A jet engine has an inlet 11 configured to introduce air, and a combustor 12 having a fuel injection port 30a that injects a fuel, and configured to combust the fuel injected from the fuel injection port 30a by using the air. The combustor 12 has a separation section 14 defining the air passage FA through which the air flows, between a rear end 15 of the inlet and the fuel injection port 30a. A plurality of turbulent flow generating sections (20;25) are arranged in the separation section 14 to makes the air flow turbulent. Each of the plurality of turbulent flow generating sections (20;25) contains a member (21;22;25B) which can restrain the turbulence of the air flow by moving or disappearing. It can be prevented that a high-pressure region reaches the inlet so that the thrust of the jet engine is reduced.

A jet engine has an inlet, and a combustor with a fuel injection opening. The combustor has wall sections that define an air flow path. The wall section has a ramp section with a ramped surface and a first wall surface extending from an end of the ramped surface on the front side. The fuel injection opening is arranged in the ramp section. A ramped angle of the ramped surface to the first wall surface is equal to or more than 13.5 degrees. The height of the ramp section protruding from the first wall surface is equal to or less than of the height of the entrance of combustor. In this way, the ramp angle of the ramped surface of the ramp section is made large to promote mixing of fuel and air, and flame stabilization behind the ramp section is made possible.

A gas turbine engine combustor includes a dome plate, an inner liner, and an outer liner. The inner liner and the outer liner are separately coupled to a rear side of the dome plate and both extend rearward therefrom. The outer and inner liners define an annular combustion chamber therebetween that extends rearward from the dome plate. At least one of the outer liner or the dome plate includes an igniter cavity formed therein that is separated from the combustion chamber by a partition wall. The partition wall defines one or more transfer holes therethrough that fluidly connect the igniter cavity to the combustion chamber.

A combustion chamber assembly of a gas turbine, for combusting a fuel in the presence of combustion air, includes: a combustion chamber, in which combustion of fuel occurs; a precombustion chamber upstream of the combustion chamber; an atomization device that feeds a liquid fuel to the precombustion chamber; and a swirl body that feeds combustion air and gaseous fuel to the precombustion chamber. The combustion chamber assembly is configured as a dual-fuel combustion chamber assembly, which, in a gas fuel operating mode, feeds a mixture of a gaseous fuel and combustion air to the combustion chamber via the swirl body, and which, in a liquid fuel operating mode, feeds liquid fuel to the combustion chamber via the atomization device and combustion air to the combustion chamber via the swirl body. The atomization device includes an atomization lance with a central atomization nozzle, and plural decentralized atomization nozzles.

Various embodiments include a trapped vortex combustor and a method for operating trapped vortex combustor. In one embodiment, the trapped vortex combustor comprises a trapped vortex combustion zone and at least one secondary combustion zone disposed downstream of the trapped vortex combustion zone. The trapped vortex combustion zone is operable to receive and combust a first fuel and a first air and produce a first combustion product flowing toroidally therein. The at least one secondary combustion zone is operable to receive and combust the first combustion product and at least one second injection consisting of fuel and/or air and produce at least one second combustion product therein. The combustor may reduce the residence time of the highest temperature combustion products and achieve the lower NOx emission.

A hybrid combustion system, and method of operation, for a propulsion system is provided. The hybrid combustion system defines a radial direction, a circumferential direction, and a longitudinal centerline in common with the propulsion system extended along a longitudinal direction. The hybrid combustion system includes a rotating detonation combustion (RDC) system comprising an annular outer wall and an annular inner wall each generally concentric to the longitudinal centerline and together defining a RDC chamber and a RDC inlet, the RDC system further comprising a nozzle located at the RDC inlet defined by a nozzle wall. The nozzle defines a lengthwise direction extended between a nozzle inlet and a nozzle outlet along the lengthwise direction, and the nozzle inlet is configured to receive a flow of oxidizer. The nozzle further defines a throat between the nozzle inlet and the nozzle outlet, and wherein the nozzle defines a converging-diverging nozzle. The hybrid combustion system further includes an inner liner extended generally along the longitudinal direction; an outer liner extended generally along the longitudinal direction and disposed outward of the inner liner along the radial direction; a bulkhead wall disposed at the upstream end of the inner and outer liners, in which the bulkhead wall extends generally in the radial direction and couples the inner liner and the outer liner, and wherein the inner liner, the outer liner, and the bulkhead wall together define a primary combustion chamber, and further wherein the RDC system and bulkhead wall together define a RDC outlet through the bulkhead wall and adjacent to the primary combustion chamber; and a fuel manifold assembly extended at least partially through the bulkhead wall, in which the fuel manifold assembly defines a fuel manifold assembly exit disposed adjacent to the primary combustion chamber.

Combustor assemblies and methods for assembling combustor assemblies are provided. For example, a combustor assembly comprises an annular inner liner and an annular outer linear, each extending generally along an axial direction. The outer liner includes an outer flange extending forward from its upstream end. The combustor assembly also comprises a combustor dome extending between an inner liner upstream end and the outer liner upstream end and including an inner flange extending forward from a radially outermost end of the combustor dome. The inner liner, outer liner, and combustor dome define a combustion chamber therebetween, and the combustor dome and a portion of the outer liner together define an annular cavity of the combustion chamber. The inner and outer flanges define an airflow opening therebetween, and a chute member is positioned within the airflow opening to define an air chute for providing a flow of air to the annular cavity.

A trapped vortex combustor for use in a gas turbine engine includes an outer vortex chamber wall and a dome attached to, or formed integrally with, the outer vortex chamber wall. The dome, the outer vortex chamber wall, or both define at least in part an outer trapped vortex chamber and a channel. The channel extends along the circumferential direction at a forward end of the outer vortex chamber wall, the channel configured to receive an airflow through or around the outer vortex chamber wall, the dome, or both and provide such airflow as a continuous annular airflow to the inner surface of the outer vortex chamber wall. The dome further defines a fuel nozzle opening, with all openings in the dome outward of the fuel nozzle opening along the radial direction, excepting any effusion cooling holes having a diameter less than about 0.035 inches, being in airflow communication with the channel.

An object is to realize combustion flame which can further reduce the amount of NOx generation. A combustor (14) includes a pilot nozzle (40); a plurality of main nozzles (44) arranged apart from the pilot nozzle (40) in the circumferential direction on the outer peripheral side of the pilot nozzle (40) and configured to perform premix combustion; a combustor basket (34) surrounding the pilot nozzle (40) and each main nozzle (44); an outlet outer ring (50) provided at a tip end of the combustor basket (34); and a combustion liner (36) fitted, at an inner surface thereof, onto the outer periphery of the combustor basket (34) and surrounding the outlet outer ring (50). The outlet outer ring (50) is formed parallel to an inner wall surface (66) of the combustion liner (36).

An engine includes an inlet tube introducing air to a combustion process and a first plurality of fuel injectors disposed in the inlet tube and used for scram-jet engine operation. The engine includes a second plurality of fuel injectors used for ram-jet engine operation. The second plurality of fuel injectors is upstream from the first plurality of fuel injectors and is disposed in the inlet tube. The engine includes a combustor swirl zone downstream of and adjacent to the first plurality of fuel injectors.