A combustor assembly for a gas turbine engine includes a combustor dome defining an opening, a cooling hole, and at least in part defining a combustion chamber. The combustor dome includes a first side and a second side, the cooling hole extending from the first side to the second side. The combustor assembly additionally includes a fuel-air injector hardware assembly positioned at least partially within the opening of the combustor dome and including a heat shield. The heat shield includes a heat deflector lip. The cooling hole in the combustor dome is oriented to direct a cooling airflow onto the heat deflector lip to maintain at least a portion of the heat shield within a desired operating temperature range.

A combustor assembly for a gas turbine engine includes a combustor dome defining an opening, a cooling hole, and at least in part defining a combustion chamber. The combustor dome includes a first side and a second side, the cooling hole extending from the first side to the second side. The combustor assembly additionally includes a fuel-air injector hardware assembly positioned at least partially within the opening of the combustor dome and including a heat shield. The heat shield includes a heat deflector lip. The cooling hole in the combustor dome is oriented to direct a cooling airflow onto the heat deflector lip to maintain at least a portion of the heat shield within a desired operating temperature range.

A combustor of an industrial gas turbine engine having a combustor secured within a combustor casing with a combustor cavity surrounding the combustor, and a flow liner forming a cooling air space between the casing and the flow liner in which high pressure cooling air can be passed to provide insulation to the casing from the high temperature gas surround the combustor. The flow liner can include a TBC or a layer of insulation to limit heat buildup of the cooling air flowing through the space to further insulate the casing.

A combustor of an industrial gas turbine engine having a combustor secured within a combustor casing with a combustor cavity surrounding the combustor, and a flow liner forming a cooling air space between the casing and the flow liner in which high pressure cooling air can be passed to provide insulation to the casing from the high temperature gas surround the combustor. The flow liner can include a TBC or a layer of insulation to limit heat buildup of the cooling air flowing through the space to further insulate the casing.

A cooling system (10) for a fuel system in a turbine engine (14) that is usable to cool a fuel nozzle (16) is disclosed. The cooling system (10) may include one or more cooling system housings (18) positioned around the fuel nozzle (16), such that the cooling system housing (18) forms a cooling chamber (20) defined at least partially by an inner surface (22) of the cooling system housing (18) and an outer surface (24) of the fuel nozzle (16). The fuel nozzle (16) may extend into a combustor chamber (26) formed at least in part by a combustor housing (32). The fuel nozzle (16) may include one or more fuel exhaust orifices (28) with an opening (30) in an outer surface (24) of the fuel nozzle (16) and configured to exhaust fluids unrestricted by the housing (18) forming the cooling system cooling chamber (20). The cooling system (10) may provide cooling fluids to cool the fuel nozzle (16) within the cooling system cooling chamber (20) regardless of whether the fuel nozzle (16) is in use.

The jet engine comprising a ramjet air path extending from an intake, into a combustion chamber, and out an exhaust nozzle, a fuel inlet leading into the combustion chamber, an oxidizer inlet leading into the combustion chamber and a partition being operable to selectively close the ramjet air path upstream of the combustion chamber to allow operation of the jet engine in rocket mode and open the ramjet air path to allow operation of the jet engine in ramjet mode.

The jet engine comprising a ramjet air path extending from an intake, into a combustion chamber, and out an exhaust nozzle, a fuel inlet leading into the combustion chamber, an oxidizer inlet leading into the combustion chamber and a partition being operable to selectively close the ramjet air path upstream of the combustion chamber to allow operation of the jet engine in rocket mode and open the ramjet air path to allow operation of the jet engine in ramjet mode.

The present disclosure is directed to a combustor having an annularly shaped liner that at least partially defines a hot gas path of the combustor. A flow sleeve circumferentially surrounds at least a portion of the liner. The flow sleeve is radially spaced from the liner to form a cooling flow annulus therebetween. A bluff body extends radially between the flow sleeve and the liner through the cooling flow annulus. A guide vane is disposed within the cooling flow annulus and extends between the flow sleeve and the liner proximate to the bluff body.

The present disclosure is directed to a combustor having an annularly shaped liner that at least partially defines a hot gas path of the combustor. A flow sleeve circumferentially surrounds at least a portion of the liner. The flow sleeve is radially spaced from the liner to form a cooling flow annulus therebetween. A bluff body extends radially between the flow sleeve and the liner through the cooling flow annulus. A guide vane is disposed within the cooling flow annulus and extends between the flow sleeve and the liner proximate to the bluff body.

A conjoined grommet assembly for a combustor wall assembly of a gas turbine engine has a first grommet defining at least in-part a first dilution hole, and a second grommet defining at least in-part a second dilution hole. The first and second dilution holes are spaced closely together such that the first grommet is in contact with the second grommet.