Abstract: Composite components and methods for forming composite components are provided. For example, a composite component of a gas turbine engine comprises a composite material, a plurality of piezoelectric fibers, and an anti-icing mechanism. The anti-icing mechanism is in operative communication with the piezoelectric fibers such that the anti-icing mechanism is activated by one or more electrical signals from the piezoelectric fibers. In exemplary embodiments, the composite component is a composite airfoil and the anti-icing mechanism is one or more heating elements. Methods for forming composite components may comprise forming piezoelectric plies comprising piezoelectric fibers embedded in a matrix material; forming reinforcing plies comprising reinforcing fibers embedded in the matrix material; laying up the piezoelectric and reinforcing plies to form a ply layup; and processing the ply layup to form the composite component.

Abstract: A holding device is used with a centripetal air sampling member of a rotor assembly that includes and downstream rotor discs and a centripetal air sampling member. The holding device has an annular support element with a housing configured to receive the air sampling member in a radial orientation, a first lateral extension, and a second lateral extension rigidly attached to the downstream rotor disc, extending substantially following the longitudinal axis and radially arranged externally with respect to the first extension. The holding device also includes a wedging ring axially arranged between a portion of the downstream rotor disc and the support element, the wedging ring being configured to simultaneously urge, under the action of a centrifugal force, the first extension in radial abutment against the second extension and the support element in radial abutment against a portion of the upstream rotor disc.

Abstract: A blade includes an airfoil section extending between leading and trailing edges, first and second opposed sides each joining the leading and trailing edges, and an inner end and a free end. The blade also includes an abrasive tip at the free end of the airfoil section. The abrasive tip includes particles disposed in a matrix material. The matrix material is a polymeric material that has a glass transition temperature greater than or equal to about 225 degrees C. (437 degrees F.). A gas turbine engine and a method of fabricating a blade are also disclosed.

Abstract: A turbine casing has a front casing segment around a turbine longitudinal axis with first and second front casing part-segments, and a rear casing segment, around the longitudinal axis with first and second rear casing part-segments. The rear casing segment is downstream of the front casing segment in the direction of flow and fastened to the front casing segment via a rear vertical flange of the rear casing segment and a front vertical flange of the front casing segment, forming a cross-shaped joint. A protrusion surrounding the longitudinal axis is formed on the front casing segment in a region of the front vertical flange and extends parallel to the longitudinal axis and protrudes from the front vertical flange in the flow direction. A circumferential seal device is on an outer side, facing away from the longitudinal axis, of the protrusion, and seals the protrusion against the rear casing segment.

Abstract: A sealing assembly positioned between a first component having a first interface and a second component having a second interface is provided. The sealing assembly includes a flat bracket fixedly attached to the first interface to define an extension and a floating seal including a body portion, a first U-shaped channel arranged to engage the extension, and a second U-shaped channel inhibiting movement of the floating seal in an axial direction while allowing movement in a radial direction, and the second U-shaped channel allowing movement in the axial direction and inhibiting movement in the radial direction.

Abstract: A turbomachine rotor has a disk carrying vanes, each vane having a blade linked by a platform to a root. For at least one vane, a recess is defined between the platform and the disk, and a vibration damper is mounted in the recess. The vibration damper includes a first structural portion configured to contact the platform of which the vibrations are to be dampened, and a second mass portion configured to dampen these vibrations. The second mass portion is a powder and the first structural portion is a box containing the powder.

Abstract: A compressor blade of a gas turbine includes a compressor blade portion including a plurality of layers; and a carbon fiber reinforcement embedded in the plurality of layers of the compressor blade portion and oriented in a direction of stress fields of the compressor blade when in operation. A method of manufacturing the compressor blade includes preparing a composite material including fiber-reinforced layers; forming a first layer of the composite material to extend in a radial direction of the compressor blade; and stacking a second layer of the composite material on the first layer in an axial direction of the compressor blade. The compressor blade is 3D-printed by forming each composite material layer in a radial direction, which layers are stacked in an axial direction. Fiber reinforcement in the composite compressor blade is oriented in line with the stress fields inherent in the operation of the compressor blade.

Abstract: Compressor components, such as blades and vanes, having an airfoil portion with an uncoated, nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1. X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each Z distance in inches. The profile sections at the Z distances are joined smoothly with one another to form a complete airfoil shape.

Abstract: A heat shield for a gas turbine engine has a main body having a first and second surface, the first surface exposed to a hot working gas, a plurality of walls upstanding from the second surface and an impingement plate. The impingement plate is on top of at least one wall and forms a chamber and has an array of impingement holes. At least one pair of divider walls are formed within the chamber and extend between the impingement plate and the second surface. The first divider wall extends from a first wall towards a second wall, the second divider wall extends from the second wall towards the first wall. The first and second divider walls both extend such that there is no clear line of sight in a perpendicular direction to the first divider wall and/or second divider wall and are spaced apart with respect to the perpendicular direction.

Abstract: Compressor components, such as blades and vanes, having an airfoil portion with an uncoated, nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1. X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each Z distance in inches. The profile sections at the Z distances are joined smoothly with one another to form a complete airfoil shape.

