Abstract: A grommet may define a dilution hole in a combustor panel. The grommet may comprise a ridge having a stepped geometry formed about an inner diameter of the grommet, the ridge comprising a passage. The passage may comprise an outlet. The ridge may further comprise a fillet about the inner diameter of the grommet, wherein the outlet is configured to direct a cooling flow circumferentially along the fillet and fill the ridge with the cooling flow.

Abstract: A grommet may define a dilution hole in a combustor panel. The grommet may comprise a ridge having a stepped geometry formed about an inner diameter of the grommet, the ridge comprising a passage. The passage may comprise an outlet. The ridge may further comprise a fillet about the inner diameter of the grommet, wherein the outlet is configured to direct a cooling flow circumferentially along the fillet and fill the ridge with the cooling flow.

Abstract: A component for a gas turbine engine includes a wall portion that includes an interior surface and an exterior surface. At least one cooling circuit is defined by the wall portion and includes a plurality of pedestals that extend across the cooling circuit. A width of the cooling circuit increases or decreases in a downstream direction. At least one cooling fluid inlet extends through the interior surface and is in fluid communication with the cooling circuit. At least one cooling fluid outlet extends through the exterior surface and is in fluid communication with the cooling circuit.

Abstract: A component for a gas turbine engine includes a wall portion that includes an interior surface and an exterior surface. At least one cooling circuit is defined by the wall portion and includes a plurality of pedestals that extend across the cooling circuit. A width of the cooling circuit increases or decreases in a downstream direction. At least one cooling fluid inlet extends through the interior surface and is in fluid communication with the cooling circuit. At least one cooling fluid outlet extends through the exterior surface and is in fluid communication with the cooling circuit.

Abstract: An airfoil may comprise an airfoil body having a leading edge, a trailing edge, an inner diameter end wall and an outer diameter end wall. A first cooling structure may be disposed within the airfoil body. The first cooling structure may comprise a first rib extending between the inner diameter end wall and the outer diameter end wall and may define a first radial passage configured to conduct a cooling airflow in a radial direction through the airfoil body. A second cooling structure may be disposed within the airfoil body. The second cooling structure may comprise a first baffle defining an axial passage configured to conduct the cooling airflow in an axial direction toward the trailing edge of the airfoil body.

Abstract: A grommet may define a dilution hole in a combustor panel. The grommet may comprise a set of upstream features such as, for example, a first inlet leading into a first plurality of passages with the first plurality of passages each having a first outlet directed towards an inner diameter of the grommet. A set of perimeter features may include a second inlet leading into a convective cooling passage having a plurality of internal cooling features. The convective cooling passage may further lead into a second plurality of passages each having a second outlet extending about the inner diameter of the grommet in a circumferential direction A set of downstream features may include a third inlet leading to a third plurality of passages each having a third outlet directed towards an outer diameter of the grommet.

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a wall having an internal surface and an outer skin, a cooling hole having an inlet extending from the internal surface and merging into a metering section, and a diffusion section downstream of the metering section that extends to an outlet located at the outer skin. The diffusion section of the cooling hole includes a first side diffusion angle, a second side diffusion angle and a downstream diffusion angle at a downstream surface of the diffusion section, the downstream diffusion angle being less than the first side diffusion angle and the second side diffusion angle.

Abstract: A grommet may define a dilution hole in a combustor panel. The grommet may comprise a ridge having a stepped geometry formed about an inner diameter of the grommet, the ridge comprising a passage. The passage may comprise an outlet. The ridge may further comprise a fillet about the inner diameter of the grommet, wherein the outlet is configured to direct a cooling flow circumferentially along the fillet and fill the ridge with the cooling flow.

Abstract: A grommet may define a dilution hole in a combustor panel. The grommet may comprise a set of upstream features such as, for example, a first inlet leading into a first plurality of passages with the first plurality of passages each having a first outlet directed towards an inner diameter of the grommet. A set of perimeter features may include a second inlet leading into a convective cooling passage having a plurality of internal cooling features. The convective cooling passage may further lead into a second plurality of passages each having a second outlet extending about the inner diameter of the grommet in a circumferential direction A set of downstream features may include a third inlet leading to a third plurality of passages each having a third outlet directed towards an outer diameter of the grommet.

Abstract: An airfoil may comprise an airfoil body having a leading edge, a trailing edge, an inner diameter end wall and an outer diameter end wall. A first cooling structure may be disposed within the airfoil body. The first cooling structure may comprise a first rib extending between the inner diameter end wall and the outer diameter end wall and may define a first radial passage configured to conduct a cooling airflow in a radial direction through the airfoil body. A second cooling structure may be disposed within the airfoil body. The second cooling structure may comprise a first baffle defining an axial passage configured to conduct the cooling airflow in an axial direction toward the trailing edge of the airfoil body.

Abstract: A gas turbine engine component includes a cooling hole. The component includes a first wall having an inlet, a second wall having an outlet and a metering section extending downstream from the inlet and having a substantially convex first surface and a substantially concave second surface. The component also includes a diffusing section extending from the metering section to the outlet. A gas turbine engine wall includes first and second surfaces and a cooling hole extending between an inlet at the first surface and an outlet at the second surface. The cooling hole includes a metering section commencing at the inlet and a diffusing section in communication with the metering section and terminating at the outlet. The metering section includes a top portion having a first arcuate surface and a bottom portion having a second arcuate surface. The first and second arcuate surfaces have arcs extending in substantially similar directions.

