Abstract: A propulsion system including: a fuel cell assembly having a fuel cell defining an outlet positioned to remove output products from the fuel cell and a fuel cell assembly operating condition; a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the combustion section configured to receive a flow of aviation fuel from an aircraft fuel supply and further configured to receive the output products from the fuel cell; and a controller comprising memory and one or more processors, the memory storing instructions that when executed by the one or more processors cause the propulsion system to perform operations including receiving data indicative of a mid-flight flameout within the combustion section; modifying the fuel cell assembly operating condition in response to receiving data indicative of the mid-flight flameout within the combustion section; and initiating a relight of the combustion section.

Abstract: A combustor for a gas turbine engine includes an inner liner and an outer liner positioned outward of the inner liner along the radial direction such that a combustion chamber is defined between the inner and outer liners. Furthermore, the combustor includes a fuel nozzle configured to supply fuel to the combustion chamber. Moreover, the combustor includes a bristle pack having a base plate and a plurality of bristles extending outward from the base plate such that the bristle pack forms at least a portion of at least one of a heat shield coupled to the fuel nozzle, a deflector of the combustor, or a flare of the combustor.

Abstract: Fuel nozzle assemblies are provided. For example, a fuel nozzle assembly for a combustor system comprises a fuel nozzle having a pilot swirler and an outlet defined in an outlet end, as well as a main mixer attached to the outlet end and extending about the outlet. A total combustor airflow through the combustor system comprises a pilot swirler airflow that is greater than about 14% and a main mixer airflow that is less than about 50% of the total combustor airflow. In further embodiments, the fuel nozzle also comprises main and pilot fuel injectors that each are configured to receive a portion of a fuel flow to the fuel nozzle. The fuel nozzle provides less than about 80% of the fuel flow to the main fuel injector at a high power operating condition of a gas turbine engine in which the fuel nozzle assembly is installed.

Abstract: A combustor for a gas turbine engine includes a forward liner segment and an aft liner segment positioned downstream from the forward liner segment relative to a direction of flow through the combustor. The forward and aft liner segments at least partially define a combustion chamber. Furthermore, the combustor includes a dilution slot frame positioned between the forward and aft liner segments along a longitudinal centerline of the gas turbine engine. Moreover, the dilution slot frame defines a plurality of dilution slots spaced apart from each other along a circumferential direction of the gas turbine engine such that the plurality of dilution slots provides an annular ring of dilution air to the combustion chamber.

Abstract: A gas turbine engine and combustor assembly are provided, the combustor assembly including a first liner and a second liner together defining at least in part a combustion chamber, wherein the first liner and the second liner are separated by a gap along the longitudinal direction, and wherein the first liner is forward of the second liner relative to a flow of fluid through the combustion chamber along the longitudinal direction, and wherein the gap is extended along the circumferential direction.

Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a forward liner segment; an aft liner segment disposed downstream from the forward liner segment relative to a direction of flow through the combustor, the forward and aft liner segments at least partially defining a combustion chamber; and an intermediate member disposed at least partially between the forward and aft liner segments and extending in the circumferential direction.

Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a liner at least partially defining a combustion chamber, the liner including a dilution slot in fluid communication with the combustion chamber; and one or more guides configured to redirect cooling medium into the dilution slot.

Abstract: A mixer assembly having a plurality of mixer vanes, each of the plurality of mixer vanes having a first end, a second end, and a body portion extending between the first end and the second end, the body portion having a length, a width, a thickness, a cross-sectional area, a curvature, and a twist, wherein each of the plurality of mixer vanes has a 3-dimensional shape defined by the length, width, thickness, cross-sectional area, curvature, and twist of the body portion, and wherein at least one of the plurality of mixer vanes has a non-uniform 3-dimensional shape.

Abstract: A swirler apparatus for a combustor, including: primary and secondary swirlers disposed axially adjacent to each other along a swirler centerline; the primary swirler including a plurality of primary swirl vanes arrayed around the swirler centerline, each primary swirl vane including opposed sides bounded between opposed forward and aft edges and opposed leading and trailing edges; wherein the forward edge is oriented at a first vane angle with respect to a radial direction; wherein the aft edge is oriented at a second vane angle with respect to the radial direction; wherein the second vane angle is different from the first vane angle; and the secondary swirler including a plurality of secondary swirl vanes arrayed around the swirler centerline.

Abstract: A combustor for a gas turbine engine includes a forward liner segment and an aft liner segment positioned downstream from the forward liner segment relative to a direction of flow through the combustor. The forward and aft liner segments at least partially define a combustion chamber. Furthermore, the combustor includes a dilution slot frame positioned between the forward and aft liner segments along a longitudinal centerline of the gas turbine engine. Moreover, the dilution slot frame defines a plurality of dilution slots spaced apart from each other along a circumferential direction of the gas turbine engine such that the plurality of dilution slots provides an annular ring of dilution air to the combustion chamber.

Abstract: A method for regulating jet wakes in a combustor including: directing compressed fluid into a passageway between a casing and a combustor liner; directing a combustion gas along a combustion zone; discharging a first portion of compressed fluid from the passageway into the combustion zone via a first through-hole which is disposed on a section of a combustor panel in a first circumferential row; and discharging a second portion of the compressed fluid from the passageway into the combustion zone via second through-holes, wherein the second through-holes are disposed on a section of the panel, spaced apart axially and circumferentially, and adjacent to the first through-hole, wherein the second through-holes include a first set of through-holes and a second set of through-holes, the first set of through-holes in a second row and the second set of through-holes in a third row, wherein the second and third rows extend circumferentially.

