An example system and corresponding method includes a jet engine combustor and a plurality of resonators. The combustor includes (i) a combustion zone, (ii) one or more fuel inlets for introducing fuel into the combustion zone for combustion, and (iii) one or more fins protruding into the combustion zone and configured to guide combustion of the fuel along a flame path. The resonators can each have a respective resonant wavelength and can each provide a respective plasma corona in the combustion zone when excited with a respective signal having a wavelength proximate to an odd-integer multiple of one-quarter () of the respective resonant wavelength. A radio-frequency power source can excite the resonators with the respective signals so as to provide the respective plasma coronas in the combustion zone and cause combustion of the fuel along the flame path.

A method and a device provide a uniform recursive sequential combustion of fuel and oxidizing agents within a thermal system having a continuous flow. Compressed fresh air is directed through the combustion chamber along a primary flow direction. A proportion of the fresh air is supplied to a burner by way of a burner entry and in the burner is combusted with fuel and exits the burner as exhaust gas. The burner is disposed at an angle in relation to the primary flow direction such that part of the exhaust gas exiting the burner exit is imparted a tangential flow in relation to the primary flow direction and circulates in the combustion chamber and enters the burner entry of a downstream burner so as to be mixed with the fresh air flowing into the downstream burner such that a recursive sequential combustion is achieved.

A turbomachine includes a plurality of transition ducts disposed in a generally annular array. Each of the plurality of transition ducts includes an inlet, an outlet, and a passage defining an interior and extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of each of the plurality of transition ducts is offset from the inlet along the longitudinal axis and the tangential axis. The turbomachine includes a support ring assembly downstream of the plurality of transition ducts along a hot gas path, and a plurality of mechanical fasteners connecting at least one transition duct of the plurality of transition ducts to the support ring assembly. The turbomachine includes a late injection assembly providing fluid communication for an injection fluid to flow into the interior downstream of the inlet of at least one transition duct of the plurality of transition ducts.

A rotating detonation combustor includes a combustion chamber configured for a rotating detonation process to produce a flow of combustion gas and an air plenum configured to contain a volume of air. The rotating detonation combustor also includes a flow passage coupled in flow communication between the combustion chamber and the air plenum and configured to channel an airflow from the air plenum. The rotating detonation combustor also includes a fuel inlet coupled in flow communication with the flow passage and configured to channel a fuel flow into the flow passage. The flow passage includes a plurality of fuel mixing mechanisms configured to mix the airflow and the fuel flow within the combustion chamber.

A ducting arrangement (12) for a combustion turbine engine is provided. The arrangement includes a ceramic liner (22) defining a hot gas path throughout a length of the ducting arrangement. A cooling sleeve (24) is disposed circumferentially outwardly onto the ceramic liner along the length. A metallic support frame (26) is disposed circumferentially outwardly onto the cooling sleeve along the length. The cooling sleeve may be structured with structural features along the length for biasing against the ceramic liner and the metallic support frame to resiliently accept mechanical and thermal growth induced loading that develops between the ceramic liner and the metallic support frame during operating conditions of the combustion turbine engine.

A combustor arrangement with a front panel, a combustor liner, and a carrier structure element is provided for carrying the front panel and the combustor liner, wherein the combustor arrangement further includes a fastening system for connecting the front panel, the combustor liner, and the carrier structure element to one another. The fastening system includes at least one elastic connection element, the latter being fixedly connected to the carrier structure element and extending therefrom to the combustor liner and to the front panel. The elastic connection element is further fixedly connected to the combustor liner and/or the front panel such as to clamp the front panel, the combustor liner, and the carrier structure element to one another in a substantially fluid tight manner.

An arrangement (10) for delivering gases from combustors (15) to a first row of blades. The arrangement (10) includes at least an upstream flow path (60) including an aft first side wall (64) and a downstream flow path (62) including a forward second side wall (66). A convergence junction trailing edge (40) is defined at a downstream terminal edge (41) of the first side wall (64), and the second side wall (66) converges toward the first side wall (64) in the direction of the convergence junction trailing edge (40). An impingement sheet structure (78) is located between and provides impingement cooling air to the first and second side walls (64, 66). Openings (88) provide a cooling air passage between the first and second side walls (64, 66) and provide a flow of post impingement air into the gas path at the convergence junction trailing edge (40).

A rotating detonation combustor includes a combustion chamber configured for a rotating detonation process to produce a flow of combustion gas and an air plenum configured to contain a volume of air. The rotating detonation combustor also includes a flow passage coupled in flow communication between the combustion chamber and the air plenum and configured to channel an airflow from the air plenum. The rotating detonation combustor also includes a fuel inlet coupled in flow communication with the flow passage and configured to channel a fuel flow into the flow passage. The flow passage includes a plurality of fuel mixing mechanisms configured to mix the airflow and the fuel flow within the combustion chamber.

An arrangement (10) for delivering gases from combustors (15) to a first row of blades. The arrangement (10) includes at least an upstream flow path (60) including an aft first side wall (64) and a downstream flow path (62) including a forward second side wall (66). A convergence junction trailing edge (40) is defined at a downstream terminal edge (41) of the first side wall (64), and the second side wall (66) converges toward the first side wall (64) in the direction of the convergence junction trailing edge (40). An impingement sheet structure (78) is located between and provides impingement cooling air to the first and second side walls (64, 66). Openings (88) provide a cooling air passage between the first and second side walls (64, 66) and provide a flow of post impingement air into the gas path at the convergence junction trailing edge (40).

The invention relates to a turbine engine combustion chamber including: an outer annular housing; a flame tube (20) connected to the outer housing, said flame tube (20) including an inner annular wall (20b) and an outer annular wall (20a) that define a first radial inlet portion of the flame tube and a second axial outlet portion of the flame tube, the flame tube also including a chamber base (30) located at the inlet of the flame tube (20); and a fuel injection system (40) configured to inject fuel into the flame tube via the inlet of the flame tube. The injection system includes an injector axis (AA), parallel to the first portion, and an air manifold (40d) configured to move air towards twists in the injection system (40). The twists are arranged around an implantation axis parallel to the injector axis. The air manifold includes a circular portion around the injector axis. The circular portion, from which extends an opening, forms an air inlet of the manifold. The opening is configured to place the entering air flow in rotation about the implantation axis so that said air flow is supplied to the twists.