Abstract: A gas turbine engine according to an aspect of the present disclosure includes, among other things, propulsor means, first compression means, second compression means, reduction means for reducing a rotational speed of an output that drives the propulsor means relative to an input, first expansion means for driving the second compression means, and second expansion means for driving the input of the reduction means. The first compression means includes a plurality of arrays of rotor blades and a plurality of arrays of stator vanes arranged in a plurality of compression stages. A factor (F) is defined as a total number of the rotor blades in an array of the plurality of arrays divided by a total number of the stator vanes in an adjacent array of the plurality of arrays. The factor (F) is between 1.19 and 1.55 for at least two of the arrays of the rotor blades.

Abstract: A gas turbine engine includes a primary flow path fluidly connecting a compressor section, a combustor section and a turbine section. A cooling air flowpath is disposed radially outward of the primary flowpath. A first seal spans from an inner diameter of the cooling air flowpath to an outer diameter of the cooling air flowpath. The first seal includes at least one axial convolution and a plurality of pass through features defining a purge airflow. A heat shield is positioned immediately downstream of the first seal and is configured in relation to the first seal such that the purge airflow enters a mixing plenum defined between the heat shield and the first seal.

Abstract: An engine system is provided that includes a compressor section, a combustor section, a turbine section, a flowpath and a flow regulator. The combustor section includes a combustion chamber. The flowpath extends sequentially through the compressor section, the combustor section and the turbine section. The flow regulator is configured to open the flowpath between the compressor section and the combustion chamber during a first mode of operation. The flow regulator is configured to at least substantially close the flowpath between the compressor section and the combustion chamber during a second mode of operation.

Abstract: A method of forming a vane preform includes braiding a sleeve from ceramic fibers, inserting a flared mandrel into the sleeve, the flared mandrel comprising a body region and at least one contoured region, manipulating the sleeve using the flared mandrel to form contours in the sleeve corresponding to the contoured region of the flared mandrel, and laying up at least one ceramic ply with the sleeve to form the vane preform.

Abstract: An aircraft propulsion system includes a propulsive fan assembly configured for assembly into an aircraft structure, the propulsive fan assembly that includes a fan rotatable about a fan axis, an inlet duct assembly disposed within the aircraft fuselage, the inlet duct assembly that includes an upper inlet duct with an upper inlet opening and a lower inlet duct with a lower inlet opening. The upper inlet duct and the lower inlet duct merge into a common inlet duct forward of the propulsive fan assembly, and an outlet duct is disposed aft of the propulsive fan assembly.

Abstract: A gas turbine engine article includes a silicon-containing ceramic wall that has an external combustion gaspath side and an internal side that borders a cooling air cavity. The external combustion gaspath side has an associated combustion gas flow direction there along. An array of cooling holes extends through the silicon-containing ceramic wall and connects the internal side with the external combustion gaspath side. The cooling holes are oriented to discharge cooling air to the external gaspath side in a direction counter to the combustion gas flow direction.

Abstract: An airfoil includes an airfoil section that has an airfoil wall that defines an arced leading end, a trailing end, and first and second sides that join the arced leading end and the trailing end. The first and second sides span in a longitudinal direction between first and second ends. The airfoil wall circumscribes an internal core cavity. There is an arced rib in the internal core cavity. A cooling passage network is embedded in the airfoil wall between inner and outer portions of the airfoil wall. The cooling passage network has a trailing edge and an arced leading edge.

Abstract: A fuel delivery system for a combustor of a gas turbine engine including a primary fuel tank configured to store a primary fuel, a secondary fuel tank configured to store a secondary fuel, a swirler configured to produce a main flame within a combustion chamber of the combustor, and a fuel nozzle configured to produce a pilot flame within the combustion chamber of the combustor. The fuel nozzle includes a nozzle outlet that is located proximate to an end of the swirler or at the end of the swirler, the end of the swirler being located at an inlet of the combustor. The fuel delivery system also includes a primary fuel line fluidly connecting the primary fuel tank to the fuel nozzle and a secondary fuel line fluidly connecting the secondary fuel tank to the swirler.

Abstract: A method is provided for a gas turbine engine. During this method, a first lubricant parameter is determined during a first period of an engine operating cycle. The first lubricant parameter is indicative of a first quantity of lubricant within a reservoir of the gas turbine engine. A second lubricant parameter is determined during a second period of the engine operating cycle. The second lubricant parameter is indicative of a second quantity of the lubricant within the reservoir. The first lubricant parameter and the second lubricant parameter are compared to determine a lubricant consumption parameter. The lubricant consumption parameter is indicative of a quantity of the lubricant consumed by the gas turbine engine during the engine operating cycle.

