Abstract: A turbine assembly for use with a gas turbine engine includes a bladed wheel assembly and a vane assembly. The bladed wheel assembly is adapted to interact with gases flowing through a gas path of the gas turbine engine. The vane assembly is located upstream of the bladed wheel assembly and adapted to direct the gases at the bladed wheel assembly.

Abstract: A curved contour of the lateral surface of a blade arrangement includes in at least one meridian section on mutually opposite sides of a blade airfoil an intersection point that is closer to the blade airfoil front edge, and an intersection point that is closer to the blade airfoil rear edge, and a best-fit line of least square distances from the curved contour. The curved contour includes first and/or the second contour section which meet specified conditions.

Abstract: A wind turbine apparatus has a tower and a plurality of blades rotatably mounted relative to the tower. Each of the plurality of blades has a reflective media on a surface thereof. The reflective media has at least a pair of primary colors. Each primary color is different from each other. The reflective media on one of the plurality of blades is arranged in an order different than in order of the reflective media of an adjacent blade. The reflective media can further include a mixed color of white. The reflective media comprises bands extending across a surface of the blade.

Abstract: A system for retaining stators and reducing air leakage in a gas turbine engine having an axis includes a stator having an inner platform, an outer platform, a low pressure side, a high pressure side, and at least one foot, and designed to turn air. The system also includes a case positioned radially outward from the stator and having at least one recess designed to interface with the at least one foot to resist movement of the stator relative to the case. The system also includes a bladder positioned between the outer platform of the stator and the case and designed to receive pressurized fluid having a greater pressure than ambient pressures experienced at the low pressure side of the stator and to further resist movement of the stator relative to the case in response to receiving the pressurized fluid.

Abstract: Methods for forming protective coatings on aerospace components are provided. In one or more embodiments, the method includes exposing an aerospace component to a first precursor and a first reactant to form a first deposited layer on a surface of the aerospace component by a first deposition process (e.g., CVD or ALD), and exposing the aerospace component to a second precursor and a second reactant to form a second deposited layer on the first deposited layer by a second deposition process. The first deposited layer and the second deposited layer have different compositions from each other. The method also includes repeating the first deposition process and the second deposition process to form a nanolaminate film stack having from 2 pairs to about 1,000 pairs of the first deposited layer and the second deposited layer consecutively deposited on each other.

Abstract: A rotor stack for a gas turbine engine includes a first rotor disk with a first rotor spacer arm, the first rotor spacer arm having a first flange with an outboard flange surface and an inboard flange surface, a first hole along an axis through the first flange, the first hole having a counterbore in the outboard flange surface; a second rotor disk with a web having a second hole along the axis; a third rotor disk with a third rotor spacer arm, the third rotor spacer arm having a third flange with an outboard flange surface and an inboard flange surface, a third hole along the axis through the third flange, the third hole having a counterbore in the inboard flange surface; and a bushing with a tubular body and a flange that extends therefrom, the tubular body comprising at least one axial groove along an outer diameter thereof, the bushing extends through the first hole, the second hole and partially into the counterbore in the inboard flange surface of the third hole.

Abstract: An illustrative example embodiment of a compressor includes an inlet defining an intake passage, a plurality of lateral inlet guide vanes in the intake passage, and a plurality of medial inlet guide vanes in the intake passage. The lateral guide vanes are selectively oriented to alter an amount of fluid flow through a first, lateral portion of the intake passage. The medial inlet guide vanes are selectively oriented to alter an amount of fluid flow through a second, medial portion of the intake passage.

Abstract: A blade for a turbomachine includes a blade root portion for securing the blade to a rotatable annular outer drum rotor. The blade root portion includes one or more radial retention features for radially retaining each of the blade root portions within the rotatable annular outer drum rotor. Further, at least one of the radial retention feature(s) includes at least one sealing member integrated therewith.

Abstract: A propeller fan includes a hub that has a side surface around a central axis, and blades that are provided on the side surface, wherein a blade of the blades includes an inner peripheral portion located on a side of a base, and an outer peripheral portion located on a side of an outer edge, the outer peripheral portion is formed as one blade, the inner peripheral portion includes blade elements arranged at a predetermined interval, a ratio r/R of a radius r which is a distance from the central axis to the outer peripheral portion and a radius R which is a distance from the central axis to the outer edge is 0.4 or less, and when a wind speed at the outer peripheral portion is V1 and a wind speed at the inner peripheral portion is V2, a relational formula of V1?V2×2.0 is established.

Abstract: The invention relates to a side-channel machine having a housing (4a), located in the housing (4a) a side-channel (28) for guiding a gas, and at least one gas inlet opening (34) which is formed in the housing (4a) and is fluidically connected to the side-channel (28). Furthermore, the side-channel machine has at least one gas inlet pipe (29a) which connects to the at least one gas inlet opening (34), The side-channel machine further comprises at least one gas outlet opening (33) and at least one gas outlet pipe (31a) which connects to the at least one gas outlet opening (33). Furthermore, the side-channel machine has an impeller that can be made to rotate in the housing (4a), with impeller blades, which bound impeller cells arranged in the side-channel (28), for delivering the gas in the impeller cells from the at least one gas inlet opening (34) to the at least one gas outlet opening (33).

