Abstract: An airfoil for use in a turbomachine includes a pressure sidewall and a suction sidewall coupled to the pressure sidewall. The suction sidewall and the pressure sidewall define a leading edge and a trailing edge, the leading edge and the trailing edge define a chord distance. The airfoil includes a tip portion extending between the pressure sidewall and the suction sidewall. The tip portion includes a planar section and a recessed section. The recessed section extends adjacent to the planar section such that a thickness of the planar section is less than a thickness of the airfoil. The recessed section is offset a predetermined distance from the leading edge and the trailing edge along the chord distance.

Abstract: A centrifugal compressor includes a casing, an impeller, a motor, a diffuser, a magnetic bearing, and a magnetic bearing backup system including at least one dynamic gas bearing and at least one hydrostatic gas bearing. The impeller is attached to a shaft rotatable about a rotation axis. The motor is configured and arranged to rotate the shaft in order to rotate the impeller. The magnetic bearing rotatably supports the shaft. The magnetic bearing backup system is configured and arranged to support the shaft when the magnetic bearing stops operating. The at least one dynamic gas bearing and the at least one hydrostatic gas bearing of the magnetic bearing backup system are disposed radially inwardly relative to the magnetic bearing.

Abstract: A device for attaching manifolds for cooling the casing of a preferably low-pressure turbine of a turbine engine by air jets. The device includes a mounting for the manifolds, shaped such as to keep the manifolds spaced apart from one another. A plurality of elements support the manifold mounting, each supporting element being attached to the casing, and connected to the manifold mounting by a connector. The device includes N cooling manifolds and N-1 supporting elements, each supporting element being arranged between two adjacent cooling manifolds.

Abstract: An airfoil assembly for use in a gas turbine engine is disclosed. The airfoil assembly includes a root, a tip shroud, and an airfoil located between the root and tip shroud. The airfoil includes a spar, a heat shield arranged around the spar, and a plurality of describe rigid load supports extending between the spar and the heat shield to transfer loads between the heat shield and the spar.

Abstract: A rotor system includes a yoke, a plurality of blade grip assemblies and a plurality of centrifugal force bearings coupling the blade grip assemblies with the yoke. A plurality of rotor blades are coupled to the blade grip assemblies such that each rotor blade has a coincident hinge and such that each rotor blade has three independent degrees of freedom including blade pitch about a pitch change axis, blade flap about a flapping axis and lead-lag about a lead-lag axis. A blade-to-blade damping ring includes a plurality of damper anchors each coupled to one of the blade grip assemblies along the respective pitch change axis and a plurality of lead-lag dampers each coupled between adjacent damper anchors. During blade pitch operations, each blade grip assembly is operable to rotate relative to the respective damper anchor, such that the blade-to-blade damping ring is operable to provide pitch independent lead-lag damping.

Abstract: There is provided a variable nozzle mechanism that can suppress wear. The variable nozzle mechanism includes a plurality of nozzle vanes, a plurality of arms, and a unison ring. The plurality of nozzle vanes are provided in the nozzle path. The plurality of arms each have a pivoted end and a free end, and the pivoted end is integrally coupled with the nozzle vane by means of a pivot. The unison ring is provided with a plurality of grooves. Each groove is engaged with the respective free end of the arm. The free end of the arm has a closing-side contact surface and an opening-side contact surface each in contact with a respective one of paired walls forming the groove of the unison ring. The closing-side contact surface has a curvature different between a side closer to the pivot and a side farther away from the pivot.

Abstract: A rotor blade configured to fold about a blade fold axis includes a first section and a second section. The first section is configured to mount to a rotor hub and includes a first connector having at least one first opening. The second section includes a second connector having at least one second opening. The second section is rotatably coupled to the first section and is configured to rotate about the blade fold axis between an aligned position and a rotated position. A linkage is operably coupled to the first section and the second section. The linkage includes a plurality of links loosely connected such that the linkage is configured to support the second section of the rotor blade only when the second section is in the rotated position.

Abstract: An axial turbine engine for an aircraft, in particular a ducted fan turbine engine. The engine comprises an oil circuit with a cooler fitted with a first heat exchange surface in contact with oil of the oil circuit, and a second heat exchange surface in contact with a secondary air flow entering the turbine engine. The cooler is of the air/oil type (ACOC) and comprises fins. To increase the capacity of the cooler, the cooler is fitted with a device for injecting a cooling liquid, e.g., water with additives. The thermal capacity of the water amplifies the cooling. Evacuation of the water increases the thrust of the turbine engine. A method is also provided for cooling the turbine engine of an aircraft by injection of a cooling liquid at the time of take-off.

Abstract: A centrifugal blower is provided. The centrifugal blower includes a hub, a shaft, a motor, a plurality of blades, a rib, and a first fin. The shaft is connected to the hub. The motor rotates the shaft. Each blade includes a rib and a first fin. The rib is connected to the hub, wherein the rib extends from the hub to an end of the blade. The first fin is disposed on a first side of the rib and connected to the hub, wherein the first fin includes a first surface, the rib protrudes from the first surface, and the thickness of the first fin is less than the thickness of the rib.

