Abstract: Aspects of the disclosure are directed to a cooling design feature for inclusion in a liner of an aircraft, comprising: a plurality of angled holes, and at least one through hole separating all combinations of any two of the angled holes, wherein the at least one through hole is oriented at an angle that is substantially perpendicular to a surface of the liner, and wherein each of the plurality of angled holes are non-parallel to the at least one through hole.

Abstract: A combustor assembly includes a first wall, a second wall, a bulkhead and a plurality of fuel injectors. The bulkhead forms a combustion chamber with the first and the second walls. The fuel injectors are configured with the first wall in a unique and/or a fluctuating pattern.

Abstract: A fuel injection system may comprise a mixer and a fuel injector disposed within the mixer. The mixer may comprise an outer housing with an exit port and a bluff body extending across the exit port of the outer housing. A flared surface of the mixer may match a contour of the bluff body.

Abstract: A combustor is provided. The combustor may include an axial fuel injection system, and a radial fuel injection system aft of the axial fuel injection system. The axial fuel injection system includes a mixer having a bluff body at an exit port of the mixer, and a fuel injector disposed within the mixer. A fuel and air mixer is also provided and comprises an outer housing with an exit port and a bluff body. The bluff body extends across the exit port of the outer housing. A fuel injection system is also provided. The systems comprise a mixer having a bluff body at an exit port of the mixer and a fuel injector disposed within the mixer.

Abstract: A combustor of a gas turbine engine including a combustor shell having an interior surface, a first panel mounted to the interior surface at a first position and a second panel mounted to the interior surface at a second position. The first panel has a first end, a first combustion chamber surface parallel with the interior surface, a first rail extending from the first combustion chamber surface toward the interior surface of the combustor shell, and a first extension extending axially from the first rail to the end of the first panel. The second panel has a second end, a second combustion chamber surface, and a second rail extending from the second combustion chamber surface toward the interior surface of the combustor shell. The first end and the second end are proximal to each other and define a circumferentially extending gap there between.

Abstract: A combustor of a gas turbine engine including a combustor shell having an interior surface defining a combustion chamber, a first panel mounted to the interior surface at a first position, the first panel having a first surface and a first rail extending from the first surface toward the combustor shell, the first rail configured at a first angle relative to the first surface, and a second panel mounted to the interior surface at a second position axially adjacent to the first panel, the second panel having a second surface and a second rail extending from the second surface toward the combustor shell, the second rail configured at a second angle relative to the second surface. The first and second rails are proximal to each other and define a circumferential gap there between and at least one of the first or second angles is an acute angle.

Abstract: A turbine vane downstream of a combustor section includes an arcuate outer vane platform defined about an axis, the arcuate outer vane platform includes a segment of the arcuate outer vane platform along the axis which follows an outer combustor liner panel structure and an arcuate inner vane platform defined about the axis, the arcuate inner vane platform includes a segment of the arcuate inner vane platform along the axis which follows an inner combustor liner panel structure.

Abstract: A method of noise control from a combustor of a gas turbine engine includes selectively forming a plurality of local circumferential zones with different fuel-air ratios within the combustor.

Abstract: A combustor liner grommet is disclosed. The grommet may include a peripheral wall defining a hole in a combustor liner and further including at least one cooling air flow channel. The cooling air flow channel in the grommet wall may be a slot or a hole. The channel may increase cooling flow to the grommet and the combustor liner around the grommet to prevent cracking from heat stress.

Abstract: Aspects of the disclosure are directed to controlling a distribution of fuel to a plurality of nozzles associated with at least one aircraft engine by: determining a state of operation associated with the at least one aircraft engine, causing a valve coupled to a fuel supply and each of the nozzles to open when it is determined that the state of operation indicates a high power state relative to at least one threshold, and causing the valve to close when it is determined that the state of operation indicates a low power state relative to the at least one threshold.

Abstract: A liner panel for a combustor of a gas turbine engine includes a multiple of heat transfer augmentors. At least one of the multiple of heat transfer augmentors includes a cone shaped pin.