Abstract: The present invention relates to an adjustment mechanism (100) for variably adjusting the cross-section of a compressor inlet (22) and further relates to a corresponding compressor (20) including such an adjustment mechanism (100). The adjustment mechanism (100) comprises a plurality of rotatable orifice elements (110) and an actuation ring (120). The actuation ring (120) is mechanically coupled to the plurality of orifice elements 110 such that rotation of the actuation ring 120 causes movement of the orifice elements 110. The movement of the orifice elements (110) thereby adjusts the cross-section of a compressor inlet (22). The adjustment mechanism (100) further comprises a plurality of support members (140) which are arranged axially between the plurality of orifice elements (110) and the actuation ring (120). Additionally, the adjustment mechanism (100) comprises a spring (130), more specific a ring-shaped wave spring.

Abstract: A rotary assembly for an aircraft engine is disclosed. A shaft is mounted for rotation about a central axis and a rotary part is mounted on the shaft for common rotation therewith. The rotary part has an axial end portion extending away from a center of mass of the rotary part, the axial end portion having a radial thickness smaller than that of the rotary part at an axial location aligned with the center of mass. The shaft and the rotary part have an interface defining an interference fit between the shaft and the axial end portion of the rotary part. The interface is located at a distance along the central axis from the center of mass. An undercut is defined in the rotary part. The undercut extends axially from adjacent the interface towards the center of mass.

Abstract: A rotor blade includes an airfoil having an airfoil shape. The airfoil shape has a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances expressed in a unit of distance by multiplying the Cartesian coordinate values of X, Y and Z by a scaling factor of the airfoil in the unit of distance. The X and Y values, when connected by smooth continuing arcs, define airfoil profile sections at each Z value. The airfoil profile sections at Z values are joined smoothly with one another to form a complete airfoil shape.

Abstract: Provided is an electrical protection system allowing transferring to the ground static electricity accumulated onto the blades of a wind generator, and the lightning current when a lightning strikes onto at least one of the blades or at the rotor, including a first discharging unit configured for continuously discharging to the ground the static electricity accumulated onto the blades, and a second discharging unit configured for discharging to the ground the lightning current when a lightning strikes in at least one of the blades or at the rotor.

Abstract: An interdigitated turbine assembly for a gas turbine engine, the interdigitated turbine assembly including a first turbine rotor assembly interdigitated with a second turbine rotor assembly. A first static frame is positioned forward of the first turbine rotor assembly and the second turbine rotor assembly. The first turbine rotor assembly is operably coupled to an inner rotatable component of a gear assembly. The second turbine rotor assembly is operably coupled to an outer rotatable component of the gear assembly. The static structure is connected to the first static frame. A driveshaft is operably coupled to the outer rotatable component. A first bearing assembly is operably coupled to the driveshaft and the first static frame. A second bearing assembly is operably coupled to the first static frame and first turbine rotor assembly. A third bearing assembly is operably coupled to the first turbine rotor assembly and the second turbine rotor assembly.

Abstract: A turbomachinery stator apparatus includes: a compressor casing including a casing wall defining an arcuate flowpath surface and an opposed backside surface, the flowpath surface defining at least two spaced-apart rotor lands; and a stator vane row of stator vanes disposed inside the compressor casing; wherein the casing wall includes at least one hollow structure; and wherein the casing wall is a single monolithic whole, wherein the stator vanes are integrally formed as part of the monolithic whole.

Abstract: A turbine blade for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface extending from a platform in a spanwise direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

Abstract: An impeller includes Z blades, where Z is an integer equal to 5 or more, arranged in a circumferential direction of the impeller and extending radially, and pitch angles between adjacent blades are all different. In terms of an arbitrary pitch angle ?, when a pitch angle ?1 adjacent to the pitch angle ?, and a pitch angle ?2 adjacent to the pitch angle ?, different from the pitch angle ?1, satisfy a relation, ?1<?2, a pitch angle ?1 different from the pitch angle ? adjacent to the pitch angle ?1 and a pitch angle ?2 adjacent to the pitch angle ?2, different from the pitch angle ?, satisfy a relation, ?2<?1.

Abstract: A turbine rotor blade includes: a root portion fixed to a rotor shaft; and an airfoil portion including a pressure surface, a suction surface, and a top surface connecting the pressure surface and the suction surface, with a cooling passage formed inside the airfoil portion. The top surface of the turbine rotor blade includes a leading edge region located on the leading edge side and formed parallel to the rotor shaft, and a trailing edge region adjacent to the leading edge region. The trailing edge region has an inclined surface inclined radially inward toward a trailing edge.

Abstract: A rotor disc for a turbomachine, the disc extending circumferentially about an axis and including a plurality of cavities configured to receive blade roots, each cavity including a downstream radial wall configured to axially block the blade root in the cavity, each downstream radial wall including a channel of ventilation of the cavity, including an inlet orifice which opens into the cavity and an outlet orifice which opens onto a downstream surface of the disc. An assembly for a turbomachine including such a disc and an upstream retention ring and a turbomachine including such an assembly.