A gas turbine engine radial impeller includes first and second impeller portions that are secured to one another along a neutral bending plane of the radial impeller. A vapor cooling cavity is provided between the first and second impeller portions. The neutral bending plane is arranged in the vapor cooling cavity.

A mid-turbine frame includes an inner frame case including a fluid passage. At least one hollow spoke is attached to the inner frame case. At least one hollow spoke includes an inlet passage. A transfer tube includes a first end that forms a first connection with at least one hollow spoke and a second end forms a second connection with the inner frame case.

A mid-turbine frame includes an inner frame case including a fluid passage. At least one hollow spoke is attached to the inner frame case. At least one hollow spoke includes an inlet passage. A transfer tube includes a first end that forms a first connection with at least one hollow spoke and a second end forms a second connection with the inner frame case.

An example system comprises an electric machine including a stator and a rotor, a cooling system configured to supply a cooling fluid to cool the electric machine, and a stator tube configured to contain the cooling fluid within a stator portion of the electric machine and prevent the cooling fluid from contacting the rotor.

An aircraft including a turbofan engine provided with an engine body and a fan located anterior to the engine body, further includes: an engine oil cooler that is a heat exchanger for cooling engine oil used in the engine body by using, as a heat source, a fan stream flowing from the fan into a gap between a core cowl surrounding the engine body and a nacelle surrounding the fan and the core cowl; and a pre-cooler that is a heat exchanger for cooling bleed air from the engine body by using the fan stream as a heat source, wherein the engine oil cooler and the pre-cooler are longitudinally arranged in one position in a circumferential direction of the nacelle, and the engine oil cooler is located anterior to the pre-cooler.

An aircraft including a turbofan engine provided with an engine body and a fan located anterior to the engine body, further includes: an engine oil cooler that is a heat exchanger for cooling engine oil used in the engine body by using, as a heat source, a fan stream flowing from the fan into a gap between a core cowl surrounding the engine body and a nacelle surrounding the fan and the core cowl; and a pre-cooler that is a heat exchanger for cooling bleed air from the engine body by using the fan stream as a heat source, wherein the engine oil cooler and the pre-cooler are longitudinally arranged in one position in a circumferential direction of the nacelle, and the engine oil cooler is located anterior to the pre-cooler.

A gas turbine having a combustion chamber with exhaust section through which combustion gas is exhaustable, plenum chamber and compressor are provided. The plenum chamber is coupled to the compressor wherein a first quantity of compressed fluid is injectable therein at a radially inner wall of the plenum chamber. A guide vane section with at least one airfoil is coupled to the exhaust section so combustion gas is flowable against the airfoil. The exhaust section and guide vane section are housed inside the plenum chamber. The airfoil has a first flow chamber where a second quantity of compressed fluid is flowable through the guide vane section from the compressor in the direction from the inner wall to a outer wall of the plenum chamber before being discharged. The second quantity of compressed fluid streamable through the guide vane section is larger than the first quantity of the compressed fluid.

A gas turbine having a combustion chamber with exhaust section through which combustion gas is exhaustable, plenum chamber and compressor are provided. The plenum chamber is coupled to the compressor wherein a first quantity of compressed fluid is injectable therein at a radially inner wall of the plenum chamber. A guide vane section with at least one airfoil is coupled to the exhaust section so combustion gas is flowable against the airfoil. The exhaust section and guide vane section are housed inside the plenum chamber. The airfoil has a first flow chamber where a second quantity of compressed fluid is flowable through the guide vane section from the compressor in the direction from the inner wall to a outer wall of the plenum chamber before being discharged. The second quantity of compressed fluid streamable through the guide vane section is larger than the first quantity of the compressed fluid.

A leakage management assembly for an orifice configured to limit a flow of at least one of a leakage fluid and a cooling fluid past an instrument positioned within the orifice are provided. The leakage management assembly includes a labyrinth ferrule including an annular ferrule body having a central bore extending therethrough. The labyrinth ferrule further includes an annular assembly cone formed at a first end of the labyrinth ferrule. The annular assembly cone includes an annular convergent lip and is configured to facilitate directing the instrument into the central bore. The labyrinth ferrule also includes a plurality of labyrinthine annular restrictions extending from a radially inner surface of the annular body toward the central bore. A combustor and a gas turbine engine including such a seal assembly are also provided.

A leakage management assembly for an orifice configured to limit a flow of at least one of a leakage fluid and a cooling fluid past an instrument positioned within the orifice are provided. The leakage management assembly includes a labyrinth ferrule including an annular ferrule body having a central bore extending therethrough. The labyrinth ferrule further includes an annular assembly cone formed at a first end of the labyrinth ferrule. The annular assembly cone includes an annular convergent lip and is configured to facilitate directing the instrument into the central bore. The labyrinth ferrule also includes a plurality of labyrinthine annular restrictions extending from a radially inner surface of the annular body toward the central bore. A combustor and a gas turbine engine including such a seal assembly are also provided.