The present invention provides a bladed rotor wheel for a gas turbine engine comprising at least a rotatable forged disc, the rotatable forged disc comprising a front surface and a back surface, at least one rim surface, and a plurality of projections located on at least a portion of at least one of the front or back surface and/or on the rim surface; wherein the projections are 3D printed features protruding outwards from the front, back and/or rim surface; the projections are arranged forming a pattern so that a heat transfer capability is created at the front, back and/or rim surface; and the ratio of the distance between projections to the forged disc external radius is lower than 0.15. Furthermore, the present invention also provides a method for manufacturing a rotatable forged disc for a bladed rotor wheel.

A turbine engine system includes a gutter and a gear train with an axial centerline. The gutter is disposed radially outside of the axial centerline. The gutter at least partially circumscribes the gear train, and includes an inner surface and a channel. The channel receives fluid directed out of the gear train. The channel extends radially into the gutter from the inner surface to a channel end, and circumferentially to a channel outlet. At least a portion of the channel has a cross-sectional channel geometry that tapers axially as the channel extends radially towards the channel end.

An apparatus and method regarding an airfoil for a turbine engine, the airfoil comprising an outer wall defining an interior bound by a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction. The airfoil further comprising at least one cooling passage extending radially within the interior and defining a primary cooling airflow; and at least one cooling hole having an inlet defining a first cross-sectional area, the inlet in communication with the cooling passage, an outlet defining a second cross-sectional area greater than the first cross-sectional area; wherein the primary cooling airflow enters the inlet in a first direction and exits the outlet in a second direction different than the first direction.

A turbine engine system includes a gutter and a gear train with an axial centerline. The gutter is disposed radially outside of the axial centerline. The gutter includes an inner surface and a channel that receives fluid directed out of the gear train. The inner surface at least partially defines a bore in which the gear train is arranged. The channel extends radially into the gutter from the inner surface, and circumferentially to a channel outlet. The bore has a cross-sectional bore area, and the channel has a cross-sectional channel area that is substantially equal to or less than about two percent of the bore area.

A strut structure with a strip for an exhaust diffuser of a gas turbine and a gas turbine having the same are provided. The strut structure with a strip for an exhaust diffuser of a gas turbine is configured to include a plurality of struts disposed along an outer circumference of a diffuser body disposed on a central side of the exhaust diffuser, and one or more strips formed on the strut, wherein an exhaust gas passing through the strut flows along the strip from a leading edge of the strut to alleviate a separated flow phenomenon, and wherein if the exhaust gas enters the strip, corner vortices are formed on the leading edge of the strut, and if the exhaust gas flows along the strip, streamwise vortices are formed to alleviate a pressure loss of the exhaust gas.

A gas turbine blade has a trench part of a film cooling unit for cooling a blade part that is formed at a tip of a film cooling hole part. As a result, cooling efficiency of the blade can be improved since the blade part is sufficiently cooled even during introduction of a large amount of cooling air, durability of the blade can be increased by inhibiting the blade from being damaged due to hot gas since the trench part is formed to have a minimum width, and efficiency of a gas turbine can be increased by an improvement in film efficiency.

A core element of an investment core for use in a casting process used to produce an airfoil includes an investment core body, an extension connected to and protruding from the investment core body, and a connection portion connected to the investment core body and to the extension. The investment core body comprises a ceramic material. A shape of the extension comprises a tube with a centerline axis passing through a center of the extension. A shape of a cross-section of the extension taken along a plane perpendicular to the extension centerline axis comprises an ellipse. The extension is connected to the investment core body by the connection portion.

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 lubricant nozzle for a planetary gear set speed reducer of a turbomachine, the nozzle having a generally elongate shape and including a body with a longitudinal axis B, the body having a longitudinal inner cavity in fluid communication with a lubricant inlet located at a longitudinal end of the body and with lubricant outlet apertures that are provided in an annular wall of the body and that extend substantially radially relative to the axis B, wherein the apertures are formed in at least one boss of the body, which boss projects radially outwards on the wall and extends, about the axis, at an angular extent that does not exceed 180.

A blade (10) for a turbomachine is provided. The blade (10) includes an airfoil (16) having at least one thickening (22) on a pressure side (20) of the airfoil (16). The thickening (22) has a region (B) of constant thickness (D). A turbomachine, a compressor and a turbine having stator vanes and/or rotor blades including at least one such blade (10) is also provided.