Abstract: A shaft and sleeve assembly of a gas turbine engine includes an outer shaft extending about an engine central longitudinal axis and having a radially inner shaft surface. The radially inner shaft surface has a shaft groove define therein. A sleeve is located radially inboard of the radially inner shaft surface. The sleeve includes a sleeve groove defined in a radially outer sleeve surface. The sleeve groove is formed at an axial location between a first axial end and a second axial end of the sleeve. A retaining element is installed in the shaft groove and the sleeve groove to retain the sleeve relative to the outer shaft. The sleeve includes one or more sleeve openings formed therein, and the sleeve groove is aligned with the one or more sleeve openings.

Abstract: An airfoil vane assembly includes a vane piece defining a first vane platform, a second vane platform, and a hollow airfoil section joining the first vane platform and the second vane platform. The first vane platform defines a collar projection therefrom. A spar piece defines a spar platform and a spar extends from the spar platform into the hollow airfoil section. The spar platform includes a radial opening defined by first and second axial faces. The radial opening is configured to receive the collar projection, and a groove in the first axial face. A seal is situated in the groove. The seal seals against the collar projection and a biasing member is configured to bias the seal towards the collar projection. A method of assembling a vane is also disclosed.

Abstract: Aspects of the disclosure are directed to a seal comprising: a shoe, and at least two beams coupled to the shoe, where a first width associated with the beams exceeds a second width associated with the shoe in a reference direction. Aspects of the disclosure are directed to an engine comprising: a compressor section, a turbine section, and a floating, non-contact seal that includes: a shoe, and at least two beams coupled to the shoe, where the beams extend past an edge of a flowpath surface associated with the shoe in a reference direction.

Abstract: A vertical axis wind turbine (VAWT) with improved and optimized wind-directing, wind-shaping, and wind-power conversion features is disclosed. The shapes of these features directly affect the ability of the VAWT to use the power of moving air, such as wind, to spin a rotor and create torque on a rotor shaft to generate electricity. The wind-power-conversion mechanical efficiency of the invention is significantly improved over previous efforts, to the point that the invention can convert wind energy into electrical power at a price-to-performance ratio that competes with or surpasses existing alternative energy technologies.

Abstract: A turbine rear structure for a gas turbine engine includes a central hub and a circumferential outer ring coaxial with the central hub. The turbine rear structure further includes a plurality of guide vanes extending radially between the central hub and the circumferential outer ring, and an intermediate guide vane located in a space defined between adjacent guide vanes. The intermediate guide vane is located closer to one of the guide vanes than the other guide vane.

Abstract: A vane includes an airfoil that defines a leading edge, a trailing edge, a pressure side, and a suction side. The airfoil includes a truss structure that has ribs that define there between a plurality of through-cavities from the pressure side to the suction side. Honeycomb cells are disposed in the cavities. A face sheet defines at least one of the pressure side or the suction side. The face sheet has perforations that correspond in location to the honeycomb cells.

Abstract: Baffle inserts for airfoils of gas turbine engines are described. The baffle inserts include a baffle insert body having a first side portion and a second side portion, wherein each side portion has a respective end, a first set of vortex generation elements is arranged at the end of the first side portion, and a second set of vortex generation elements is arranged at the end of the second side portion. The first set of vortex generation elements and the second set of vortex generation elements are arranged at an aft end of the baffle insert body.

Abstract: A seal assembly for a gas turbine engine includes a seal including a main body extending circumferentially between opposed mate faces. The main body has a sealing portion and an engagement portion extending outwardly from sealing portion along at least one of the mate faces. The main body has a core including one or more core plies arranged to establish an internal cavity. An overwrap has one or more overwrap plies arranged to follow a perimeter of the one or more core plies to establish the engagement portion and the sealing portion. A platform insert extends between portions of the core and the overwrap to establish the sealing portion. A method of fabricating a seal for a gas turbine engine is also disclosed.

Abstract: A turbine includes an outer casing and an inner casing surrounded by said outer casing. The outer casing and inner casing are coaxially arranged with respect to a machine axis. The outer casing and inner casing are each divided in a split plane into an upper part and a lower part. The upper and lower parts are connected with each other in said split plane by means of a flanged connection. The disassembly process is improved by the flanged connection of the upper and lower part of the inner casing comprising a plurality of bolts, which extend each through through holes in respective flanges of said flanged connection of the inner casing, protrude from said through holes at both ends with a threaded section, and are tightened by means of nuts screwed on said threaded sections at both ends of said bolt.

Abstract: A tip rail cooling insert for attaching into a tip rail pocket in a tip rail of a turbine blade is disclosed. The insert includes a first inner layer defining at least one first insert cooling channel therein, the first inner layer including a pair of spaced legs defining a first coolant collection plenum with at least the tip rail pocket for directing coolant from at least one internal cooling cavity in the turbine blade to the at least one first insert cooling channel. Each of the pair of spaced legs has an angled outer end configured to accommodate rounded inner corners of the tip rail pocket. A first outer layer is on a first side of the first inner layer, and a second outer layer is on a second side of the first inner layer.