Abstract: A brush seal system includes a component including a first mating surface. The brush seal system further includes a brush seal including a brush seal backing plate, a retaining ring, and a plurality of bristles retained between the brush seal backing plate and the retaining ring. The brush seal backing plate includes a second mating surface mounted to the first mating surface. The brush seal system further includes at least one cooling channel extending from an exterior side of the component to an interior side of the component so as to bypass the brush seal.

Abstract: A thermal management system for a gas turbine engine includes an additively manufactured nacelle component, at least a portion of the additively manufactured nacelle component forming an additively manufactured heat exchanger that extends into a fan bypass flow.

Abstract: A combustor for use in a gas turbine engine including a heat shield panel having a first surface and a second surface opposite the first surface of the heat shield panel and a combustor shell having an inner surface and an outer surface opposite the inner surface. The inner surface of the combustor shell and the second surface of the heat shield panel being in a facing spaced relationship defining an impingement cavity therebetween. The combustor shell further includes an impingement aperture that has a nonaxisymmetric shape. The impingement aperture extending from the outer surface to the inner surface through the combustor shell.

Abstract: An assembly is provided for an aircraft. This aircraft assembly includes a carrier, a first gear system, a second gear system and an idler. The first gear system includes a first sun gear, a first ring gear and a plurality of first intermediate gears. Each of the first intermediate gears is between and is meshed with the first sun gear and the first ring gear. Each of the first intermediate gears is rotatably mounted to the carrier. The second gear system includes a second sun gear, a second ring gear and a plurality of second intermediate gears. Each of the second intermediate gears is between and is meshed with the second sun gear and the second ring gear. Each of the second intermediate gears is rotatably mounted to the carrier. The idler gear couples the first ring gear to the second ring gear.

Abstract: A method of preparing a casting article for use in manufacturing a gas turbine engine part according to an exemplary aspect of the present disclosure includes, among other things, communicating a powdered material to an additive manufacturing system and preparing a casting article that includes at least one trunk and a skin core that extends from the at least one trunk out of the powdered material. A casting article is also disclosed.

Abstract: Disclosed is a multi-phase abradable coating including a ceramic matrix and a dislocator phase.

Abstract: A method for abrasive flow machining includes moving an abrasive media through a high-aspect passage of a workpiece. Local pressure of the abrasive media is increased at target abrasion surfaces of the high-aspect passage using a passage geometry that is configured to direct flow of the abrasive media into the target abrasion surfaces such that the target abrasion surfaces are preferentially polished by the abrasive media over other, non-targeted surfaces of the high-aspect passage at which the flow of the abrasive media is not directed into.

Abstract: A turbine engine has: a compressor; a combustor; a turbine; a gaspath passing downstream from the compressor through the combustor and then through the turbine; a fuel source; a fuel flowpath from the fuel source to the combustor; and a heat exchanger for transferring heat from the gaspath to the fuel flowpath The heat exchanger has: an inner wall in heat transfer relation with the gaspath; an outer wall; tubes between the inner wall and the outer wall bounding respective segments of the fuel flowpath; and a heat transfer fluid between the inner wall and the outer wall and in heat transfer relation with the tubes and the inner wall.

Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a carrier, a first gear system, a second gear system and a lock device. The first gear system includes a sun gear, a first ring gear and a plurality of first intermediate gears between and meshed with the sun gear and the first ring gear. Each of the first intermediate gears is rotatably mounted to the carrier. The second gear system includes a second ring gear and a plurality of second intermediate gears meshed with the second ring gear. Each of the second intermediate gears is rotatably mounted to the carrier. The lock device is configured to lock rotation of the carrier about an axis during a first mode. The lock device is configured to unlock rotation of the carrier about the axis during a second mode.

Abstract: A vane multiplet includes first and second ceramic matrix composite (CMC) singlet vanes that are arranged circumferentially adjacent each other. Each of the CMC singlet vanes includes an airfoil section and a platform at one end of the airfoil section. The platform defines forward and trailing platform edges and first and second circumferential side edges. A CMC overwrap conjoins the CMC singlet vanes. The CMC overwrap includes fiber plies that are fused to the platforms of the CMC singlet vanes.

Abstract: A coating for a blade root/disk interface includes a layer of soft metal matrix, and a solid lubricant distributed through the soft metal matrix. Examples of materials include CuAl as the soft metal matrix and MoS2 as the solid lubricant, although others are also disclosed.