An apparatus and method for a turbine engine for can include an engine component. The engine component can include an interior cooling passage at least partially defining a cooling circuit for passing a flow of cooling fluid through the component. Film holes provide for exhausting a portion of the cooling fluid to an exterior of the component, to form a cooling film along an exterior hot surface of the engine component. A deflector can be position within the cooling passage upstream form the film hole.

A turbine blade including an airfoil portion extending between a leading edge and a trailing edge, a base portion positioned below the airfoil portion, the base portion including an outwardly-extending wing positioned below the airfoil portion, and an aero-brake feature positioned between the outwardly-extending wing and the airfoil portion and extending outward from the base portion, where the aero-brake feature is structurally configured to disrupt axial airflow across the turbine blade.

A turbocharger includes a first scroll member made of sheet metal, extending in a circumferential direction of an impeller shaft, having an opening that is opened to the bearing housing, and forming a part of a wall surface of the first turbine scroll passage, a second scroll member made of sheet metal, extending in the circumferential direction of the impeller shaft, having an opening that is opened to the bearing housing, and forming a part of a wall surface of the second turbine scroll passage, and a closing member closing the opening of the first scroll member and forming the wall surface of the first turbine scroll passage on the side thereof adjacent to the bearing housing, and closing the opening of the second scroll member and forming the wall surface of the second turbine scroll passage on the side thereof adjacent to the bearing housing.

A vane includes a ceramic airfoil section that has an airfoil wall defining a leading edge, a trailing edge, a pressure side, and a suction side. The ceramic airfoil section has an internal cavity. A support spar extends through the internal cavity for supporting the ceramic airfoil section. The support spar is spaced from the airfoil wall such that there is a gap there between. The support spar has an internal through-passage that is fluidly isolated from the gap in the ceramic airfoil section. A baffle is disposed in the gap and is spaced apart from the airfoil wall and the support spar so as to divide the gap into a plenum space between the support spar and the baffle and an impingement space between the baffle and the airfoil wall. The baffle has impingement holes directed toward the airfoil wall that connect the plenum space and the impingement space.

A gas turbine engine comprising: a shaft about an axis; a first turbine assembly mounted to the shaft, a first flow path and a second flow path extending through first turbine assembly along the axis. The second flow path is located radially inward of the first flow path relative to the axis. A second turbine assembly is about the axis downstream of the first turbine assembly, with a gap defined between the first turbine assembly and the second turbine assembly, the gap in fluid communication with the first flow path and the second flow path. A washer is downstream of the first turbine assembly. The washer has an annular body including a deflector between the first turbine assembly and the second turbine assembly, the deflector obstructing the first flow path and extending toward the second flow path.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk and a plurality of rotor blades. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is mounted to the first rotor disk and to the second rotor disk. The rotor blades include a first rotor blade. The first rotor blade includes an attachment projecting axially along the rotational axis into a first pocket in the first rotor disk and a second pocket in the second rotor disk. The attachment has a dovetail cross-sectional geometry when viewed in a plane perpendicular to the rotational axis. A portion of the first rotor disk extends circumferentially across and covers the attachment.

Exhaust frequency mitigation apparatus, exhaust diffusers, and turbomachines are provided. An exhaust frequency mitigation apparatus includes a main body extending along an axial centerline from a base to a tip. the base defines a first diameter and the tip defining a second diameter. the main body converges from the first diameter to the second diameter with respect to the axial centerline. The base is configured to extend from an inner shell of a turbomachine exhaust diffuser. The exhaust frequency mitigation apparatus further includes at least one rib extending from a root coupled to the main body to a free end.

A seal system for a bearing chamber (22) of a turbomachine includes the baffle element (70) encircling the axis (25) of the machine. The baffle element is formed with front surface including both a recess (73) which defines a circumferentially-extending oil-receiving channel (74), and an oil-deflecting surface (78) on a gutter (77). The channel decreases in cross-sectional area close to the oil-deflecting surface, and a circular line which is within the channel distant from the gutter, intersects with the oil-deflection surface. This forces the oil to change direction at the gutter, and urges the oil radially outward. Thus, a high rotational velocity of the body of oil may be maintained, whilst improving the drainage efficiency of the seal system at the gutter.

A turbine vane assembly adapted for use with a gas turbine engine includes an airfoil and a spar. The airfoil is formed to define a cavity that extends into the airfoil. The spar is located in the cavity to define a cooling passage that extends around the spar between the spar and the airfoil. The turbine vane assembly includes cooling features to aid heat transfer of the turbine vane assembly during operation in the gas turbine engine.

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.