Abstract: A heat shield (100) and method for assembling such is disclosed. The heat shield (100) may comprise an outer wall (122) and an inner wall (124). The outer wall (122) includes a first member (130), a flange (132) extending outward from the first member (130) and a first inner edge (134). The first member (130) extends from the flange (132) to the first inner edge (134). The inner wall (124) includes a second member (140), a rim (142) extending outward from the second member (140) and a second inner edge (144). The second member (140) extends from the rim (142) to the second inner edge (144). The inner wall (124) is spaced apart from the outer wall (122), the first and second edges form an air gap (146) between them, and the inner wall (124) and the outer wall (122) form a cavity (148).

Abstract: An acoustic liner can comprise a support layer including a set of partitioned cavities defining a set of cells with open faces, as well as a first and second facing sheet operably coupled to the support layer that overlie and close the open faces. A set of aerogel fillings can be provided within at least some of the set of cells.

Abstract: A gas turbine engine including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft mounting structure that includes supporting bearings connected to the fan shaft; and an epicyclic gearbox that can receive input from the core shaft, and output drive to the fan shaft so as to drive the fan at a lower rotational speed than the core shaft. The gearbox includes a sun gear, planet gears, a ring gear, and a planet carrier for mounting the planet gears. A torsional stiffness of the planet carrier is greater than or equal to 1.6×108 Nm/rad. The fan shaft mounting structure has: a radial bending stiffness greater than or equal to 7.00×108 N/m, and/or a tilt stiffness greater than or equal to 1.50×107 Nm/rad.

Abstract: A centrifugal force (CF) bearing for a rotorcraft rotor assembly includes an inner-member, an outer-member, and an elastic member interposed between the inner- and outer-members. The outer-member includes an upper flange extension and a lower flange extension for location to and engagement with a grip. The upper flange extension is disposed over an upper surface portion of the elastic member, and the lower flange extension is disposed under a lower surface portion of the elastic member. The CF bearing is configured for attachment to a grip of the rotor assembly. The outer-member is configured to communicate mechanical loads (e.g., centrifugal force, lateral shear, vertical shear) from the grip to the yoke during operation of the rotor assembly.

Abstract: A gas turbine engine includes: an engine core with a turbine, compressor, and core shaft; a fan upstream the engine core; a gearbox that can receive an input from the core shaft and output drive to a fan shaft to drive the fan at a lower rotational speed than the core shaft. A fan-gearbox axial distance is greater than or equal to 0.35 m. The fan shaft has: (i) a radial bending stiffness ratio of a radial bending stiffness of the fan shaft at the fan input to a radial bending stiffness of the fan shaft at the gearbox output that is greater than or equal to 6.0×10?3; and/or (ii) a tilt stiffness ratio of a tilt stiffness of the fan shaft at the fan input to a tilt stiffness of the fan shaft at the gearbox output that is greater than or equal to 2.5×10?2.

Abstract: An engine for an aircraft includes an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox having a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted; and a gearbox support arranged to mount the gearbox within the engine. The fan shaft, core shaft, gearbox and the gearbox support together may form a transmission.

Abstract: A variable guide vane assembly is provided for a turbine defining a core air flowpath. The variable guide vane assembly includes an airfoil band defining a flowpath surface and a cavity. The variable guide vane assembly further includes an airfoil including a first end extending at least partially into the cavity of the airfoil band and an opposite second end, the airfoil extending generally along an axis between the first end and the second end and being moveable generally about the axis relative to the airfoil band. The variable guide vane assembly further includes a sealing element operable to form a seal between the first end of the airfoil and the airfoil band.

Abstract: An outlet tube arrangement for an aircraft nacelle may comprise a vent tube defining a vent tube outlet, the vent tube outlet configured to be flush with an outer surface of the aircraft nacelle, and an inner tube disposed within the vent tube, the inner tube configured to increase momentum of fluid flowing through the inner tube to urge the fluid to separate from the inner tube without wetting the aircraft nacelle upon exiting from the vent tube, wherein the inner tube comprises an inner tube inlet comprising an inlet cross-sectional area and an inner tube outlet comprising an outlet cross-sectional area, the inlet cross-sectional area being less than a cross-sectional area of the vent tube at the inner tube inlet, the inlet cross-sectional area being greater than the outlet cross-sectional area, and the inner tube outlet being flush with the vent tube outlet.

Abstract: A gas turbine engine for an aircraft has an engine core having a turbine, compressor, and core shaft connecting the turbine and compressor; a fan upstream the engine core, the fan having fan blades; and a gearbox. The gearbox receives an input from a core shaft and outputs drive to a fan to drive the fan at a lower rotational speed than the core shaft. The gearbox is an epicyclic gearbox and has a sun gear, planet gears, ring gear, and planet carrier on which the planet gears are mounted. The gearbox has a gear mesh stiffness between the planet gears and the ring gear and a gear mesh stiffness between the planet gears and the sun gear. The gear mesh stiffness between the planet gears and the ring gear divided by that between the planet gears and the sun gear is in the range from 0.90 to 1.28.

Abstract: A trailing edge for a rotor blade tip of an aerodynamic rotor of a wind turbine. The trailing edge comprises a trailing edge delimiting line, which replicates the contour of the trailing edge, and multiple serrations to improve flow behavior at the trailing edge. The serrations are provided at the trailing edge in dependence on the trailing edge delimiting line, and consequently on geometrical and operating parameters.