Abstract: A system and method for operating a combustion system comprising a fuel nozzle defining at least one main fuel circuit and at least one pilot fuel circuit is generally provided. The method includes determining an overall flow of fuel, the overall flow of fuel defining a sum total fuel through the main fuel circuit and the pilot fuel circuit; determining a plurality of ranges of ratios of main fuel flow through the main fuel circuit versus pilot fuel flow through the pilot circuit from the overall flow of fuel, wherein each range of ratios is based on a combustion criterion different from one another; determining a resultant range of ratios of main fuel flow versus pilot fuel flow based on a hierarchy of combustion criteria, wherein the hierarchy of combustion criteria provides a priority ranking of the combustion criterion; and flowing the overall flow of fuel to the main fuel circuit and the pilot fuel circuit based on the resultant range of ratios of main fuel flow versus pilot fuel flow.

Abstract: A hydrocarbon fluid containment article having a wall with a surface that is wetted by hydrocarbon fluid. The surface includes an anti-coking coating. The anti-coking coating includes a copper salt, a silver salt, or a combination thereof. A gas turbine engine component including a wall having a first surface and an anti-coking coating on the first surface of the wall that is wetted by hydrocarbon fluid. The anti-coking coating including a copper salt, a silver salt, or a combination thereof that prevents the formation of gum or coke on a surface thereon. Methods for reducing the deposition of thermal decomposition products on a wall of an article are also provided.

Abstract: A combustor for a gas turbine engine includes a forward liner segment and an aft liner segment positioned downstream from the forward liner segment relative to a direction of flow through the combustor. The forward and aft liner segments at least partially define a combustion chamber. Furthermore, the combustor includes a dilution slot frame positioned between the forward and aft liner segments along a longitudinal centerline of the gas turbine engine. Moreover, the dilution slot frame defines a plurality of dilution slots spaced apart from each other along a circumferential direction of the gas turbine engine such that the plurality of dilution slots provides an annular ring of dilution air to the combustion chamber.

Abstract: A combustor for a gas turbine engine includes an inner liner and an outer liner positioned outward of the inner liner along the radial direction such that a combustion chamber is defined between the inner and outer liners. Furthermore, the combustor includes a fuel nozzle configured to supply fuel to the combustion chamber. Moreover, the combustor includes a bristle pack having a base plate and a plurality of bristles extending outward from the base plate such that the bristle pack forms at least a portion of at least one of a heat shield coupled to the fuel nozzle, a deflector of the combustor, or a flare of the combustor.

Abstract: A fuel nozzle apparatus includes: an outer body having an exterior surface and a plurality of openings in the exterior surface. An inner body is disposed inside the outer body, cooperating with the outer body to define an annular space. A main injection ring is disposed in the annular space and includes an array of fuel posts extending outward therefrom, each fuel post including a perimeter wall defining a lateral surface and a recessed floor. Each fuel post is aligned with one of the openings and separated from the opening by a perimeter gap defined between the opening and the lateral surface. A main fuel gallery extends within the main injection ring. The main injection ring includes plurality of main fuel orifices, each main fuel orifice communicating with the main fuel gallery and extending through one of the fuel posts.

Abstract: A fuel nozzle apparatus includes: an outer body having an exterior surface and a plurality of openings in the exterior surface. An inner body is disposed inside the outer body, cooperating with the outer body to define an annular space. A main injection ring is disposed in the annular space and includes an array of fuel posts extending outward therefrom, each fuel post including a perimeter wall defining a lateral surface and a recessed floor. Each fuel post is aligned with one of the openings and separated from the opening by a perimeter gap defined between the opening and the lateral surface. A main fuel gallery extends within the main injection ring. The main injection ring includes plurality of main fuel orifices, each main fuel orifice communicating with the main fuel gallery and extending through one of the fuel posts.

Abstract: A method and structure for operating a combustion system of a gas turbine engine to mitigate low frequency combustion acoustics is generally provided. The method includes flowing an oxidizer through a fuel nozzle passage defining an inner wall and an outer wall, in which each of the inner wall and the outer wall are contoured from a first radius to a second radius smaller than the first radius; flowing the oxidizer at a higher axial velocity at the inner wall relative to the outer wall upstream of a fuel injection port; flowing a fuel through the fuel injection port to the fuel nozzle passage to mix with the flow of oxidizer to produce a fuel-oxidizer mixture; and igniting the fuel-oxidizer mixture downstream of the fuel injection port.

Abstract: A method and structure for operating a combustion system of a gas turbine engine to mitigate low frequency combustion acoustics is generally provided. The method includes flowing an oxidizer through a fuel nozzle passage defining an inner wall and an outer wall, in which each of the inner wall and the outer wall are contoured from a first radius to a second radius smaller than the first radius; flowing the oxidizer at a higher axial velocity at the inner wall relative to the outer wall upstream of a fuel injection port; flowing a fuel through the fuel injection port to the fuel nozzle passage to mix with the flow of oxidizer to produce a fuel-oxidizer mixture; and igniting the fuel-oxidizer mixture downstream of the fuel injection port.

Abstract: A fuel nozzle for a gas turbine engine is generally provided. The fuel nozzle includes an outer sleeve extended circumferentially around a fuel nozzle centerline and extended along a longitudinal direction co-directional to the fuel nozzle centerline. The outer sleeve defines a plurality of first radially oriented air inlet ports through the outer sleeve in circumferential arrangement relative to the fuel nozzle centerline. A centerbody is positioned radially inward of the outer sleeve. The centerbody includes an outer annular wall extended circumferentially around the fuel nozzle centerline and extended along the longitudinal direction co-directional to the fuel nozzle centerline. The outer sleeve further includes a first radial wall extended from a downstream end of the outer annular wall. The centerbody defines a plenum radially inward thereof. The outer annular wall and the outer sleeve together define a first fuel air mixing passage extended along the longitudinal direction therebetween.