Abstract: A method for manufacturing a burner assembly for a gas turbine combustor, having a pilot burner extending along a longitudinal axis, and a premix burner surrounding the pilot burner, wherein the method includes manufacturing at least one portion of the pilot burner by an additive manufacturing technique which includes manufacturing at least one first thermal bridge connecting parts of the pilot burner having a temperature difference not greater than a threshold value.

Abstract: An ingestion restricting device is provided, which is particularly useful in a turbine engine. The ingestion restricting device can include a first component defining a first edge and including a plurality of ceramic matrix composite (CMC) plies forming outer, inner, and intermediate layers. The ingestion restricting device can also include a second component adjacent to the first component and defining a second edge. The first edge of the first component is opposite the second edge of the second component and defines a gap therebetween. Also, at least one ply of the plurality of CMC plies in the outer layer of the first component comprises an integral protrusion extending from the first edge a distance sufficient to traverse the gap and to overlap the second edge of the second component.

Abstract: A combustion chamber assembly for an engine with at least one tile component, on the cold side of which facing away from a combustion space, a positioning aid with at least two positioning elements, wherein a defined position relative to the combustion chamber component is predefined for the tile component via the contact of the at least two positioning elements on at least one contact face of the combustion chamber component in two spatial axes.

Abstract: A fuel swirler, for a gas turbine engine, has a primary cone housing defining an interior chamber. The interior chamber has an inlet in communication with a source of pressurized fuel. The interior chamber has a transition portion and a socket portion with an axisymmetric interior surface. A swirler core is disposed within the interior chamber. The swirler core has a downstream end and an upstream shank portion having an exterior surface mating the axisymmetric interior surface of the socket portion. The shank portion has a plurality of axially extending grooves. The grooves are disposed axisymmetrically about the exterior surface of the shank portion.

Abstract: A fuel nozzle for a gas turbine engine is generally provided. The fuel nozzle includes an outer sleeve extended circumferentially around a fuel nozzle centerline and extended along a longitudinal direction co-directional to the fuel nozzle centerline. The outer sleeve defines a plurality of first radially oriented air inlet ports through the outer sleeve in circumferential arrangement relative to the fuel nozzle centerline. A centerbody is positioned radially inward of the outer sleeve. The centerbody includes an outer annular wall extended circumferentially around the fuel nozzle centerline and extended along the longitudinal direction co-directional to the fuel nozzle centerline. The outer sleeve further includes a first radial wall extended from a downstream end of the outer annular wall. The centerbody defines a plenum radially inward thereof. The outer annular wall and the outer sleeve together define a first fuel air mixing passage extended along the longitudinal direction therebetween.

Abstract: A combustor with axially staged fuel injection includes an endcover and a fuel injector that extends axially downstream from the endcover. The fuel injector includes a cylindrical shell formed by an outer wall and an inner wall. A first plurality of outlets is circumferentially spaced across the inner wall. A first plurality of premix channels is defined between the outer wall and the inner wall. Each premix channel of the first plurality of premix channels is in fluid communication with a fuel supply, a compressed air supply and a respective outlet of the first plurality of outlets.

Abstract: A method of controlling at least part of a start-up or re-light process of a gas turbine engine, the method comprising: controlling ignition within a combustion chamber of the gas turbine engine; controlling rotation of a low pressure compressor using a first electrical machine, and controlling rotation of a high pressure compressor using a second electrical machine, the combustion chamber downstream of the low pressure compressor and high pressure compressor; determining if an exit pressure of the high pressure compressor is equal to or greater than a self-sustaining threshold pressure; and in response to determining that the exit pressure of the high pressure compressor is equal to or greater than the self-sustaining threshold pressure, ceasing controlling rotation of the low pressure compressor using the first electrical machine, and/or the high pressure compressor using a second electrical machine.

Abstract: A combustor includes a nozzle, a burner cylinder surrounding the outer circumference of the nozzle, and an outside flow path cylinder surrounding the outer circumference of a downstream end portion of the burner cylinder. A purge air flow path through which air flows is formed between the inner circumferential side of the outside flow path cylinder and the outer circumferential side of the burner cylinder. A tapered surface is formed in the downstream end portion of the burner cylinder such that the plate thickness of a burner cylinder-forming plate that forms the burner cylinder gradually thins toward the downstream side.

Abstract: A device for regulating the fuel supply flow rate for a turbine engine, comprising at least a pump, an upstream line leading to the pump, a downstream line downstream of the pump, a recirculation duct branched off the downstream line, allowing to collect the fluid in the downstream line and channel it towards the upstream line, a means for regulating the flow rate disposed on the recirculation duct, a phase measurer disposed in the upstream line, a computing unit connected to the phase measurer and to the flow rate regulating means, said computing unit controlling the flow rate regulating means such as to control the degree of opening of said flow rate regulating means, according to the gas content value in the upstream line.

Abstract: The present disclosure provides systems and methods for power production. In particular, the systems and methods utilize the addition of heat to an expanded turbine exhaust stream in order to increase the available quantity of heat for recuperation and use therein for heating a compressed carbon dioxide stream for recycle back to a combustor of the power production system and method.