Abstract: A liner panel for use in a combustor for a gas turbine engine includes a forward region including a multiple of effusion passages each directed at an orientation consistent with a local swirl direction of combustion gases. Another liner panel for use in a combustor for a gas turbine engine includes a forward region forward of a dilution passage and an aft region aft of the forward region. The forward region includes a multiple of effusion passages each directed generally circumferentially. The aft region includes a multiple of effusion passages each directed at an orientation generally axially. A method of cooling a wall assembly within a combustor of a gas turbine engine includes orienting a multiple of effusion passages within a forward region consistent with a local swirl direction of combustion gases.

Abstract: Aspects of the disclosure are directed to a liner of an aircraft. The liner comprises a panel including at least one dilution hole, a rail coupled to the panel, and at least one effusion cooling hole located between a wall of the at least one dilution hole and a side of the rail. The liner may be associated with a combustion engine of the aircraft.

Abstract: The present disclosure relates to gas turbine engines, and in particular to combustor swirlers and swirler configurations. In one embodiment, a swirler includes an inner passage for receiving a fuel injector and a plurality of outer passages concentrically arranged around the inner passage. The plurality of outer passages include an outer vane assembly including a plurality of vane elements arranged at a first angle and a first middle vane assembly include a plurality of vane elements arranged at a second angle, wherein the outer vane assembly is concentrically arranged around the first middle vane assembly, and wherein the outer vane assembly and first middle vane assembly are configured to produce a low net-swirl to control the penetration depth and improve premixing of a fuel air mixture for the fuel injector. According to another embodiment, the swirler includes a second middle vane assembly include a plurality of vane elements arranged at a third angle.

Abstract: The present disclosure relates to combustor and heat shield configurations for gas turbine engines. A heat shield for a gas turbine engine can include one or more effusion holes downstream of a dilution hole to restore a cooling film applied to the heat shield in the combustor. Effusion holes may be clustered in one or more configurations and rows following each dilution hole and/or on a trailing edge of each heat shield panel structure. One or more embodiments are directed to effusion holes positioned along a trailing edge of a heat shield panel.

Abstract: A combustor module for a gas turbine engine is provided that includes a first annular liner assembly extending along a longitudinal axis of the engine. The first annular liner assembly includes a first annular support shell and a plurality of first heat shield panels coupled to the first annular support shell. The first heat shield panels form a segmented ring defining a plurality of first axial seams therebetween. The combustor module further includes a bulkhead coupled to the first annular liner assembly. The bulkhead provides a plurality of fuel nozzles for passing a first mass flow comprising fuel and air. The combustor module further includes a second annular liner assembly coupled to the bulkhead. The second annular liner assembly is in spaced-apart generally coaxial relationship from the first annular liner assembly by a channel height H. The second annular liner assembly includes an air admittance hole having a mean diameter D extending along a hole axis.

Abstract: A fuel injector nozzle assembly includes a body extending along an axis and a core swirl plug positioned at least partially within the body. The core swirl plug has a flow modifying structure configured to swirl fuel at a location upstream from a distal end of the nozzle assembly.

Abstract: A combustor is provided. The combustor may include an axial fuel injection system, and a radial fuel injection system aft of the axial fuel injection system. The axial fuel injection system includes a mixer having a bluff body at an exit port of the mixer, and a fuel injector disposed within the mixer. A fuel and air mixer is also provided and comprises an outer housing with an exit port and a bluff body. The bluff body extends across the exit port of the outer housing. A fuel injection system is also provided. The systems comprise a mixer having a bluff body at an exit port of the mixer and a fuel injector disposed within the mixer.

Abstract: Aspects of the disclosure are directed to a cooling design feature for inclusion in a liner of an aircraft, comprising: a plurality of angled holes, and at least one through hole separating all combinations of any two of the angled holes, wherein the at least one through hole is oriented at an angle that is substantially perpendicular to a surface of the liner, and wherein each of the plurality of angled holes are non-parallel to the at least one through hole.

Abstract: Aspects of the disclosure are directed to controlling a distribution of fuel to a plurality of nozzles associated with at least one aircraft engine by: determining a state of operation associated with the at least one aircraft engine, causing a valve coupled to a fuel supply and each of the nozzles to open when it is determined that the state of operation indicates a high power state relative to at least one threshold, and causing the valve to close when it is determined that the state of operation indicates a low power state relative to the at least one threshold.