Abstract:
A gas turbine engine component includes an airfoil and a platform. The airfoil has a pressure side and an opposite suction side. The platform is connected to the airfoil and has a first curved edge adjacent the suction side and a second curved edge spaced from the pressure side such that more than half of the platform is located to the pressure side. The platform located to the pressure side has a non-axisymmetrical surface contouring.
Abstract:
A gas turbine engine component includes an airfoil extending radially from a root section to a tip section and having a trailing edge cooling passageway and first, second and third flow dividers in the cooling passageway. The first, second and third flow dividers have longitudinal axes that are angled based upon a position of the flow divider relative to the tip section of the airfoil.
Abstract:
A gas turbine engine component includes a wall with an inner face and an outer skin. A plurality of cooling air holes extend from the inner face to the outer skin. The cooling holes include an inlet merging into a metering section, and a diffusion section downstream of the metering section, and extend to an outlet at the outer skin. The diffusion section includes a plurality of lobes. A coating layer is formed on the outer skin, with at least a portion of the plurality of lobes formed within the thermal barrier coating. A method of forming such a component is also disclosed.
Abstract:
An airfoil includes a body that includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. A cooling passage is arranged interiorly of the exterior airfoil surface and provides an interior surface. The interior cooling surface includes micro-bumps that protrude from the interior cooling surface into the cooling passage. The micro-bumps are discrete from and noncontiguous relative to one another in multiple directions along the interior cooling surface. The micro-bumps may be provided while forming the airfoil or using correspondingly shaped micro-depressions on an airfoil core.
Abstract:
A method of manufacturing a fuel component for a gas turbine engine combustor includes additive manufacturing a sacrificial core and manufacturing a fuel component body at least partially around the sacrificial core. The sacrificial core is at least partially removed to at least partially define an internal geometry of the fuel component. An additively manufactured sacrificial core for a fuel component of a gas turbine engine combustor includes a first structure and a second structure. The first structure at least partially defines a first passage of the fuel component. The second structure at least partially defines a second passage of the fuel component. The second structure at least partially surrounds the first structure.
Abstract:
A composite component is disclosed. The composite component may comprise a metal plating deposited on a surface of the composite component, and a metallic feature adhesively bonded to the metal plating. The composite component may further comprise an adhesive layer between the metal plating and the metallic feature. The metal plating may provide a metal -to-metal interface between the surface of the composite component and the metallic feature.
Abstract:
A thermal management system for a gas turbine engine includes a heat exchanger in fluid communication with a geared architecture and a heating compartment.
Abstract:
An airfoil member and platform assembly for a gas turbine engine is provided. The platform to which the airfoil member is anchored is made of a composite material and includes an endwall defining a contoured region for improved aerodynamics. The contoured region influences the flow of gases through the flow passages between the airfoil members, thereby reducing endwall losses due to horseshoe vortexing. The composite material may be woven ceramic matrix composite fibers infiltrated with a ceramic matrix material or woven organic matrix composite fibers infiltrated with an organic matrix material.
Abstract:
A gas turbine engine component includes a wall having first and second wall surfaces and first and second cooling holes extending through the wall. The first and second cooling holes each include an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. Each diffusing section includes first and second lobes, each lobe diverging longitudinally and laterally from the metering section. The outlets of each cooling hole include first and second lateral ends and a trailing edge. One of the lateral ends of the outlet of the first cooling hole and one of the lateral ends of the outlet of the second cooling hole meet upstream of the trailing edge of the first cooling hole and the trailing edge of the second cooling hole.
Abstract:
A gas turbine engine component includes a wall having first and second wall surfaces, a cooling hole extending through the wall and a convexity. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally and laterally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section. The convexity is located near the outlet.