Abstract: A combustion system may include a detonation combustor comprising one or more detonation chamber walls defining a detonation chamber, a deflagration combustor comprising one or more deflagration chamber walls defining a deflagration chamber, and one or more conjugate chamber walls defining a conjugate chamber, with the conjugate chamber in fluid communication with the detonation chamber and the deflagration chamber. The detonation chamber includes a detonation region and a nozzle region, with the nozzle region providing fluid communication between the detonation region and the conjugate chamber.

Abstract: A detonation combustion system includes a detonation combustor. The detonation combustor includes a detonation manifold and one or more detonation chamber walls defining a detonation chamber. The detonation manifold includes a plurality of detonation fluid pathways defined by a monolithic structure of the detonation manifold, and a plurality of detonation orifice groups respectively including a plurality of detonation orifices disposed about a surface of the detonation manifold. Respective ones of the plurality of detonation orifice groups provide fluid communication from a corresponding one of the plurality of detonation fluid pathways to the detonation chamber through the plurality of detonation orifices corresponding to the respective one of the plurality of detonation orifice groups. The plurality of detonation orifices may be symmetrically oriented about a reference element of the detonation combustor.

Abstract: A computer-implemented method for multi-mode operation of a combustion system, a combustion system, and a heat engine are provided. The method includes initializing combustion of a fuel/oxidizer mixture, determining whether conditions at the combustion system meet or exceed a first threshold operating parameter, transitioning to detonation combustion of the fuel/oxidizer mixture if conditions at the combustion system meet or exceed the first threshold operating parameter, and maintaining or increasing fuel flow through a deflagrative fuel circuit if conditions at the combustion system do not meet or exceed the first threshold operating parameter.

Abstract: A rotating detonation combustion (RDC) system is provided. The RDC includes a first outer wall and a second outer wall each extended around a centerline axis, and a detonation chamber formed radially inward of the second outer wall. A fuel passage extended between the first outer wall and the second outer wall, the fuel passage including a first inlet opening proximate to the aft end through which a flow of fuel is received into the fuel passage. The flow of fuel is provided through the fuel passage from the aft end to the forward end of the RDC system and to the detonation chamber.

Abstract: A combustion system includes an annular tube disposed between an inner wall and an outer wall, the annular tube extending from an inlet end to an outlet end; at least one fluid inlet disposed in the annular tube proximate the inlet end, the fluid inlet providing a conduit through which fluid flows into the annular tube; at least one outlet disposed in the annular tube proximate the outlet end; at least one primary fuel injector, the primary fuel injector dispersing fuel into a fluid stream entering the annular tube via the fluid inlet; and at least one secondary fuel injector, the secondary fuel injector disposed in the annular tube.

Abstract: A computer-implemented method for multi-mode operation of a combustion system, a combustion system, and a heat engine are provided. The method includes initializing combustion of a fuel/oxidizer mixture, determining whether conditions at the combustion system meet or exceed a first threshold operating parameter, transitioning to detonation combustion of the fuel/oxidizer mixture if conditions at the combustion system meet or exceed the first threshold operating parameter, and maintaining or increasing fuel flow through a deflagrative fuel circuit if conditions at the combustion system do not meet or exceed the first threshold operating parameter.

Abstract: A rotating detonation combustion (RDC) system is provided. The RDC includes a first outer wall and a second outer wall each extended around a centerline axis, and a detonation chamber formed radially inward of the second outer wall. A fuel passage extended between the first outer wall and the second outer wall, the fuel passage including a first inlet opening proximate to the aft end through which a flow of fuel is received into the fuel passage. The flow of fuel is provided through the fuel passage from the aft end to the forward end of the RDC system and to the detonation chamber.

Abstract: Systems for rotating detonation combustion are provided herein. The system includes an inner wall and an outer wall each extended around a centerline axis, wherein a detonation chamber is defined between the inner wall and the outer wall, and an iterative structure positioned at one or both of the inner wall or the outer wall. The iterative structure includes a first threshold structure corresponding to a first pressure wave attenuation and a second threshold structure corresponding to a second pressure wave attenuation. The iterative structure provides for pressure wave strengthening along a first circumferential direction in the detonation chamber or pressure wave weakening along a second circumferential direction opposite of the first circumferential direction. The first circumferential direction corresponds to a desired direction of pressure wave propagation in the detonation chamber.

Abstract: A combustor for a turbine engine includes an inner combustion liner and an outer combustion liner together defining at least in part an interior. The interior includes a combustion chamber and a main portion. The combustor also includes an inlet combustion liner at least partially defining the combustion chamber of the interior and including an inlet assembly. The inlet assembly includes at least two cavity air tubes arranged along the axial direction, each cavity air tube extending between an inlet and an outlet, the outlet of each cavity air tube in airflow communication with the combustion chamber for providing the combustion chamber with a flow of air.

Abstract: A combustor assembly for use in a gas turbine engine is provided. The combustor assembly includes a dome assembly, a fuel nozzle coupled to the dome assembly, and a cowl assembly. The cowl assembly includes a radially inner cowl coupled to the dome assembly, and a radially outer cowl including a first cowl member and a second cowl member each positioned about the fuel nozzle. The first and second cowl members substantially align to define a substantially continuous flow path configured to direct a flow of air about the dome assembly.

Abstract: A combustor for use in a turbine engine that includes an inner combustion liner and an outer combustion liner. An interior is defined between the inner combustion liner and the outer combustion liner, and the interior includes a cavity portion and a main portion extending radially inward from the cavity portion. The cavity portion includes a flow inlet and the main portion includes a flow outlet. A plurality of film cooling holes are formed in at least one of the inner combustion liner and the outer combustion liner. The plurality of film cooling holes are configured such that cooling airflow discharged therefrom flows helically relative to a centerline of the turbine engine and towards the flow outlet.