Abstract: An airfoil of a gas turbine engine is provided including a leading edge extending in a radial direction, a tip extending in an axial direction from the leading edge, a first rib extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity within the airfoil, a second rib, the second rib and the first rib defining a serpentine cavity therein, a third rib extending axially within the tip, a flag tip cavity defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity, and a bypass aperture formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity.

Abstract: A turbocharger assembly includes a compressor wheel with a base surface, a nose surface, a z-plane disposed between the base surface and the nose surface and a bore extending from the base surface to the nose surface and a shaft that includes a first pilot surface disposed in the bore of the compressor wheel at a position between the z-plane and the nose surface, a second pilot surface disposed in the bore of the compressor wheel at a position between the z-plane and the base surface, and a recessed surface disposed between the first pilot surface and the second pilot surface. A nut adjustably disposed on the shaft adjacent to the nose surface can tension the shaft to apply a compressive load between the base surface and the nose surface of the compressor wheel. Various other examples of devices, assemblies, systems, methods, etc., are also disclosed.

Abstract: A blade for a turbomachine, in particular a compressor or turbine stage of a gas turbine, having at least one matrix having a first impact chamber (10) in which at least one impulse element (11) is disposed with play, at least one second impact chamber (20) whose volumetric centroid is offset from a volumetric centroid of the first impact chamber (10) along a first matrix axis (A) and in which at least one impulse element (21) is disposed with play, and at least one third impact chamber (30) whose volumetric centroid is offset from the volumetric centroid of the first impact chamber (10) along a second matrix axis (B) transversely to the first matrix axis (A) and in which at least one impulse element (31) is disposed with play, the first matrix axis (A) and an axis of rotation (R) of the turbomachine forming an angle of at least 60° and no more than 120° is provided.

Abstract: Wind turbine blades comprising one or more deformable trailing edge sections, each deformable trailing edge section comprising a first and a second actuator, wherein the second actuator is arranged substantially downstream from the first actuator, and wherein the first actuator is of a first type and wherein the second actuator is of a second type, the second type being different from the first type. The application further relates to wind turbines comprising such blades and methods of operating a wind turbine comprising one or more of such blades.

Abstract: A ribbed composite shell includes an annular grid of relatively thick crack arresting ribs embedded in a relatively thin annular shell and relatively thin panels in thin annular shell between arresting ribs wherein each of panels are completely surrounded by a set of relatively thick adjoining ones of ribs. A shell forward flange may extend radially inwardly from thin annular shell. Arresting ribs may include radially stacked layers of strips between radially stacked annular layers of shell. Annular grid may include a rectangular grid pattern, a diamond grid pattern, or a hexagonal grid pattern. A nacelle inlet may have the ribbed composite shell within one or both of radially spaced apart composite inner and outer skins of an inner barrel. Nacelle inlet may be part of attached to a fan casing and axially disposed forward of fan blades circumscribed by the casing. The inlet may be on an engine nacelle.

Abstract: An airfoil for use in a turbomachine includes a pressure sidewall and a suction sidewall coupled to the pressure sidewall. The suction sidewall and the pressure sidewall define a leading edge and an opposite trailing edge. The leading edge and the trailing edge define a chord distance. The airfoil further includes a root portion, and a tip portion. The tip portion extends between the pressure sidewall and the suction sidewall such that the tip portion is substantially perpendicular to each sidewall. The tip portion includes at least one planar section and at least one oblique section that forms a recess within the tip portion. The at least one oblique section extends from the at least one planar section towards the root portion along the chord distance. The tip portion is configured to reduce airfoil wear during contact with a surrounding casing.

Abstract: An airfoil for a gas turbine engine includes an airfoil including spaced apart pressure and suction side walls joined at leading and trailing edges to provide an exterior airfoil surface that extends in a radial direction from a platform to a tip. A cavity is provided between the pressure and suction side walls near the trailing edge. The cavity includes an interior region bounded by first and second exit regions arranged at angle relative to one another. The first and second exit regions are respectively in low and high pressure regions relative to one another. First and second pedestal groups respectively are arranged at the first and second exit regions. The second pedestal group has first and second pedestals each terminating in an end. The ends of the second pedestals extend beyond the ends of first pedestals.

Abstract: A vibration-proofed fan frame includes a housing and a base surrounded by the housing. A hollow tube is arranged on the base and has the rotary connection with an impeller of the heat dissipation fan. The base includes a disc and a plurality of ribs extending outwards from periphery of the disc. The fan frame further includes a buffer member within the structural interconnections between the disc and the housing. The buffer member includes first elements. The first elements are embedded in the disc. A heat dissipation fan having the fan frame described is also presented.

Abstract: The invention relates to a fan assembly having a rotatingly driven hub (7) and fan blades (10) which can be pivoted relative to the hub body (29). The fan blades (10) are adjusted by means of a linear drive in the form of a spindle drive (13) or a piston drive (52).

Abstract: One exemplary embodiment according to this disclosure relates to a system includes a blade outer air seal (BOAS), and a meter plate. A portion of the meter plate is provided radially outward of a radially outermost surface of the BOAS.

Abstract: A turbine rotor blade is provided including an airfoil having an airfoil shape. The turbine rotor blade airfoil having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I wherein the Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances in inches by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil in inches, and wherein X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z, the airfoil profile sections at Z distances being joined smoothly with one another to form a complete airfoil shape.