Abstract: A wind-turbine rotor blade having a rotor blade portion which comprises an inner rotor blade portion with a first end having a plurality of fastening units for fastening to a hub of a wind turbine and a second end having a flange for fastening further portions of the wind-turbine rotor blade, wherein the inner rotor blade portion is produced from polymer concrete, wherein the rotor blade portion has a plurality of resin-impregnated fiber-composite laid scrims or rovings which are wound around the inner rotor blade portion.

Abstract: A bushing anti-rotation system and apparatus. A bushing anti-rotation system includes an impeller and diffuser stage including a rotatable shaft extending longitudinally through the stage, the diffuser having an inner diameter defining a bore, a bearing set including a bushing pressed into the inner diameter, and a rotatable sleeve inward of the bushing, and means for wedging the bushing against the inner diameter of the diffuser, wherein the means for wedging provides a mechanical lock against rotation of the bushing. The means for wedging the bushing against the diffuser includes one of a key wedged between an eccentric cavity in the diffuser and a notch in the bushing, a pin protruding from the diffuser against a flattened portion of the bushing, a square-shaped bushing secured into a square-shaped bore, a bent retaining ring having a bent end that wedges against a diffuser groove wall, or a combination thereof.

Abstract: An aerofoil body for a gas turbine engine is provided. The aerofoil body has leading and trailing edge portions, wherein one of the leading and trailing edge portions is a morphable edge portion having a composite layer structure. The aerofoil body further has a non-morphing central portion which forms pressure and suction surfaces of the aerofoil body between the leading and trailing edge portions. The composite layer structure includes a return spring, one or more shape memory alloy layers, and a flexible cover for the return spring and the one or more shape memory alloy layers. The flexible cover defines pressure and suction surfaces of the aerofoil body at the morphable edge portion. The one or more shape memory alloy layers are electrically heatable to deform the layers against the resistance of the return spring, and thereby alter the pitch of the aerofoil body at the morphable edge portion.

Abstract: Systems and methods for reducing heat exposure of a turbine casing in a gas turbine engine may be provided. The system may include blade track segments. The blade track segments may include a cavity formed at a circumferential end of the blade track segment. The blade track segments may further include an assembly groove configured to provide a clearance for axial insertion of an inter-segment strip seal. An inter-segment strip seal may be axially inserted past the assembly grooves and into the cavity. The inter-segment strip seal may cover a gap between adjacent blade track segments, thus blocking hot, high pressure fluid from contacting the turbine casing.

Abstract: A plurality of circumferential blowout passages that communicate with a cavity into which cooling air flows and that open in a circumferential end surface are formed in a shroud of a vane. Of a central region of the circumferential end surface, an upstream-side region of the circumferential end surface, and a downstream-side region of the circumferential end surface, at least the central region has openings of the plurality of circumferential blowout passages formed therein. A density that is the number of the openings of the circumferential blowout passages per unit length in an axial direction is highest in the central region.

Abstract: Airfoil bodies having a first core cavity within the airfoil body and a second core cavity located within the airfoil body and adjacent the first core cavity, wherein the second core cavity is defined by a first cavity wall, a second cavity wall, a first exterior wall, and a second exterior wall, wherein the first cavity wall is located between the second core cavity. The first cavity wall includes a first surface angled toward the first exterior wall and a second surface angled toward the second exterior wall and at least one first cavity impingement hole is formed within the first surface. A funneling feature extends the second core cavity along the first exterior wall to funnel the first impingement flow along the first exterior wall.

Abstract: A method of sealing one or more openings provided in a wall of an aerofoil for a gas turbine engine, the aerofoil comprising at least one cavity which is at least partly filled with a vibration damping material, the method comprising steps to provide a metallic material onto the wall of the aerofoil in order to cover the opening and bond the metallic material to the wall of the aerofoil to seal the opening.

Abstract: Embodiments of the present invention generally relate to a runner unit of a tidal power plant, and more particularly to a device for reversing a blade of the runner unit. The device according to the embodiments is lighter and more efficient with respect to known solutions which involve articulated mechanisms as it is based on a reversing servomotor including an annular piston which acts on the blade to be reversed.

Abstract: A locking spacer, which is fitted in a dovetail slot provided on an outer circumferential surface of a disk put on a rotor shaft, includes: a pair of first blocks each provided with a dovetail joint, and configured to have a size occupying a portion of an internal space of the dovetail slot; a pair of second blocks having a size occupying a portion of the internal space of the dovetail slot, the portion not being occupied by the pair of first blocks, and each being provided with a locking groove; and a locking block having a size occupying a portion of the internal space of the dovetail slot, the portion not being occupied by the first and second blocks, and being provided with a rotating locking arm configured such that opposite end portions thereof are inserted into the locking grooves.

