Abstract:
One exemplary embodiment of this disclosure relates to a gas turbine engine including a component. The component includes a platform having a mateface on a circumferential side thereof. The platform including a core passageway configured to communicate fluid to the mateface.
Abstract:
A method is provided for inspecting at least one aperture of a component with curable material and an inspection system. At least a portion of the curable material is injected into the aperture. The curable material conforms to at least a portion the aperture and subsequently cures and forms a mold of at least a portion of the aperture. The mold is removed from the aperture. At least a portion of a geometry of the mold is compared to at least a portion of a geometry of a reference model for the aperture using the inspection system.
Abstract:
Aspects of the disclosure are directed to a damper configured to be located in a cavity formed between first and second bases configured to seat respective first and second airfoils, the damper having a first face and a second face, where an aspect ratio between the first face and the second face ensures that the damper is installed in the cavity in accordance with a predetermined orientation.
Abstract:
An airfoil has a body that includes leading and trailing edges that adjoin pressure and suction sides to provide an exterior airfoil surface. A cooling passage extends in a radial direction from a root to a tip. A trailing edge cooling passage interconnects the cooling passage to the trailing edge. The trailing edge cooling passage includes first and second pedestals of different sizes that are arranged in a repeating pattern with respect to pedestals of the same size and with respect to pedestals of different sizes.
Abstract:
A turbine blade for a gas turbine engine includes an airfoil that extends in a first radial direction from a platform. A root extends from the platform in a second radial direction and has opposing lateral sides that provide a firtree-shaped contour. The contour includes first, second and third lobes on each of the lateral sides and that tapers relative to the radial direction away from the platform. The first, second and third lobes each provide contact surfaces arranged at about 45° relative to the radial direction. A contact plane on each lateral side at an angle of about 11° relative to the radial direction defining a contact point on each of the contact surfaces. The first, second and third lobes each include first, second and third grooves that are substantially aligned with one another along an offset plane spaced a uniform offset distance from the contact plane.
Abstract:
An assembly according to an exemplary aspect of the present disclosure includes, among other things, a disk, a cover plate providing a cavity at a first axial side of the disk, a passageway including an inlet provided by a notch in at least one of the disk and the cover plate in fluid communication with the cavity, and the passageway extending from the inlet to an exit provided at a second axial side of the disk opposite the first axial side, the exit in fluid communication with the inlet, and the passageway configured to provide fluid flow from the cavity to the exit.
Abstract:
A cast component includes a cast body that has a single crystal microstructure and an internal corner bounding an internal cavity. The single crystal microstructure defines a critical internal residual stress with respect to investment casting of the cast body using a refractory metal core beyond which the single crystal microstructure recrystallizes under a predetermined condition. The internal corner has a corner radius that is greater than a critical corner radius below which an amount of internal residual stress in the single crystal microstructure exceeds the critical internal residual stress. The internal cavity includes a cross section less than about 20 mils near the corner radius.
Abstract:
A gas turbine engine component comprises a blade having a leading edge and a trailing edge. The blade is mounted to a disc and configured for rotation about an axis. A platform supports the blade, and has a fore edge portion at the leading edge and an aft edge portion at the trailing edge. At least one of the fore edge portion and aft edge portion includes a mouth portion defined by an inner wing and an outer wing spaced radially outward of the inner wing. At least one coverplate is retained against the disc by the inner wing. A gas turbine engine is also disclosed.
Abstract:
A rotor blade according to an exemplary aspect of the present disclosure includes, among other things, a platform, an airfoil that extends radially from the platform, a first cooling core that extends at least partially inside the airfoil, a second cooling core inside of the platform, a first cooling hole that extends circumferentially between a mate face of the platform and the second cooling core, a second cooling hole that extends between a gas path surface of the platform and the second cooling core, the second cooling core radially disposed between the gas path surface and a non-gas path surface, and the second cooling core circumferentially disposed between the first cooling core and the mate face. A method of cooling a blade is also disclosed.
Abstract:
A rotor blade according to an exemplary aspect of the present disclosure includes, among other things, a platform, an airfoil that extends from the platform, a first cooling core that extends at least partially inside the airfoil, a second cooling core inside of the platform and a first cooling hole that extends between a mate face of the platform and the second cooling core.