the input shaft, a second gear coupled to the turbine frame, and an inner spool coupling the first gear and the second gear, in which the inner spool defines a gear axis extended therethrough. The inner spool, the first gear, and the second gear are together rotatable about the gear axis. The gear axis is rotatable about the axial centerline of the engine.

Abstract: A premixing fuel nozzle assembly for a turbomachine is disclosed as having at least one fuel injector with an outer jacket having a first end and a second end and a threaded surface disposed therebetween. The first end has a cylindrical sealing surface and the second end has at least one wrenching pocket. An annular hub is disclosed as having an aperture with a first end and a second end and a threaded surface disposed therebetween, wherein the aperture is configured to at least partially house the fuel injector therein and the annular hub further has a conical seat about the first end. The fuel injector cylindrical sealing surface swages inward to form a hermetic seal when contacting the annular hub conical seat as the fuel injector is threaded into the aperture.

Abstract: An afterburner for use with a gas turbine engine includes a plurality of vanes distributed downstream of a turbine of the gas turbine engine. The vanes can include one or more exit apertures through which hot combustion flow from a pilot can be injected. The exit apertures can be protrusions or slots in some forms. In some embodiments, cooling passages are arranged around the exit apertures. An upstream vane portion can be positioned to inject fuel to be combusted via interaction with hot flow that is discharged through the exit apertures.

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: The present disclosure is directed to a nozzle assembly including a header manifold and a ring manifold, which defines a liquid fuel plenum, spaced from the header manifold. An outer sleeve connects to the ring manifold, and a nozzle body connects to the outer sleeve. The ring manifold, the outer sleeve, and the nozzle body define a fluid chamber. An inner fuel tube extends from the header manifold to the nozzle body. A portion of the inner fuel tube extends helically about an axial centerline of the nozzle assembly between the ring manifold and the nozzle body. A first fuel tube extends helically around a portion of the inner fuel tube. The first fuel tube fluidly couples the liquid fuel plenum to the header manifold. A second fuel tube extends helically around a portion of the inner fuel tube and fluidly couples the liquid fuel plenum to the header manifold.

Abstract: The present disclosure is directed to a dual-fuel fuel nozzle including a center body having a tube shape and a gas fuel plenum defined within the center body. The fuel nozzle also includes a plurality of turning vanes extending radially outward from the center body. Each turning vane includes at least one fuel port in fluid communication with the gas fuel plenum. A plurality of apertures is disposed through the plurality of turning vanes. The fuel nozzle further includes a ring manifold disposed within the center body downstream of the plurality of turning vanes. Additionally, the fuel nozzle includes a first fuel tube extending helically around a centerline of the center body. Furthermore, the fuel nozzle includes an air shield disposed within the center body and extending circumferentially around the first fuel tube.

Abstract: The present disclosure is directed to a combustion section for a gas turbine engine including a combustor assembly, an outer casing, a fuel injector assembly, and an inner casing. An inner liner and an outer liner are each extended at least partially along a lengthwise direction and at an acute angle. A dome assembly is extended between the inner liner and the outer liner, and the inner liner, and the outer liner together define a combustion chamber therebetween. A bulkhead assembly defining a plurality of walls is coupled to the inner liner and generally surrounding the outer liner and dome assembly. The outer casing surrounds the combustor assembly. The inner casing, the outer casing, and the combustor assembly together define a primary flowpath. An oxidizer flows through the primary flowpath in serial flow through the plenum and the fuel injector assembly into the combustion chamber.

Abstract: Aspects of the disclosure are directed to a system associated with an aircraft engine, comprising: a gearbox, a link coupled to the gearbox, and a bracket coupled to the link, where the bracket includes at least one notch formed in the bracket to provide a pilot for a fuse trigger, and where a portion of the bracket contained within the at least one notch breaks away from the bracket when the system experiences a load that is less than a nominal load associated with a fan blade out event of the aircraft engine.

Abstract: A fitting for forming a double-walled tube includes an inner fitting having a first inner weld lip, a second inner weld lip, and an inner body extending from the first inner weld lip to the second inner weld lip and having an outer surface with a first threading. The fitting also includes an outer fitting having a first outer weld lip, a second outer weld lip, and an outer body extending from the first outer weld lip to the second outer weld lip and having an inner surface with a second threading that is configured to mate with the first threading.

Abstract: A fan duct seal for a gas turbine engine has at least one seal body with a hollow center. A first opening is formed in the seal body that is in fluid communication with the hollow center. The first opening is configured to receive fan air flow. A second opening is formed in the seal body that is in fluid communication with the hollow center such that cooling flow enters the first opening, moves through the seal body and exits from the second opening. A fan duct for a gas turbine engine and a method of cooling a fan duct seal assembly are also disclosed.

Abstract: The present disclosure is directed to the operation and turndown of a segmented annular combustion system. The method includes injecting, via a fuel nozzle, a combustible mixture into a primary combustion zone between an adjacent pair of integrated combustor nozzles and burning the combustible mixture. The method further includes flowing air and injecting fuel into a premixing channel defined within a first integrated combustor nozzle to produce a second combustible mixture. The second combustible mixture is injected into a secondary combustion zone where it is combusted. The flow of combustion gases is accelerated, via turbine nozzles of the integrated combustor nozzles, toward turbine blades of a downstream turbine section. The method permits turndown of the combustion system by reducing or shutting off fuel to various components of the combustion system.