Abstract: An apparatus and method of reducing a flow of fluid and heat between a first space and a second space in a rotatable machine and an integral seal and heat shield device are provided. The device includes an annular flange configured to couple to the rotating member of the rotatable machine and a multi-walled seal shield member extending axially from the flange. The multi-walled seal shield member is formed integrally with the flange. The seal shield member includes a first wall including a plurality of surface features, a second wall spaced radially inwardly with respect to the first wall, and a cavity formed between the first and second walls. The integral seal and heat shield device also includes a cap end integrally formed and configured to seal the first and second walls. Each of the flange, the seal shield member, and the cap end are formed of a sintered metal.

Abstract: A tangential on-board injection (TOBI) system for gas turbine engines includes a first endwall and a second endwall radially spaced apart from the first endwall. A plurality of airfoils extend radially from the first endwall to the second endwall. Each airfoil has a leading edge and a trailing edge. Each airfoil includes a first fillet portion transitioning between the airfoil and the first endwall and a second fillet portion transitioning between the airfoil and the second endwall. The radius of curvature of the first and second fillet portions proximate to the leading edge of the airfoil is larger than the radius of curvature of the first and second fillet portions proximate to the trailing edge.

Abstract: The present invention generally relates to a vane for a gas turbine, and more in particular it provides an innovative vane with improved flexibility leading to a reduction of stresses at the transition from the vane trailing edge to the vane platform, without interfering into the cooling scheme of such component. The present invention can increase flexibility of the vane platform by introducing on the vane platform a material cutback confined in the proximity of the trailing edge portion of the vane airfoil.

Abstract: A recirculation seal for use within a gas turbine engine. The recirculation seal includes a first seal base, including a first seal base axis. The recirculation seal further includes a second seal base, including a second seal base axis The recirculation seal further includes a resilient bulb member coupled to the first seal base. The resilient bulb member includes an exterior bulb wall and an interior bulb wall, wherein the interior bulb wall defines an interior space.

Abstract: A turbomachine, particularly an aircraft engine, having a rotor, which is mounted rotatably around its longitudinal axis in a stator, and which has at least one row of rotating blades, which is formed by a plurality of rotating blades, is disclosed, wherein the stator has at least one abradable layer, and wherein each rotating blade has a blade tip that is lowered radially inward, at least in sections, proceeding from a leading edge on the side of the rotating blades in the direction of a trailing edge on the side of the rotating blades, and with a blade tip region that extends downstream from the leading edge runs into the abradable layer during operation of the turbomachine.

Abstract: An airfoil according to an exemplary aspect of the present disclosure includes, among other things, an airfoil section having an external wall and an internal wall. The internal wall defines a first reference plane extending in a spanwise direction and through a thickness of the internal wall. A first cavity and a second cavity are separated by the internal wall. A plurality of crossover passages within the internal wall connects the first cavity to the second cavity. The plurality of crossover passages are arranged such that the passage axis of each of the plurality of cooling passages intersects a surface of the second cavity.

Abstract: An airfoil for a turbomachine such as a gas turbine engine includes a pressure side and a suction side, a flow-distributing forward wall and an inner cooling wall. A switchbacked passage extends through the airfoil, and the flow-distributing forward wall has a plurality of jetting orifices formed therein and connecting the switchbacked passage to a forward cavity. The jetting orifices are oriented to produce wall jets of cooling fluid directed in an upstream direction toward a back side of the leading edge. A second passage extends between the forward cavity and outlets of the airfoil, so as to convey cooling fluid in a downstream direction toward the outlets.

Abstract: An airfoil for a turbomachine such as a gas turbine engine includes a switchbacked passage for conveying cooling fluid, and flow-distribution passages in a forward wall so as to direct cooling fluid from the internal passage for back side impingement upon a leading edge of the airfoil. An inner cooling wall forms a passage for discharging the spent cooling air from a trailing edge of the airfoil after impingement, and cools one of a pressure side and a suction side of the airfoil by way of conduction. The switchbacked passage may have a serpentine form.

Abstract: A turbocharger has a shaft with a rotational axis, a radial/axial turbine wheel which is arranged in a turbine housing and which is connected to the shaft in a non-rotatable manner, and a bearing housing which adjoins the turbine housing and which contains a lateral wall facing the turbine housing. A sub-region of the lateral wall of the bearing housing forms a sub-region of the rear wall of the turbine housing. The sub-region of the bearing housing has two sub-sections, one of which runs diagonally to the rotational axis in an inflow direction of an exhaust gas flow conducted into the turbine housing and the second of which runs in a radial direction relative to the rotational axis of the shaft and parallel to the rear wall of the turbine wheel. The two sub-sections are connected to each other via an exhaust gas flow separation edge of the bearing housing.

Abstract: Horizontal axis W/T of low wind speeds of propeller type, bearing a main rotor of three blades (1), while in the space between two successive blades (1) and diametrically opposite to the third, an additional 4th blade (2) also of propeller type but of significantly longer length, is interposed. This additional blade (2) is not permanently coupled but selectively engaged in the system of the W/T at low wind speeds, contributing to the startup and enhancing the energy production. The blade (2) bears diametrically opposite a counterweight (15) to balance the forces developed, and rotates in a plane parallel to the main rotor. The coupling of the blade (2) is preferably made at the stand-by state or at low wind speed operation of the W/T, while the uncoupling will be performed during operating state and at the rated power. The blade (2) after uncoupling, gets in vertical position and remains immobilized attached to the tower.

Abstract: A rotor for a turbine has a plurality of rotor components lined up in the axial direction and connected by a connecting rod, wherein a groove which extends in the circumferential direction and is open in the axial direction is disposed on one of the rotor components, wherein a coupling element running around the connecting rod in order to support the connecting rod is disposed in the groove. The rotor is adapted to enable particularly stable support of the connecting rod in order to prevent vibrations. The coupling element is disposed on a retaining element connected to the connecting rod.

Abstract: A rotorcraft rotor having each of its blades fitted with a lead/lag damper. A connection device between the blade root of a blade under consideration and a lead/lag damper comprises a rod passing through a laminated ball joint secured to a distal endpiece of the lead/lag damper. The rod extends in the general plane in which the blade extends, being prevented from moving firstly in rotation by being embedded in the blade root and secondly in translation along the general transversely-extending direction of the blade by being pinned to a U-shaped yoke fastened to the blade root.

Abstract: The invention relates to a compressor blade, defined at each one of the points of its surface by a sweep angle and a dihedral angle, comprising: a root, a tip, the distance between the root and the tip, measured along an axis referred to as the radial axis, perpendicular to an axis of rotation of the compressor, being referred to as radial height (h), a zone, between the root and the tip, of which a first portion has a strictly positive leading edge dihedral angle, and a second portion has a strictly negative leading edge dihedral angle, the zone of maximum dihedral angle being, along the radial axis, between r=0.25 h and r=0.7 h.

Abstract: A stator vane shiplap assembly is provided. The stator vane shiplap assembly may comprise a plurality of stator clusters, coupled together to form an annular shape. Each stator cluster may comprise a shiplap stator shroud and a plurality of stator vanes. The shiplap stator shroud may comprise a female end and a male end, with each end comprising a complimentary forward shiplap surface and outward shiplap surface. In response to a coupling of adjacent shiplap stator shrouds, the female forward shiplap surface and the female outward shiplap surface may form a compound shiplap seal joint with the male forward shiplap surface and the male outward shiplap surface. A feather seal may also be introduced between the coupling of adjacent shiplap stator shrouds.