Abstract: The present disclosure is directed to a propulsion system including a rotating detonation combustion (RDC) system defining a plurality of fuel-oxidizer mixing nozzles each defined by a converging-diverging nozzle wall defining a nozzle flowpath. The nozzle wall defines a throat and a lengthwise direction extended between a nozzle inlet and nozzle outlet along the lengthwise direction. The longitudinal centerline of the propulsion system and the radial direction together define a reference plane, and the lengthwise direction of the nozzle intersects the reference plane and defines a nozzle angle greater than zero degrees and approximately 80 degrees or less relative to the reference plane.

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: The present disclosure is directed to a rotating detonation combustion system for a propulsion system, the rotating detonation combustion system defining a radial direction, a circumferential direction, and a longitudinal centerline in common with the propulsion system extended along a longitudinal direction. The rotating detonation combustion system includes an annular outer wall and an annular inner wall each generally concentric to the longitudinal centerline and together defining at least in part a combustion chamber and a combustion chamber inlet. The outer wall and the inner wall together define an annular nozzle concentric to the longitudinal centerline at the combustion chamber inlet. The nozzle defines a lengthwise direction and extending between a nozzle inlet and a nozzle outlet along the lengthwise direction, the nozzle inlet configured to receive a flow of oxidizer. The nozzle further defines a throat between the nozzle inlet and nozzle outlet.

Abstract: The present disclosure is directed to a rotating detonation combustion system for a propulsion system, the rotating detonation combustion system defining a radial direction, a circumferential direction, and a longitudinal centerline in common with the propulsion system extended along a longitudinal direction.

Abstract: The present disclosure is directed to a method of operating a propulsion system at an approximately constant detonation cell quantity in the combustion chamber of a detonation combustion system. The propulsion system defines an inlet section upstream of the rotating detonation combustion system and an exhaust section downstream of the rotating detonation combustion system.

Abstract: A laser ignition system for an internal combustion engine, and more specifically a gas turbine engine, is provided. The system including a laser light source configured to generate a laser beam, an ignition port configured to provide optimized optical access of the laser beam to a combustion chamber and an optical beam guidance component disposed between the laser light source and the ignition port. The optical beam guidance component is configured to include optimized optic components to transmit the laser beam to irradiate on a fuel mixture supplied into the combustion chamber to generate a combustor flame in a flame region. A method for igniting a fuel mixture in an internal combustion engine is also presented.

Abstract: A laser ignition system for an internal combustion engine, and more specifically a gas turbine engine, is provided. The system including a laser light source configured to generate a laser beam, an ignition port configured to provide optimized optical access of the laser beam to a combustion chamber and an optical beam guidance component disposed between the laser light source and the ignition port. The optical beam guidance component is configured to include optimized optic components to transmit the laser beam to irradiate on a fuel mixture supplied into the combustion chamber to generate a combustor flame in a flame region. A method for igniting a fuel mixture in an internal combustion engine is also presented.

Abstract: A combustor liner includes a liner having an upstream end and a downstream end having a longitudinal axis extending therethrough, and a plurality of cooling holes formed in the liner, the cooling holes are arranged along the longitudinal axis into a plurality of circumferentially extending rows that are variably spaced apart along the longitudinal axis.

Abstract: A combustor liner includes a liner having an upstream end and a downstream end having a longitudinal axis extending therethrough, and a plurality of cooling holes formed in the liner, the cooling holes are arranged along the longitudinal axis into a plurality of circumferentially extending rows that are variably spaced apart along the longitudinal axis.

Abstract: An igniter tube assembly includes an axis of symmetry extending therethrough, an igniter tube that includes a first opening extending coaxially therethrough having a diameter sized to receive a portion of the igniter therethrough such that the igniter tube circumscribes the igniter and such that a gap is defined between the igniter tube and the igniter, a ferrule coupled to the igniter tube, and a plurality of cooling air openings extending through at least one of the igniter tube and the ferrule to facilitate channeling cooling air into the gap.

