Abstract: The present disclosure is directed to a rotating detonation combustion system for a propulsion system including a plurality of combustors in adjacent arrangement along the circumferential direction. Each combustor defines a combustor centerline extended through each combustor, and each combustor comprises an outer wall defining a combustion chamber and a combustion inlet. Each combustion chamber is defined by an annular gap and a combustion chamber length together defining a volume of each combustion chamber. Each combustor defines a plurality of nozzle assemblies each disposed at the combustion inlet in adjacent arrangement around each combustor centerline. Each nozzle assembly defines a nozzle wall extended along a lengthwise direction, a nozzle inlet, a nozzle outlet, and a throat therebetween, and each nozzle assembly defines a converging-diverging nozzle. A first array of combustors defines a first volume and a second array of combustors defines a second volume different from the first volume.

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 rotating detonation combustion system includes an outer wall and an inner wall together defining at least in part a combustion chamber and a combustion chamber inlet. A nozzle of the rotating detonation combustion system is located at the combustion chamber inlet, the nozzle defining a lengthwise direction and extending between a nozzle inlet and a nozzle outlet along the lengthwise direction. The nozzle further defines a throat between the nozzle inlet and nozzle outlet. A fuel injection port is also provided, the fuel injection port defining a fuel outlet located between the nozzle inlet and the nozzle outlet for providing fuel to a flow of oxidizer received through the nozzle inlet.

Abstract: The present disclosure is directed to a fuel injector assembly of a gas turbine engine, including a centerbody, an outer sleeve surrounding the centerbody, an inner sleeve disposed between the centerbody and at least a section of the outer sleeve, and at least one axially oriented vane defined between the centerbody and a section of the inner sleeve. A section of the outer sleeve includes at least one radially oriented vane. A portion of the inner sleeve includes a contoured shroud and defines at least one fuel injection port disposed upstream of the contoured shroud. The axially oriented vane is disposed upstream of the fuel injection port.

Abstract: A low emission fuel injection system and combustion chamber for use in gas turbine engines comprises one fuel injection body having a dual circuit to supply both pilot and main fuel systems. Both pilot fuel circuit and a main fuel circuit inject fuel at essentially the same axial and radial location. The recessed pilot fuel injection site is along the combustor centerline into a swirling air passage produced by axial air swirlers. The main fuel is injected radially through a plurality of injection sites, at a compound angle, into the inner diameter of a swirling air passage produced by radial air swirlers. The fuel/air residence time prior to entering the combustion chamber is relatively short, minimizing the likelihood of auto ignition. During pilot circuit only operation, the flame is stabilized by a swirler produced recirculation zone, producing high temperatures to completely burn the fuel producing low CO and UHC emissions.

Abstract: An annular combustor having fuel preparation chambers mounted in the dome of the combustor. The fuel preparation chamber comprises an annular wall extending axially from an inlet to an exit that defines a mixing chamber. Mounted to the inlet are an air swirler and a fuel atomizer. The air swirler provides swirled air to the mixing chamber while the atomizer provides a fuel spray. On the downstream side of the exit, the fuel preparation chamber has an inwardly extending conical wall that compresses the swirling mixture of fuel and air exiting the mixing chamber.

Abstract: A swirling apparatus for directing air into an annular combustion chamber is disclosed comprising a substantially elliptical vane array disposed around a cylindrical fuel injector. The vanes comprising the vane array extend substantially radially from the fuel injector and define first and second air passages therebetween. The first air passages permit an air mass flow through the vane array having a tangential component greater than that of the air mass flow permitted by the second air passages. Each vane has a helical pitch of 60 degrees and comprises a radially outermost edge, a leading edge and a trailing edge. The vane array has major and minor axes of predetermined length with the length of the major axis being greater than the length of the minor axis by a factor of at least 1.3. When used in conjunction with a conventional annular combustion chamber, the minor axis of each of the elliptical vane arrays is aligned radially with respect to the longitudinal axis of the combustion chamber.