An assembly is provided for a turbine engine. This assembly includes a static structure and an impeller rotor housed within the static structure. The impeller rotor includes a vane structure and a shroud. The vane structure includes a first sidewall, a second sidewall and a plurality of vanes arranged circumferentially about a rotational axis. The vanes include a first vane. The first vane includes a first portion, a second portion and a third portion. The first portion is axially between the first sidewall and the second sidewall. The second portion is radially between the first sidewall and the shroud. The third portion is radially between the second sidewall and the shroud. The shroud circumscribes the vane structure. A gap is formed by and extends between the shroud and the static structure. A dimension of the gap changes as the gap extends along the shroud.

An assembly is provided for a turbine engine. This assembly includes a static structure and an impeller rotor housed within the static structure. The impeller rotor includes a vane structure and a shroud. The vane structure includes a first sidewall, a second sidewall and a plurality of vanes arranged circumferentially about a rotational axis. The vanes include a first vane. The first vane includes a first portion, a second portion and a third portion. The first portion is axially between the first sidewall and the second sidewall. The second portion is radially between the first sidewall and the shroud. The third portion is radially between the second sidewall and the shroud. The shroud circumscribes the vane structure. A gap is formed by and extends between the shroud and the static structure. A dimension of the gap changes as the gap extends along the shroud.

The present technique presents a blade 1 for a gas turbine 10. The blade 1 includes an airfoil 100 having an airfoil tip part 100a and a pressure side 102 and a suction side 104 meeting at a leading edge 106 and a trailing edge 108 and defining an internal space 100s of the airfoil 100. A squealer tip 80, 90 is arranged at the airfoil tip part 100a. The squealer tip 80, 90 comprises a suction side rail 90. The suction side rail 90 comprises a chamfer part 90x and at least one squealer tip cooling hole 99. The chamfer part 90x comprises a chamfer surface 9. An outlet 99a of the at least one squealer tip cooling hole 99 is disposed at the chamfer surface 9.

The present technique presents a blade 1 for a gas turbine 10. The blade 1 includes an airfoil 100 having an airfoil tip part 100a and a pressure side 102 and a suction side 104 meeting at a leading edge 106 and a trailing edge 108 and defining an internal space 100s of the airfoil 100. A squealer tip 80, 90 is arranged at the airfoil tip part 100a. The squealer tip 80, 90 comprises a suction side rail 90. The suction side rail 90 comprises a chamfer part 90x and at least one squealer tip cooling hole 99. The chamfer part 90x comprises a chamfer surface 9. An outlet 99a of the at least one squealer tip cooling hole 99 is disposed at the chamfer surface 9.

A gas turbine engine is disclosed having a turbine, one or more hydrocarbon gas combustors, and a compressor. The compressor has a rotor assembly with one or more rotor blade rows extending radially outward from an inner wheel disk. The compressor also has a stator assembly with one or more stator vane rows extending radially inward from an inner casing and positioned between adjacent rotor blade rows. The inner casing extends circumferentially around the rotor assembly and is constructed from at least one low-alpha metal alloy.

A nozzle assembly for a gas turbine engine includes a nozzle having a first material defining a first coefficient of thermal expansion, the nozzle having an airfoil defining a fluid passage therein, an inlet wall defining a fluid inlet that is fluidly connected to the fluid passage, and a passive valve assembly comprising an annular band, the annular band comprising a second material having a second coefficient of thermal expansion less than the first coefficient of thermal expansion such that the passive valve assembly is at least partially moveable relative to the fluid inlet.

A nozzle assembly for a gas turbine engine includes a nozzle having a first material defining a first coefficient of thermal expansion, the nozzle having an airfoil defining a fluid passage therein, an inlet wall defining a fluid inlet that is fluidly connected to the fluid passage, and a passive valve assembly comprising an annular band, the annular band comprising a second material having a second coefficient of thermal expansion less than the first coefficient of thermal expansion such that the passive valve assembly is at least partially moveable relative to the fluid inlet.

According to an embodiment, a turbine stage sealing mechanism reduces leak flow, which bypasses a working fluid flow path, between a rotating part and a stationary part in each turbine stage of a gas turbine. A radially inner side end portion of the stationary part is formed to be on a circle centered at a center axis of the rotating part under a predetermined operating condition of the gas turbine.

A device for cooling a turbine shroud comprising at least one annular flange configured to be attached to an annular radial collar of a shroud support structure being arranged upstream, with respect to a circulation direction of an air flow, of the turbine shroud, and comprising at least one cooling air circulation channel, a diffuser configured to be attached to said annular radial collar downstream of the annular flange and comprising at least one intake channel in fluid communication with the circulation channel of the annular flange, and comprising an injection cavity comprising a plurality of injection holes and being configured to inject on a radially external face of the shroud, via the injection holes, the cooling air originating in the intake channel, and a particle filter arranged on an inlet section of the circulation channel of the annular flange, the particle filter comprising a plurality of openings.

A gas turbine engine including a first turbine rotor assembly having a plurality of first turbine rotor blades extended within a gas flowpath, and a second turbine rotor assembly positioned aft along the gas flowpath of the first turbine rotor assembly. The second turbine rotor assembly is rotatably separate from the first turbine rotor assembly. A casing surrounds the first turbine rotor assembly. The casing has a unitary, integral outer casing wall extended forward of the first turbine rotor assembly and aft of the first turbine rotor assembly. The casing includes a plurality of vanes extended from the outer casing wall and through the gas flowpath aft of the first turbine rotor assembly and forward of the second turbine rotor assembly. The casing includes a plurality of walls forming thermal control rings extended outward along the radial direction from the outer casing wall. The outer casing wall and the thermal control rings is a unitary, integral structure.