Abstract: An inlet bleed heat (IBH) system for use in a turbine engine including a silencer assembly. The inlet bleed heat (IBH) system includes a feed pipe for delivering compressor discharge air. The feed pipe includes a plurality of orifices along at least a portion of a length of the feed pipe, and each orifice of the plurality of orifices extends through a wall of the feed pipe for allowing the compressor discharge air to exit the feed pipe. The system also includes a heat shielding component that extends across the feed pipe, wherein the heat shielding component is configured to reduce heat transfer between the feed pipe and the silencer assembly of the turbine engine.

Abstract: The invention concerns a turbine for a gas turbine comprising a blade, a vane and an abradable lip attached to the blade or to the vane, wherein the blade and the vane are separated by a gap and the abradable lip extends part of the distance across the gap. Embodiments include the addition of an abrasive layer attached on the other side of the gap from the abradable lip. A method of manufacturing is also described.

Abstract: A variable pitch fan assembly includes variable pitch fan blades circumscribed about engine centerline axis coupled to a drive shaft centered about the engine centerline axis. Each blade pivotable about pitch axis perpendicular to centerline axis and having blade turning lever connected thereto. One or more linear actuators non-rotatably mounted parallel to engine centerline axis and operably linked to fan blades for pivoting fan blades and connected to spider ring through thrust bearings for transmission of axial displacement of non-rotatable actuator rods of actuators while the fan blades are rotating. Spider arms extending away from spider ring towards blade roots and each spider arm connected to one of the turning levers. Turning levers may be connected and caromed to spider arms by pin and slot joint. Each spider arm may include joint pin disposed through joint slot of turning lever. Joint slot may be angled or curved.

Abstract: A fan blade capable of cooling an enclosed space is disclosed, comprising: a leading edge; a trailing edge; a top side; a bottom side; and a plurality of passageways wherein each of the passageways narrows within the fan blade. The plurality of passageways may narrow at any point within the interior of the fan blade and may be shaped in any way to ensure that a narrowing occurs. The plurality of passageways may also be in thermal communication with the rest of the fan blade. The fan blades may be used as part of a fan having a hub and a plurality of fan blades. Also disclosed, is a method of cooling a room comprising rotating a fan, the fan comprising: a plurality of fan blades, each fan blade having a plurality of passageways, wherein each of the passageways narrows within the fan blade.

Abstract: A quill shaft which translates rotational energy between a gear and a drive shaft having a drive shaft spline includes a first leg having an inner spline which meshes with the drive shaft spline and a second leg concentric with the first leg and radially offset and separated from the first leg by a gap. The second leg includes an outer spline which meshes with the gear. The first leg is connected to the second leg at one end to allow for change to a gap size during while translating the rotational energy between the gear and a drive shaft.

Abstract: A gas turbine engine including a first rotor assembly is generally provided. The first rotor assembly includes outer drum and an outer drum airfoil. The outer drum airfoil is coupled to the outer drum and extended inward along a radial direction. A damper structure is coupled to one or more of the outer drum or the outer drum airfoil.

Abstract: A thermal management article is disclosed including a substrate and a first coating disposed on the substrate. The first coating includes a first coating surface and at least one passageway disposed between the substrate and the first coating surface. The at least one passageway defines at least one fluid pathway. A method for forming a thermal management article is disclosed including attaching at least one passageway to a substrate. The at least one passageway includes a passageway wall having a wall thickness and defines at least one fluid pathway. A first coating is applied to the substrate and the passageway wall, forming a first coating surface. The at least one passageway is disposed between the substrate and the first coating surface.

Abstract: Various embodiments include gas turbine seals and methods of forming such seals for advanced gas path (AGP) heat transfer designed gas turbine engines. In some cases, a turbine includes: a first arcuate component adjacent to a second arcuate component, each arcuate component formed of a ceramic matrix composite and including a seal slot located in an end face, each seal slot having a plurality of axial surfaces and radially facing surfaces extending from opposite ends of the axial surfaces and an overhanging seal assembly disposed in the seal slot. The overhanging seal assembly including an intersegment seal including a plurality of seal segments defining two vertical seal segments and a plurality of horizontal seal segments and a first overhanging positioner shim seal and a second overhanging positioner shim seal, each including an extended portion. The overhanging seal assembly disposed in the seal slot so as to provide position constrainment of the two vertical seal segments.

Abstract: A composite rotatable assembly for an axial flow compressor comprises a spool having a plurality of blade assemblies arranged in stages on the spool and attached thereto by a wound band. Each blade assembly comprises a blade and a base, with the base having a forward tang extending axially forward of a leading edge of the blade and an aft tang extending axially aft of a trailing edge of the blade. The band is wound over at least a portion of the forward and aft tangs of the plurality of blade assemblies to hold the blade assemblies to the spool under centrifugal loading. An abradable layer may be added over the wound band.

Abstract: An airfoil includes an airfoil body that has a geometrically segmented coating section. The geometrically segmented coating section includes a wall that has an outer side. The outer side includes an array of cells and there is a coating disposed in the array of cells. The cells include respective secondary mechanical bonding features that secure the coating to the wall.