Abstract: A boundary layer ingestion engine includes a fan section configured to extend into a boundary layer of a full annulus of an aft end of a fuselage of an aircraft. The fan section includes a first fan stage and a second fan stage. The boundary layer ingestion engine also includes a differential planetary gear system is operable to transform rotation of an input shaft into counter rotation of a first shaft coupled to the first fan stage and a second shaft coupled to the second fan stage. The boundary layer ingestion engine further includes a motor operable to drive rotation of the input shaft.

Abstract: A method for operating a gas turbine system includes utilizing a gas turbine controller to determine a schedule for a firing temperature for operative burners of a second combustor located downstream of a first combustor when the gas turbine system is operating in a low part load mode. During the low part load mode, multiple burners for the second combustor are switched-off. Further, the schedule is determined based on a position of inlet guide vanes of a compressor of the gas turbine system located upstream of both the first and second combustors. The method also includes controlling the firing temperature of the operative burners utilizing the schedule during the low part load mode to keep the gas turbine system within relevant operational limits of the gas turbine system.

Abstract: The invention concerns an assembly comprising: a fuel supply circuit (15, 15a, 15b) configured to supply fuel to a turbine heat engine, an electronic module (14, 14a, 14b), a power source (13, 13a, 13b) for supplying power to the electronic module (14, 14a, 14b), and a heat exchanger (16, 16a, 16b) positioned to allow a flow of heat from the electronic module (14, 14a, 14b) to the fuel supply circuit (15, 15a, 15b), the assembly being characterised in that the electronic module (14, 14a, 14b) comprises a phase-change material (PCM), configured to change state when the temperature of same reaches a predetermined phase-change temperature (Tf).

Abstract: A turbulence generating structure for liner cooling enhancement is a liner cooling structure applied to a double-structured side portion formed by a liner and a flow sleeve and includes the liner and a first turbulence generator protruding from a surface space of the liner and including a plurality of ribs arranged in an axial direction, each of the ribs comprising blocks arranged at regular distances in the axial direction and interspaced by cooling holes. The ribs include a first set of adjacent ribs separated by a first passage distance, the first passage distance being repeated in a circumferential direction, and a second set of adjacent ribs separated by a second passage distance, the second passage distance being repeated in the circumferential direction, wherein the first and second sets of adjacent ribs have exactly one rib in common, and the first passage distance is greater than the second passage distance.

Abstract: Systems and methods for adjusting airflow distortion in a gas turbine engine using a valved airflow passage assembly are provided. A gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. The compressor section can include a compressor. The gas turbine engine can further include a valved airflow passage assembly comprising a valve and a duct, the duct defining an inlet in airflow communication with the engine airflow path at a location downstream of the compressor and an outlet in airflow communication with the engine airflow path at a location upstream of the compressor, the duct comprising an airflow passage extending between the inlet and outlet. The valve can be operable with the airflow passage for controlling an airflow through the airflow passage to adjust airflow distortion.

Abstract: A gas turbine engine has a circumferential staging configuration of fuel injectors in a combustor. As a turbine blade revolves within the gas turbine engine it is subjected to lift and drag forces based on the configuration of lit injectors. A configuration of lit injectors that results in the minimum unsteady forces the turbine blade experiences is determined in order to increase the life span of the turbine blade and limit any structural failures.

Abstract: In accordance with one exemplary embodiment of the present disclosure, a method for providing overspeed protection for a gas turbine engine is provided. The gas turbine engine may include an engine core and an engine shaft. The method may include determining an overspeed condition of the engine. The overspeed condition may be indicative of an above normal rotational speed of the engine shaft. The method may also include reducing an airflow through the engine core of the gas turbine engine in response to the determined overspeed condition to reduce the rotational speed of the engine shaft.

Abstract: An accessory system for a gas turbine engine having a driveshaft with an axis of rotation is provided. The accessory system includes a towershaft coupled to the driveshaft and driven by the driveshaft along a towershaft axis of rotation. The accessory system also includes a shaft including a first shaft bevel gear coupled to a towershaft bevel gear. The shaft is rotatable by the towershaft along a shaft axis of rotation and defines an air-oil separator. The shaft axis of rotation is transverse to the towershaft axis of rotation. The accessory system includes a first accessory drive shaft having a first accessory bevel gear driven by the shaft. The accessory system also includes a second accessory drive shaft having a second accessory bevel gear driven by the shaft.

Abstract: An acoustic panel includes a base, a cantilevered portion, a gap, and a support member. The base has a surface defining a plurality of cavities configured to attenuate noise from an engine. The cantilevered portion extends from the base and is configured to be removably coupled with a portion of a transcowl. The gap is defined by the base and the cantilevered portion. The support member is coupled to the cantilevered portion and the base, and the supporting member is configured to support the cantilevered portion.

Abstract: A propulsion system for an aircraft. The propulsion system comprises a core, an internal fixed structure secured to the core and arranged around the core, and a nacelle surrounding the core and the internal fixed structure in which a secondary flow path is delimited between the internal fixed structure and the nacelle. The internal fixed structure has a slot which permits fluidic communication between a compartment on the inside of the internal fixed structure and the secondary flow path.