Abstract: Bleed air system includes bleed port with annular bleed slot in a compressor inner casing surrounding flowpath circumscribing a centerline in compressor inner casing. Inner and outer slot walls bound bleed flow passage through the bleed slot. Annular array of arcuate shields radially inwardly bound the bleed flow passage downstream and aft of inner slot wall. Each shield includes integrally-formed axially forward retaining portion, downstream portion, and upstream portion therebetween. Clipping means clips retaining portion to aft facing radial wall of annular lip on inner slot wall. Clipping means may be arcuate or annular clip having forward annular portion attached to annular aft flange substantially perpendicular to the centerline and forward annular portion disposed in annular slot extending into inner slot wall. Forward annular portion and annular slot may be cylindrical or conical. Clip may include aft conical portion radially offset from forward conical portion and transition portion therebetween.

Abstract: Disclosed is a flutter damper, including a chamber having an internal space, a movable diaphragm disposed at least partially within the chamber, the diaphragm separating the chamber into a first chamber and a second chamber, the second chamber forming an acoustic volume, wherein a size of the acoustic volume configures the chamber for peak acoustical energy absorption at a frequency range that is associated with one or more fan flutter modes, and a biasing member that moves the diaphragm responsive to a signal from an aircraft or engine electronic control (EEC) unit, wherein movement of the diaphragm increases or reduces the size of the acoustic volume of the second chamber.

Abstract: A centrifugal compressor casing, comprising a bundle which rotatably supports a rotating shaft, and an impeller fixed to the rotating shaft and that rotates with the rotating shaft about an axis of the rotating shaft, and in which an annular suction flow channel centered on the axis introduces fluid into the flow channel of the impeller, and an annular discharge flow channel centered on the axis that discharges the fluid from the flow channel of the impeller, are formed; a casing main body which covers the bundle from the outer circumferential side; and a volute piece which is provided on an inner surface of the casing main body in the discharge flow channel.

Abstract: A propulsive assembly for an aircraft includes a nacelle and a turbojet. The nacelle includes an upstream air inlet section. The turbojet includes a fan with removable blades. The air inlet section extends longitudinally. The fan includes a rotary disk including an annular peripheral wall and a plurality of curvilinear recesses positioned circumferentially in the wall. Each blade includes a blade body extending between a blade root and a blade head in a blade axis which is roughly radial relative to a fan rotation axis. Each blade root includes a curvilinear attachment housed in a respective recess of the rotary disk. Each attachment has a support surface forming an arc. The function of the support surface is to hold the attachment in the recess. The arc extends in or parallel to an attachment plane containing the blade axis or forming with the blade axis an angle of less than 15°.

Abstract: A reverse flow gas turbine engine has a low pressure (LP) spool and a high pressure (HP) spool arranged sequentially in an axial direction. The LP spool comprises an LP compressor disposed forward of an LP turbine and drivingly connected thereto via an LP compressor gear train. The HP spool comprises an HP compressor in flow communication with the LP compressor, and an HP turbine disposed forward of the HP compressor and drivingly connected thereto via an HP shaft.

Abstract: The invention relates to an autogyro having a rotor (12). According to the invention, a gas pressure spring (32) is provided and is arranged for trimming the rotor (12).

Abstract: A locking spacer, which is fitted in a dovetail slot provided on an outer circumferential surface of a disk put on a rotor shaft, includes: a pair of first blocks each provided with a dovetail joint and a stepped seating surface with a first bolt hole, and configured to have a size occupying a portion of an internal space of the dovetail slot; a second block having a size to be inserted into a remaining portion of the internal space of the dovetail slot, and having a height corresponding to the seating surfaces; a fixing plate seated on both the seating surfaces of the first blocks and an upper surface of the second block, and provided with second bolt holes corresponding to the first bolt holes; and a bolt screwed into the first bolt hole through an associated second bolt hole.

Abstract: A compressor blade locking device can include: a first support including a first locking groove; a second support including a left groove and a first protrusion corresponding to the first locking groove; a third support including a second locking groove; a fourth support including a right groove and a second protrusion corresponding to the second locking groove; a center support disposed between the second support and the fourth support; and a tab disposed in the center support, wherein the tab is configured to be turned such that the tab is disposed in the left groove and the right groove.

Abstract: Turn caps for airfoils of gas turbine engines including cavity sidewalls, a first turn cap divider extending between the cavity sidewalls and defining a turning cavity between the first turn cap divider and the cavity sidewalls, and a second turn cap divider disposed radially inward within the turning cavity. A first turning path is defined between the first turn cap divider and the second turn cap divider and a second turning path is defined radially inward of the second turn cap divider and a merging chamber is formed in the turn cap wherein fluid flows through the first turning path and the second turning path are merged, the merging chamber, the first turning path, and the second turning path forming the turning cavity.

Abstract: A device for adjusting a gap between the housing of an impeller and the impeller in a radial compressor of a turbomachine is provided. The device comprises at least one mechanical setting device by means of which the housing of the impeller can be axially adjusted in a targeted manner for the purpose of adjusting the gap, wherein the at least one setting device can be actuated via an adjusting ring, and a transmission of a force and/or a torsional moment from the adjusting ring can be effected via at least one adjustment lever device, in particular a thread lever, or a tilting lever.

Abstract: A compressor blade locking device can include: a first support including a first locking groove; a second support including a second locking groove; an upper plate disposed in the first locking groove and the second locking groove; and a plurality of bolts passing through the upper plate and coupled with the first support and the second support. The compressor blade locking device can further include a center support interposed between the first support and the second support.