Abstract: A fabricated cowl for a double annular combustor of a gas turbine engine having a longitudinal axis extending therethrough, including an outer annular portion, an inner annular portion, and a middle annular portion. A plurality of circumferentially spaced radial slots are formed in at least one of the annular portions. The outer annular portion includes a plurality of circumferentially spaced windows of desired size formed therein, wherein a plurality of outer radial members connecting the outer annular portion and the middle annular portion are provided between adjacent windows therein. Similarly, the inner annular portion includes a plurality of circumferentially spaced windows of desired size formed therein, wherein a plurality of inner radial members connecting the inner annular portion and the middle annular portion are provided between adjacent windows therein.

Abstract: A method facilitates washing a gas turbine engine combustor. The method comprises coupling a nozzle assembly against the combustor, wherein the nozzle assembly includes an inlet end, a discharge end, a hollow nozzle body extending therebetween, and a centerbody positioned within the nozzle body, coupling the nozzle assembly to a fluid source, and discharging an annulus of fluid from the nozzle assembly into the combustor to facilitate removing particulate matter from the combustor.

Abstract: A method facilitates washing a gas turbine engine combustor. The method comprises coupling a nozzle assembly against the combustor, wherein the nozzle assembly includes an inlet end, a discharge end, a hollow nozzle body extending therebetween, and a centerbody positioned within the nozzle body, coupling the nozzle assembly to a fluid source, and discharging an annulus of fluid from the nozzle assembly into the combustor to facilitate removing particulate matter from the combustor.

Abstract: A fabricated cowl for a double annular combustor of a gas turbine engine having a longitudinal axis extending therethrough, including an outer annular portion, an inner annular portion, and a middle annular portion. A plurality of circumferentially spaced radial slots are formed in at least one of the annular portions. The outer annular portion includes a plurality of circumferentially spaced windows of desired size formed therein, wherein a plurality of outer radial members connecting the outer annular portion and the middle annular portion are provided between adjacent windows therein. Similarly, the inner annular portion includes a plurality of circumferentially spaced windows of desired size formed therein, wherein a plurality of inner radial members connecting the inner annular portion and the middle annular portion are provided between adjacent windows therein.

Abstract: A combustor for a gas turbine engine includes a plurality of igniter tubes that facilitate reducing temperature gradients within the combustor in a cost effective and reliable manner. The combustor includes an annular outer liner that includes a plurality of openings sized to receive igniter tubes. Each igniter tube maintains an alignment of each igniter received therein, and includes an air impingement device that extends radially outward from the igniter tube. During operation, airflow contacting the air impingement device is channeled radially inward for impingement cooling of the igniter tubes and the combustor outer liner.

Abstract: A one-piece deflector-flare cone assembly for a gas turbine engine combustor that facilitates extending a useful life of the combustor in a cost-effective and reliable manner is described. The one-piece assembly includes a deflector portion and a flare cone portion. The deflector portion includes an integral opening that extends through the deflector portion for receiving cooling fluid therein. The cooling opening extends circumferentially within the deflector portion. Cooling fluid discharged from the cooling opening is used for film cooling a portion of the deflector portion to facilitate reducing an operating temperature and extending a useful life of the combustor.

Abstract: A one-piece deflector-flare cone assembly for a gas turbine engine combustor that facilitates extending a useful life of the combustor in a cost-effective and reliable manner is described. The one-piece assembly includes a deflector portion and a flare cone portion. The deflector portion includes an integral opening that extends through the deflector portion for receiving cooling fluid therein. The cooling opening extends circumferentially within the deflector portion. Cooling fluid discharged from the cooling opening is used for impingement cooling a portion of the flare cone portion to facilitate reducing an operating temperature and extending a useful life of the combustor.

Abstract: A combustor for a gas turbine engine includes a plurality of igniter tubes that facilitate reducing temperature gradients within the combustor in a cost effective and reliable manner. The combustor includes an annular outer liner that includes a plurality of openings sized to receive igniter tubes. Each igniter tube maintains an alignment of each igniter received therein, and includes an air impingement device that extends radially outward from the igniter tube. During operation, airflow contacting the air impingement device is channeled radially inward for impingement cooling of the igniter tubes and the combustor outer liner.