Disclosed are a blade ring segment for a turbine section, a turbine section having the blade ring segment, and a gas turbine having the turbine section. Multiple blade ring segments is installed in a turbine casing accommodating turbine blades rotated by combustion gas from a combustor. The blade ring segment includes an inner panel provided in the turbine casing and having multiple air holes through which cooling air fed from the outside of the turbine casing flows, an outer panel disposed on one side of the inner panel, and a cooling structure protruding from one side of the outer panel so as to form a flowing channel in a zigzag pattern so that cooling air fed through the air holes flows therethrough.

At least one of a variable turbine geometry and a variable compressor geometry for an exhaust gas turbocharger may include a housing including a first housing wall and a blade bearing ring having at least one guide blade rotatably mounted thereon. A control lever may be included for adjusting the at least one guide blade between a closing position and an opening position. An actuating shaft may be connected to the control lever in a rotationally fixed manner along a rotation axis. The actuating shaft may be rotatably mounted on the housing via a passage opening disposed in the first housing wall. The actuating shaft may directly support itself on the first housing wall in the passage opening.

At least one of a variable turbine geometry and a variable compressor geometry for an exhaust gas turbocharger may include a housing including a first housing wall and a blade bearing ring having at least one guide blade rotatably mounted thereon. A control lever may be included for adjusting the at least one guide blade between a closing position and an opening position. An actuating shaft may be connected to the control lever in a rotationally fixed manner along a rotation axis. The actuating shaft may be rotatably mounted on the housing via a passage opening disposed in the first housing wall. The actuating shaft may directly support itself on the first housing wall in the passage opening.

A blade of a turbomachine is provided, including at least one cooling channel in the interior of the blade for cooling the blade with the aid of a fluid flowing through the cooling channel, the cooling channel having at least one inlet and at least one outlet, between which the cooling channel extends along its longitudinal axis, and the cooling channel being radially delimited by at least one wall, at least one displacement body being situated in the cooling channel, so that an annular or tubular gap between the displacement body and the wall of the cooling channel results in the area of the displacement body/bodies, which is available for the through-flow of the fluid, or at least two or multiple subchannels being formed in the area of the displacement body/bodies. The invention also relates to a method for manufacturing a corresponding blade.

A gas turbomachine includes a compressor portion, a turbine portion operatively connected to the compressor portion, and a combustor assembly fluidically connected to each of the compressor portion and the turbine portion. A turbine casing includes a body having an outer surface and an inner surface. A clearance control system includes a plurality of fluidically connected fluid channels extending through the turbine casing. The plurality of fluidically connected fluid channels includes a first fluid channel configured to direct a fluid flow in a first axial direction, a circumferential fluid channel configured to direct the fluid flow in a circumferential direction, and a second fluid channel configured to direct the fluid flow in a second axial direction substantially opposite the first axial direction. The first fluid channel includes a first outlet passing through the inner surface, and the second fluid channel including a second outlet passing through the inner surface.

A conical hub for a fan of a gas turbine engine is provided. The conical hub having: a plurality of attachment features located on an outer circumferential surface of the conical hub, wherein at least some of the plurality attachment features are axially aligned with each other and at least some of the plurality of attachment features are off set from each other, and wherein each of the plurality of attachment features have an opening configured to receive a portion of a pin; and the outer circumferential surface of the conical hub increases in diameter with respect to an axis of the conical hub in a forward to aft direction of the conical hub.

In one embodiment, a hub for a fan of a gas turbine engine is provided. The hub having: a plurality of attachment features located on an outer circumferential surface of the hub, wherein at least some of the plurality attachment features extend radially away from the outer circumferential surface and are axially aligned with each other and at least some of the plurality of attachment features extending radially away from the outer circumferential surface and are off set from each other, and wherein the plurality of attachment features have an opening configured to receive a portion of a pin; and wherein at least some of the plurality of attachment features are located on a forward leading edge of the hub.

A single degree of freedom (SDOF) acoustic liner includes a porous face sheet, a substantially imperforate back sheet generally parallel to and opposing said porous face sheet and defining a thickness therebetween, and an acoustic core layer of contiguous adjacent resonant cavities disposed between the porous face sheet and the imperforate back sheet. The acoustic core layer includes a first resonant cell having a first internal volume therein and a second resonant cell having a second internal volume therein different than the first internal volume. A cell partition wall extends between the porous face sheet and the imperforate back sheet, and separates and seals the first resonant cell from the second resonant cell. In a thickness direction, and perpendicular to a plane generally parallel with the porous face sheet and the substantially imperforate back sheet, the first internal volume overlaps the second internal volume over the cell partition wall.

An abradable structure (10) for a turbomachine (40) that is designed to be at least partially deformed and/or at least partially abraded by at least one abrading element (44, 47) of the turbomachine (40) during operation thereof; the at least one abradable structure (10) being formed at least regionally of structural elements (12, 14, 16, 18) that have a respective polygonal cross section (20, 22) that is oriented in a rub direction (X_A) of the abradable structure (10). The structural elements (12, 14, 16, 18) have a greater extent (X.sub.h, X.sub.r) in the rub direction (X_A) than in a direction of the abradable structure (10) that is orthogonal to the rub direction (X_A). Other aspects of the present invention relate to a turbomachine (40) having an abradable structure (10), as well as to a method for manufacturing an abradable structure (10).

An axial compressor of a turbine engine includes a plurality of disk and spacer pairs oriented along a common axis of rotation. Each of a disk and a spacer of the disk and spacer pairs has a contacting face defining an engagement between the disk and the spacer. The contacting face of each of the disk and the spacer includes a recessed area. A pin has a stem received within the recessed area of the disk and a head received within the recessed area of the spacer. The head of the pin includes at least two flats corresponding to complementary surfaces of the recessed area of the spacer.