Abstract: Nozzles and nozzle assemblies for gas turbine engines are provided. A nozzle includes an airfoil having an exterior surface defining a pressure side and a suction side extending between a leading edge and a trailing edge, an outer band disposed radially outward of the airfoil, the outer band including a radially outwardly-facing end surface, and an inner band disposed radially inward of the airfoil, the inner band including a radially inwardly-facing end surface. The nozzle further includes a flange extending radially from one of the radially outwardly-facing end surface or the radially inwardly-facing end surface. The flange is formed from a ceramic matrix composite material and includes a plurality of ceramic matrix composite plies stacked together and generally having an L-shape in a circumferential cross-sectional view.

Abstract: An airfoil assembly for a gas turbine engine and a method of transferring load from the ceramic matrix composite (CMC) airfoil assembly to a metallic vane assembly support member are provided. The airfoil assembly includes a forward end and an aft end with respect to an axial direction of the gas turbine engine. The airfoil assembly includes a radially outer end component, a radially inner end component, and a hollow airfoil body extending therebetween. The radially outer end component including a radially outwardly-facing end surface having a non-compression load-bearing feature extending radially outwardly and formed integrally with the outer end component, the load-bearing feature configured to mate with a complementary feature formed in a radially inner surface of a first airfoil assembly support structure and selectively positioned orthogonally to a force imparted into the airfoil assembly.

Abstract: A fiber preform for a hollow turbine engine vane, the preform including a main fiber structure obtained by three-dimensional weaving and including at least one main part, wherein the main part extends from a first link strip, includes a first main longitudinal portion forming a pressure side wall of an airfoil, an U-turn bend portion forming a leading edge or a trailing edge of the airfoil, a second main longitudinal portion facing the first main longitudinal portion and forming a suction side wall of the airfoil, and terminating at a second link strip. The first and second link strips are secured to each other and form a link portion of the main fiber structure. The main longitudinal portions are spaced apart so as to form a gap between the main longitudinal portions forming a hollow in the airfoil.

Abstract: A fan assembly for an aviation turbine engine, the assembly including a fan disk having at least one tooth and at least one platform mounted on the tooth of the fan disk, the platform including a box of composite material made from fiber reinforcement densified by a matrix, the box having a flow passage wall, a bottom wall, and two side walls extending radially between the bottom wall and the flow passage wall. The box of the platform includes an upstream opening at an upstream end of the platform and a downstream opening, and the assembly also includes a locking key housed in the box and passing through the upstream and downstream openings of the box, the locking key being blocked at each of its ends by a blocking element.

Abstract: A sealing ring segment for a stator of a turbine, the sealing ring segment has substantially the shape of a cylinder casing segment and has on its outer side a groove for fixing a plurality of guide vanes. The sealing ring segment has, for each guide vane that is fixable to the sealing ring segment, at least one pressure bolt that acts on the respective guide vane by a restoring force, the pressure bolt being able to be fixed and accordingly oriented by a corresponding opening in the sealing ring segment and being configured as a cylindrical element which can be compressed in the axial direction, and wherein the respective pressure bolt has a disk spring and the restoring force of the respective pressure bolt acts in the radial direction with respect to the rotation axis of the turbine in which the sealing ring segment is placed.

Abstract: A geared architecture for a gas turbine comprising an output shaft for connection with a fan, an input shaft and a gearbox connecting the input shaft with the output shaft. The gearbox has a forward planet carrier plate supported by a forward radially fixed bearing structure and a rearward planet carrier plate supported by a rearward radially fixed bearing structure. By supporting the carrier plates on radially fixed bearing structures the planet carriers can rotate about their own axis to lower vibration and mesh forces within the gearbox.

Abstract: A turbine engine assembly and methods involving a turbine engine assembly are provided. In one method, the turbine engine assembly is received. The turbine engine assembly includes an annular stator vane structure disposed at a first orientation. The stator vane structure is reconfigured with the turbine engine assembly to be disposed at a second orientation.

Abstract: The invention is related to a rotor blade for the generation of electrical power. The rotor blade transforms the kinetic energy of a fluid, into rotational movement of a mechanical shaft. The shape of the rotor blade is characterized in that, along an axis, it is longitudinally bound by a root (a) and a tip (b), which are connected through multiples curved segments, called neutral sectional axes [Eni]. All [Eni] generate a continuous or discontinuous curvature called Primary Neutral Axis [En]. The point corresponding to a leading edge and a trailing edge, configure an airfoil [PAij]. The curvature of the blade (e) has an arch of length L, and is defined by the neutral sectional axes [Eni]. The blade (e) is defined by at least one continuous curved section called primary neutral axis [En] having a length [Ln]. The blade's shape has a variable cross section along the Primary Neutral Axis [En].

Abstract: The invention relates to the field of centrifugal pumps, and in particular it relates to a centrifugal pump (303, 403) comprising at least: an impeller (303a, 403a); a rotary shaft (302, 402) secured to said impeller (303a, 403a); a casing (320, 420) having an axial admission passage (325, 425); at least one first bearing (305, 405) supporting said rotary shaft (302, 402) in said casing (320, 420); and at least one dynamic seal (311, 411) around the rotary shaft (302, 402), the impeller (303a, 403a) being situated between the at least one dynamic seal (311, 411) and the axial admission passage (325, 425) of the pump (303, 403). The centrifugal pump (303, 403) also comprises, between the impeller (303a, 403a) and at least one dynamic seal (311, 411), an axial force compensation disk (330, 430) secured to the shaft (302, 402) and presenting a diameter greater than 70% of a diameter of the impeller (303a, 403a).