Abstract: A combustion system includes an annular tube disposed between an inner wall and an outer wall, the annular tube extending from an inlet end to an outlet end; at least one fluid inlet disposed in the annular tube proximate the inlet end, the fluid inlet providing a conduit through which fluid flows into the annular tube; at least one outlet disposed in the annular tube proximate the outlet end; at least one primary fuel injector, the primary fuel injector dispersing fuel into a fluid stream entering the annular tube via the fluid inlet; and at least one secondary fuel injector, the secondary fuel injector disposed in the annular tube.

Abstract: A fuel injector to reduce NOx emissions in a combustor system. The fuel injector including a housing, at least one oxidizer flow path, extending axially through the fuel injector housing and defining therein one or more oxidizer flow paths for an oxidizer stream and a fuel manifold, extending axially through the fuel injector housing and defining therein one or more fuel flow path. The fuel manifold includes a forward portion and an aft portion including an aft face. A plurality of fuel injector outlets are defined in the aft portion, wherein the plurality of fuel injector outlets are configured to inject a fuel flow along a mid-plane of the fuel injector and away from a downstream wall. The fuel flow exiting the fuel manifold undergoes circumferential and radial mixing upon interaction with the oxidizer stream. Additionally disclosed is a combustor system including the fuel injector.

Abstract: An air-shielded fuel injection assembly for use in a combustion chamber of a turbine assembly. The air-shielded fuel injection assembly generally includes a fuel manifold including a plurality of fuel injection ports and an air manifold including a plurality of air injection ports. Each of the plurality of fuel injection ports is configured to introduce a fuel column into an annular cavity of a mixer assembly. Each of the plurality of air injection ports is configured to introduce an air curtain about an associated fuel injection column to minimize recirculation upstream of the fuel injection column and increase penetration of the fuel injection column into the cavity. Also disclosed are a mixer assembly and a turbine assembly including the air-shielded fuel injection assembly.

Abstract: An air-shielded fuel injection assembly for use in a combustion chamber of a turbine assembly. The air-shielded fuel injection assembly generally includes a fuel manifold including a plurality of fuel injection ports and an air manifold including a plurality of air injection ports. Each of the plurality of fuel injection ports is configured to introduce a fuel column into an annular cavity of a mixer assembly. Each of the plurality of air injection ports is configured to introduce an air curtain about an associated fuel injection column to minimize recirculation upstream of the fuel injection column and increase penetration of the fuel injection column into the cavity. Also disclosed are a mixer assembly and a turbine assembly including the air-shielded fuel injection assembly.

Abstract: A gas turbine engine and combustor assembly including a combustor liner defining therein a combustion chamber for the downstream flow of a main fluid. At least two axially spaced apart annular trapped vortex cavities are located on the combustor liner and staged axially and radially spaced apart. A cavity opening is located at a radially inner end of each of the at least two annular trapped vortex cavities. A plurality of injectors are configured tangentially relative to circular radially outer wall extending between an aft wall and a forward wall of each cavity to provide for an injection of air and fuel to form an annular rotating trapped vortex of a fuel and air mixture within a respective annular trapped vortex cavity. The annular rotating trapped vortex of the fuel and air mixture at the cavity openings is substantially perpendicular to the downstream flow of the main fluid.

Abstract: A combustor for a turbine engine includes an inner combustion liner and an outer combustion liner together defining at least in part an interior. The interior includes a combustion chamber and a main portion. The combustor also includes an inlet combustion liner at least partially defining the combustion chamber of the interior and including an inlet assembly. The inlet assembly includes at least two cavity air tubes arranged along the axial direction, each cavity air tube extending between an inlet and an outlet, the outlet of each cavity air tube in airflow communication with the combustion chamber for providing the combustion chamber with a flow of air.

Abstract: A combustor for use in a turbine engine that includes an inner combustion liner and an outer combustion liner. An interior is defined between the inner combustion liner and the outer combustion liner, and the interior includes a cavity portion and a main portion extending radially inward from the cavity portion. The cavity portion includes a flow inlet and the main portion includes a flow outlet. A plurality of film cooling holes are formed in at least one of the inner combustion liner and the outer combustion liner. The plurality of film cooling holes are configured such that cooling airflow discharged therefrom flows helically relative to a centerline of the turbine engine and towards the flow outlet.

Abstract: A combustor assembly for use in a gas turbine engine is provided. The combustor assembly includes a dome assembly, a fuel nozzle coupled to the dome assembly, and a cowl assembly. The cowl assembly includes a radially inner cowl coupled to the dome assembly, and a radially outer cowl including a first cowl member and a second cowl member each positioned about the fuel nozzle. The first and second cowl members substantially align to define a substantially continuous flow path configured to direct a flow of air about the dome assembly.

Abstract: A combustor assembly including a combustor liner defining therein a combustion chamber for the downstream flow of a main fluid. At least two annular trapped vortex cavities are located on the combustor liner and staged axially spaced apart. A cavity opening is located at a radially inner end of each of the at least two annular trapped vortex cavities spaced apart from a radially outer wall and extending between an aft wall and a forward wall of each cavity. A plurality of injectors are configured tangentially relative to the circular radially outer wall to provide for an injection of air and fuel to form an annular rotating trapped vortex of a fuel and air mixture within a respective annular trapped vortex cavity. The annular rotating trapped vortex of the fuel and air mixture at the cavity openings is substantially perpendicular to the downstream flow of the main fluid. A gas turbine engine including the combustor assembly is disclosed.