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a wall having an internal surface and an outer skin, a cooling hole having an inlet extending from the internal surface and merging into a metering section, and a diffusion section downstream of the metering section that extends to an outlet located at the outer skin. The diffusion section of the cooling hole includes a first side diffusion angle, a second side diffusion angle and a downstream diffusion angle at a downstream surface of the diffusion section, the downstream diffusion angle being less than the first side diffusion angle and the second side diffusion angle.

Abstract: An airfoil, and in a disclosed embodiment a rotor blade, has film cooling holes formed at a leading edge. A supplemental film cooling channel is positioned near the leading edge, but spaced toward the trailing edge from the leading edge. The supplemental film cooling channel directs film cooling air onto a suction wall. The supplemental film cooling channel air is generally directed to a location on the suction wall that has raised some challenges in the past. In a disclosed embodiment, the airfoil is provided with a thermal barrier coating, and the supplemental film cooling air protects this thermal barrier coating.

Abstract: A component for use in a gas turbine engine is provided. The component has an airfoil portion with a plurality of internal cooling passages and a non-linear trailing edge. The component further has a non-linear array of teardrop shaped assemblies which form a plurality of injection slots for injecting a coolant fluid into a fluid passing over the airfoil portion. The teardrop shaped assemblies are designed to maximize thermal performance of the component by reducing a relative diffusion angle between the injected coolant flow and the streamline direction of the fluid passing over the airfoil portion.

Abstract: The present invention provides a turbine blade having a revised under-platform structure including a unique coating combination that reduces mechanical stress factors within the turbine blade. The turbine blade includes a platform with an airfoil extending upwardly from the airfoil and a root portion extending downwardly from the platform. Two suction side tabs extend a first distance outward from a suction side of the root potion. Two pressure side tabs extend outward from a pressure side of the root portion. One of the two pressure side tabs extends outward a distance similar to the first distance, however, the other of the two pressure side tabs extends outward a distance much smaller than the first distance, which reduces stresses acting on the turbine blade. In addition, a plurality of coatings are systematically applied to the turbine blade to further reduce mechanical stress factors and improve cooling.

Abstract: The present invention provides an improved cooling circuit for a trailing edge of a turbine blade. The cooling circuit includes an inlet passage that receives a airflow and distributes the airflow through a feed passage. The feed passage primarily includes trip strips, at least one barrier including cross-over holes, teardrop shaped protrusions, and pockets disposed along a trailing edge. The geometry and positioning of both the cross-over holes and teardrop shaped protrusions downstream of the cross-over holes have been optimized to maximize cooling efficiency and reduce airflow. An improved transition between the inlet passage and the feed passage is also provided, which is arcuate and allows the airflow to maintain attachment and flow unimpeded from the inlet passage to the feed passage. The geometry of the pockets disposed along the trailing edge is optimized to improve cooling.

Abstract: An airfoil, and in a disclosed embodiment a rotor blade, has a serpentine cooling path. To best account for the Coriolis effect, the paths of the serpentine cooling channel have trapezoidal cross-sections. An area of the rotor blade between a smaller side of the trapezoidal-shaped paths, and a facing wall of the rotor blade has high thermal and mechanical stresses, and is a challenge to adequately cool. A microcircuit, which is a very thin cooling circuit having crossing pedestals, is embedded into the blade in this area. The microcircuit provides additional cooling, and addresses the challenges with regard to cooling these areas.

Abstract: An airfoil, and in a disclosed embodiment a rotor blade, has film cooling holes formed at a leading edge. A supplemental film cooling channel is positioned near the leading edge, but spaced toward the trailing edge from the leading edge. The supplemental film cooling channel directs film cooling air onto a suction wall. The supplemental film cooling channel air is generally directed to a location on the suction wall that has raised some challenges in the past. In a disclosed embodiment, the airfoil is provided with a thermal barrier coating, and the supplemental film cooling air protects this thermal barrier coating.

Abstract: An airfoil, and in a disclosed embodiment a rotor blade, has a serpentine cooling path. To best account for the Coriolis effect, the paths of the serpentine cooling channel have trapezoidal cross-sections. An area of the rotor blade between a smaller side of the trapezoidal-shaped paths, and a facing wall of the rotor blade has high thermal and mechanical stresses, and is a challenge to adequately cool. A microcircuit, which is a very thin cooling circuit having crossing pedestals, is embedded into the blade in this area. The microcircuit provides additional cooling, and addresses the challenges with regard to cooling these areas.

Abstract: A rotor blade is provided having a hollow airfoil and a root. The hollow airfoil has a cavity and one or more cooling apertures. The root is attached to the airfoil, and has a leading edge conduit, at least one mid-body conduit, and a trailing edge conduit. The conduits are operable to permit airflow through the root and into the cavity. Each conduit has a centerline. The leading edge conduit includes an inlet having a forward side, a suction side, and a pressure side that diverge from the centerline of the leading edge conduit, and an aft side. Each of the mid-body conduits includes an inlet having a suction side and a pressure side that diverge from the centerline of the mid-body conduit, and an aft side and a forward side. The trailing edge conduit includes an inlet having a suction side and a pressure side that diverge from the centerline of the trailing edge conduit. The trailing edge conduit inlet further includes a forward side and an aft side.