Abstract: A geared architecture for a gas turbine comprising an output shaft for connection with a fan, an input shaft and a gearbox connecting the input shaft with the output shaft. The gearbox has a forward planet carrier plate supported by a forward radially fixed bearing structure and a rearward planet carrier plate supported by a rearward radially fixed bearing structure. By supporting the carrier plates on radially fixed bearing structures the planet carriers can rotate about their own axis to lower vibration and mesh forces within the gearbox.

Abstract: A heat shield (44) for an external component, such as a gearbox (42), of a gas turbine engine (20) for limiting heat transfer from the engine (20) to the external component is disclosed. The heat shield (44) may be formed to the shape of the gearbox (42) to reduce the geometric envelope of the gearbox (42) and heat shield (44). The heat shield (44) may be formed by a thin thermal insulation core (50) surrounded by a metal sheet (51). The small thickness of the heat shield (44) also reduces the geometric envelope of the heat shield (44). Since no extra equipment is needed to provide cooling fluid to the gearbox (42), the heat shield (44) reduces the overall weight of the gas turbine engine (20) as well.

Abstract: Hub assembly and propeller blade assemblies incorporating partial hub assembly having a backing plate with a set of partial root seats and a set of clamps selectively operably couplable to the backing plate with each clamp defining a second partial root seat and wherein when a clamp is coupled to the backing plate a root seat is formed, propeller blades having roots are received within the root seats.

Abstract: A gas turbine includes at least one rotor assembly and at least one compressor casing. The compressor casing has at least one inner compressor casing chamber for arranging the rotor assembly and at least one outer compressor casing chamber for tempering the compressor casing. The inner and outer compressor casing chambers are separated from each other by a separating casing wall. The outer compressor casing chamber has at least one boundary casing wall. The boundary casing wall and the separating casing wall are oppositely spaced such that the outer compressor casing chamber is formed. The boundary casing wall has at least one inlet opening for leading in an inlet tempering gas flow with tempering gas into the outer compressor casing chamber such that a tangential material temperature variation of the compressor casing is reduced in comparison to a non tempered compressor casing.

Abstract: The invention relates to an impeller for an axial fan, having an impeller body, which has an outer impeller shell and a carrier disc having a hub for connection for conjoint rotation to a drive shaft. A multiplicity of radially outward-pointing blades, which each have a blade root and a vane, is arranged on the carrier disc. Each blade root has a fastening portion accommodated with positive engagement between two fastening segments each arranged opposite one another on both sides of the carrier disc. The two opposite fastening segments and the carrier disc each have at least two through-holes for accommodating a long sleeve and a short sleeve. The fastening portion can be clamped firmly between the two opposite fastening segments with the aid of screw bolts, in each case inserted through the sleeves, and nuts, wherein each blade root is guided in a respective recess of the carrier disc.

Abstract: A bladed rotor arrangement comprises a rotor, a plurality of rotor blades and a plurality of lock plates. The rotor blades are mounted in circumferentially spaced axially extending slots in the periphery of the rotor. A plurality of lock plates are arranged at a first axial end of the rotor and a plurality of lock plates are arranged at a second axial end of the rotor. The radially outer ends of the lock plates engage corresponding grooves defined by radially inwardly extending flanges on the platforms of the rotor blades. The radially inner ends of the lock plates also engage circumferentially extending grooves. The radially inner end of each lock plate has three crushable, deformable, circumferentially spaced feet to reduce the risk of the lock plates chocking against the rotor, or seal plates and additional radial loads being reacted by the rotor blades into the rotor.

Abstract: A gas turbine engine airfoil assembly includes an airfoil and an attachment structure respectively bonded to opposing sides of a platform. At least one of the airfoil, the platform and the attachment structure are constructed from a ceramic matrix composite.

Abstract: An engine includes circumferentially spaced turbine blades radially inward a casing and circumscribed by a carrier section having segments, each having a carrier wall radially inward the casing and radially outward the turbine blades. The wall has one or more portions facing the casing. At least one of the portions has one or more impingement apertures for air passage of a predetermined temperature from a feed source into impingement onto the casing. The segments are radially inward the casing and radially outward the turbine blades, with the portions of their respective walls facing the casing. A method of controlling the gas turbine engine turbine casing temperature includes: passing air of a predetermined temperature from a feed source through the apertures in the one or more portions and into impingement on the casing; and optionally exhausting the air impinged onto the casing from a space between the segment and the casing.

Abstract: A rotor assembly for a gas turbine engine includes a rotor disk. The rotor disk includes a heat barrier feature radially inward of a rotor blade and radially outward of a rotor disk.

Abstract: An outer diffuser case of a diffuser case assembly for use in a gas turbine engine may have a housing orientated about an engine axis, a forward flange projecting radially outward from the housing, a mid-flange spaced axially aft of the forward flange and projecting radially inward from the housing, and an aft flange projecting radially outward from the housing. The mid flange is spaced axially between the forward and aft flanges and is configured to detachably engage an inner diffuser case of the case assembly. The forward flange is configured to detachably engage a high pressure compressor and the aft flange is configured to detachably engage a high pressure turbine of the engine.

Abstract: A gas turbine engine includes a compressor section in series flow with a turbine section. The compressor section includes one or more compressors and the turbine section includes one or more turbines, the compressor section and turbine section together defining a core air flowpath. A sump is positioned inward of the core air flowpath for containing and collecting lubrication oil. Additionally, a heat exchanger is positioned in direct thermal communication with the sump for removing heat from lubrication oil contained therein to reduce a size and/or amount of heat exchangers located radially outward of the core air flowpath.