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 combustor includes a combustion tube having a cylindrical shape with a combustion space where fuel is combusted and including an inlet through which the fuel is introduced, an outlet through which a gas generated when the fuel is combusted is discharged, and a protrusion protruding inward from a wall surface between the inlet and the outlet; an injection unit configured to inject fuel into the combustion tube through the inlet of the combustion tube; and an additional injection unit located on the protrusion of the combustion tube and configured to inject fuel into the combustion tube.

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.

An afterburner for use with a gas turbine engine includes a plurality of vanes distributed downstream of a turbine of the gas turbine engine. The vanes can include one or more exit apertures through which hot combustion flow from a pilot can be injected. The exit apertures can be protrusions or slots in some forms. In some embodiments, cooling passages are arranged around the exit apertures. An upstream vane portion can be positioned to inject fuel to be combusted via interaction with hot flow that is discharged through the exit apertures.

There is disclosed an igniter for a gas turbine engine including: a base, a glow plug heater rod, the glow plug heater rod extending from the base along an axis and terminating in a rod end, a sleeve extending circumferentially around the glow plug heater rod along at least a portion of a length of the glow plug heater rod, the sleeve having an inner surface spaced from the glow plug heater rod to provide a gap between the inner surface and the glow plug heater rod, the gap extending radially relative to said axis.

There is disclosed an igniter for a gas turbine engine including: a base, a glow plug heater rod extending from the base along an axis and terminating in a rod end, a sleeve extending circumferentially around the glow plug heater rod along at least a portion of a length of the glow plug heater rod, the igniter having a ventilation path extending from an inlet to an outlet, the inlet fluidly connectable to a spacing located between a casing and a combustor liner of the gas turbine engine, the outlet extending circumferentially around the glow plug heater rod and oriented axially relative to the axis, the outlet fluidly connectable to a combustion chamber of the gas turbine engine delimited by the combustor liner.

There is disclosed an igniter for a gas turbine engine including: a base; a glow plug heater rod extending from the base along an axis and terminating in a rod end; and a fuel receiver adjacent the heater rod, the fuel receiver including a portion located closest to the heater rod, the rod end protruding axially relative to the axis from the fuel receiver portion located closest to the heater rod.

A combustion system for a gas turbine engine includes a combustor dome. The combustor dome includes a fuel manifold with an inlet and nozzle components of a plurality of nozzles circumferentially spaced around the combustor dome. The fuel manifold and nozzle components are integrated with the combustor dome for fluid communication from the inlet to the nozzle components.

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.

System comprising a gas turbine engine fuel injection apparatus (11) arranged to deliver to a combustion chamber (9) of a gas turbine engine a first injection fluid comprising a first fuel and a second injection fluid comprising a second fuel. The apparatus is arranged to deliver the first and second injection fluids in a manner such that the first injection fluid is delivered in a first stream and the second injection fluid is delivered in a second stream. Further, such that there is a delivery zone (47) corresponding to a first location at which both the first and second streams have been delivered in which the first stream is substantially radially surrounded by the second stream.