A cooling system for a gas turbine is provided that may include a drive shaft, a shield of a combustor of the gas turbine engine, and a conduit. The drive shaft may be configured to mechanically couple a compressor of the gas turbine engine to a turbine of the gas turbine engine. The shield may be positioned between the compressor and the turbine. The combustor may be configured to drive the turbine. The conduit may be configured to supply a cooling fluid to a gap between an outer surface of the drive shaft and the shield of the combustor.

Disclosed is an apparatus for decreasing a thrust of a radial flow turbine. The apparatus includes a rotary shaft having an axial through-hole in the interior thereof, a rotor assembled in the rotary shaft and having a rotor hub and rotor blades formed on an outer peripheral surface thereof, a casing configured to isolate the rotary shaft and the rotor from the outside, and a shaft seal configured to maintain a seal between the rotary shaft and the casing.

An aspirating seal assembly for use in a turbine engine is provided. The aspirating seal assembly includes a face seal and a rotary component. The face seal includes a first annular seal surface, and the rotary component includes a second annular seal surface positioned adjacent the first annular seal surface and defining a seal interface therebetween. The face seal is configured to discharge a flow of air towards the seal interface. The seal assembly also includes a first seal member extending between the first and second annular seal surfaces such that the flow of air induces a back pressure across the seal interface, and a second seal member positioned radially inward from the first seal member and extending between the first and second annular seal surfaces. A length of the second seal member is selected to increase the back pressure induced across the seal interface.

A turbine wheel (100), a gas turbine (200) and a method for cooling a turbine wheel (100) are presented. A turbine wheel (100) has a rotation axis (5), a first side (110) and a second side (120), said sides (110, 120) being opposite sides in a direction of the rotation axis (5). The turbine wheel (100) comprises a first opening (10) at the first side (110) and a number of second openings (20) at the second side (120), wherein the first opening (10) is in a fluid communication, via at least one fluid channel (15), with the at least one of the number of second openings (20), the at least one fluid channel (15) comprising at least one curved portion, and a shaft position (12) at the first side (110) for attaching a shaft (160), wherein the first opening (10) is arranged to a different position with respect to the shaft position (12).

An air intake module for an aircraft turbomachine extending along a longitudinal axis oriented from upstream to downstream and including an inner wall and an outer wall defining together a vein for circulating an air flow, a plurality of guide vanes of the air flow extending radially in the vein and each comprising an upstream leading edge, at least one stator member fixedly mounted on the inner wall and on the outer wall and extending radially in the vein upstream of the guide vanes, the stator member comprising an infrared wave emitting device configured to emit an infrared beam on the leading edge of at least one of the guide vanes for de-icing it.