A gas turbine engine including a trapped vortex combustor assembly is generally provided. The combustor assembly includes an inner liner wall extended annularly around a combustor centerline, and an outer liner wall extended annularly around the combustor centerline. The inner liner wall and the outer liner wall together define an involute wall extended at least partially as a spiral curve from a circumferential reference line around the combustor centerline. The involute wall defines an involute combustion chamber. One or more of the inner liner wall and the outer liner wall each define a first dilution opening and a shaped dilution opening.

A fuel nozzle assembly and a gas turbine having the fuel nozzle assembly includes a fuel nozzle guide disposed in a compressed air channel formed between a body and a housing of a gas turbine and includes a nozzle body disposed in the housing, a shroud mounted on an outer side of the nozzle body, and two or more flow guides arranged at predetermined distances from each other between the shroud and the outer side of the nozzle body and formed to correspond to the shape of an end of the shroud and the shape of the outer surface of the nozzle body.

A fuel nozzle assembly and a gas turbine having the fuel nozzle assembly includes a fuel nozzle guide disposed in a compressed air channel formed between a body and a housing of a gas turbine and includes a nozzle body disposed in the housing, a shroud mounted on an outer side of the nozzle body, and two or more flow guides arranged at predetermined distances from each other between the shroud and the outer side of the nozzle body and formed to correspond to the shape of an end of the shroud and the shape of the outer surface of the nozzle body.

A combined cooling heating and power micro gas turbine device includes a micro gas turbine. The micro gas turbine includes a gas compressor, a turbine and a combustion chamber assembly. The combustion chamber assembly includes a combustion chamber, an air inlet cavity, an air inlet channel and an exhaust channel. The air inlet cavity includes an interior air inlet cavity and an exterior air inlet cavity that are integrated, an air outlet end of the exterior air inlet cavity is communicated with an air inlet end of the interior air inlet cavity, an air inlet end of the exterior air inlet cavity is communicated with the air inlet channel, the air inlet channel is communicated with the gas compressor, the combustion chamber is arranged between the interior air inlet cavity and the exterior air inlet cavity, and an air outlet of the combustion chamber is communicated with the exhaust channel.

A combustor for a gas turbine engine includes an outer liner having a first end interconnected to an opposite second end by an outer liner wall composed of a plurality of outer segments. The inner liner has a first inner end interconnected to an opposite second inner end by an inner liner wall composed of a plurality of inner segments. The outer liner wall and the inner liner wall define a combustion chamber, and each of the outer wall segments extend at an angle of at least 40 degrees relative to a longitudinal axis. The outer wall segments includes a first segment, a second segment that extends at a second angle relative to the first segment, which is less than a third angle defined between the second segment and a third segment and is substantially the same as a fourth angle defined between the third segment and a fourth segment.

A combined cooling heating and power micro gas turbine device includes a micro gas turbine. The micro gas turbine includes a gas compressor, a turbine and a combustion chamber assembly. The combustion chamber assembly includes a combustion chamber, an air inlet cavity, an air inlet channel and an exhaust channel. The air inlet cavity includes an interior air inlet cavity and an exterior air inlet cavity that are integrated, an air outlet end of the exterior air inlet cavity is communicated with an air inlet end of the interior air inlet cavity, an air inlet end of the exterior air inlet cavity is communicated with the air inlet channel, the air inlet channel is communicated with the gas compressor, the combustion chamber is arranged between the interior air inlet cavity and the exterior air inlet cavity, and an air outlet of the combustion chamber is communicated with the exhaust channel.

An apparatus is provided for a turbine engine. This turbine engine apparatus includes a fuel nozzle. The fuel nozzle includes a nozzle passage and a nozzle orifice. The nozzle passage extends longitudinally along a centerline within the fuel nozzle to the nozzle orifice. The nozzle passage has a solid polygonal cross-sectional geometry at the nozzle orifice.

A module for an aircraft turbine engine including at least one annular casing of an annular combustion chamber, at least one sealing ring for a turbine wheel, and at least one annular bearing support, wherein this module is made in one piece.

A ring assembly for holding a double-skin combustor liner has an inner skin of ceramic matrix composite and an outer skin. The assembly comprises an annular end body, a first projecting portion projecting relative to the annular end body at least partially in an axial direction, the first projecting portion configured for supporting the inner skin, a second projecting portion projecting relative to the annular end body at least partially in the axial direction, the second projecting portion configured for supporting the outer skin. The first and second projecting portions are radially spaced apart to provide a radial spacing between the inner skin and the outer skin. An expansion joint is between the ring assembly and the inner skin configured to permit motion resulting from a variation in thermal expansion between the inner skin and the ring assembly.

A low-leakage seal assembly for a turbomachine's combustor-turbine interface. A turbomachine includes a combustion chamber that receives compressed air for combustion from a compressor. The combustion chamber is defined by a combustion liner terminating in a seal ring where the seal ring has an enlarged head that is thicker than the remainder of the seal ring. A transition liner directs combustion gases from the combustion chamber to the turbine and has three walls forming a cavity with an open end. The seal ring extends through the open end and the head nests in the cavity. The transition liner and the seal ring are parts of a low-leakage seal assembly that is exposed on one side to the compressed air and that is exposed on another side to the combustion gases.