Abstract: A guide vane for a turbomachine, having a blade airfoil and at least one platform, to which the blade airfoil is connected. A cooling channel system is provided for cooling the platform and the blade airfoil. The platform has, on the side thereof facing the airfoil, at least one sealing lip for sealing to a rotating system of the turbomachine. At least one cooling channel extends through the sealing lip, which cooling channel forms part of the cooling channel system.

Abstract: A guide vane for a turbomachine, having a blade airfoil and at least one platform, to which the blade airfoil is connected. A cooling channel system is provided for cooling the platform and the blade airfoil. The platform has, on the side thereof facing the airfoil, at least one sealing lip for sealing to a rotating system of the turbomachine. At least one cooling channel extends through the sealing lip, which cooling channel forms part of the cooling channel system.

Abstract: A refractory ceramic component for a gas turbine engine is employed that is more cost effective than typical components used in the gas turbine engine. The refractory ceramic component may be a refractory ceramic liner that is easily replaceable. The refractory ceramic liner may be a unitary construction or made of numerous bricks that are interlocked. The ceramic used is a refractory oxide material.

Abstract: A turbine airfoil (10) with an internal cooling system (12) having one or more bladders (14) forming near-wall cooling channels (16) is disclosed. The bladder (14) may be conformed to a shape of an inner surface (44) forming a cavity (18) within the internal cooling system (12). One or more standoff ribs (56) may extend radially inward from the inner surface (44) forming the cavity (18) to maintain the bladder (14) in position off of the inner surface (44) so that the near-wall cooling channel (16) is formed between the bladder (14) and the inner surface (44). The near-wall cooling channel (16) may be formed by inserting a bladder (14) into the cavity (18) in a first insertable position (22) and expanding the bladder (14) into a second expanded position (24). In at least one embodiment, the chamber (26) formed by the bladder (14) may be dead space that does not contain cooling fluids as a part of the cooling system (12).

Abstract: A hybrid die cast system (10) having an inner liner insert (12) that enables the configuration of a component (18) produced by the system (10) to be easily changed by changing the inner liner insert (12) without having to rework the die housing (16) is disclosed. Because the inner liner insert (12) only need be removed and replaced to change the configuration of an outer surface (18) of a component (18) produced by the system (10), the cost savings is significant in contrast with conventional systems in which the die would have to be reworked. The system (10) may also include a cooling system (20) for controlling the casting process by controlling the rate of solidification and the rate of cooling of the casting. Local heating and cooling may be used to control the microstructure, enhance mold fill and reduce casting defects such as porosity.

Abstract: A turbine engine airfoil structure including an airfoil adapted to be supported to extend across a gas passage for a hot working gas in a turbine engine. The airfoil structure further includes a platform structure located at one end of the airfoil and positioned at a location forming a boundary of the gas passage. The platform structure includes a platform member including a gas side surface extending generally perpendicular from the airfoil at a junction with the airfoil, and providing a structural connection to the airfoil. The platform structure further includes a separately formed platform cover attached to the platform member at the gas side surface. The platform cover extends from a location adjacent to one of the sidevvalls of the airfoil, and includes an outer surface located for contact with the hot working gas passing through the gas path.

Abstract: The present invention is directed to a hollow turbine airfoil having a cooling system designed to provide enhanced cooling to the fillet of a turbine airfoil. The turbine airfoil may include at least one fillet cooling channel, passing proximate to the fillet. A portion of the fillet cooling channel may be positioned proximate to the fillet outer surface without breaching an outer surface of the turbine airfoil. The turbine airfoil may include a vortex plate positioned adjacent to the end wall inner surface proximate to the fillet and an opening of the fillet cooling channel may be in fluid communication with the vortex chamber. The turbine airfoil may also include at least one end wall film cooling channel that may extend obliquely through the end wall and may be in fluid communication with the vortex chamber.

Abstract: The invention relates to a turbine blade or vane for a gas turbine, having a blade or vane root, which is successively adjoined by a platform region with a transversely running platform and then a blade or vane profile which is curved in the longitudinal direction, having a platform surface, which is provided at the platform and can be exposed to hot gas, and having at least one cavity, which is open on the root side, through which a coolant can flow and which extends through the blade or vane root and at least into the platform region and is surrounded by an inner wall, the contour of which, running in the platform region, is set back with respect to the contour running in the blade or vane root, so as to form a recess.

Abstract: The present invention is directed to a hollow turbine airfoil having a cooling system designed to provide enhanced cooling to the fillet of a turbine airfoil. The turbine airfoil may include at least one fillet cooling channel, passing proximate to the fillet. A portion of the fillet cooling channel may be positioned proximate to the fillet outer surface without breaching an outer surface of the turbine airfoil. The turbine airfoil may include a vortex plate positioned adjacent to the end wall inner surface proximate to the fillet and an opening of the fillet cooling channel may be in fluid communication with the vortex chamber. The turbine airfoil may also include at least one end wall film cooling channel that may extend obliquely through the end wall and may be in fluid communication with the vortex chamber.

Abstract: The invention relates to a gas turbine blade or vane ring having a supporting structure and gas turbine blades or vanes secured to it. The gas turbine blades or vanes have a blade or vane root, which is successively adjoined by a platform and then a blade profile which is curved in the longitudinal direction, the blade or vane root running in the longitudinal direction of the blade profile, and the platform having two platform longitudinal edges which are bent parallel and run in the longitudinal direction. To provide an alternative gas turbine blade or vane ring with simplified assembly, it is proposed that the blade or vane root be shaped in such a manner that the suction-side or pressure-side blade or vane root surface of the associated platform longitudinal edge be curved convexly or concavely, respectively. Moreover, the invention relates to the use of a gas turbine blade or vane ring of this type.

Abstract: A vortex-reducing fluid guide component is disclosed. The fluid guide component includes a base portion and an upstanding body portion. An internally-cooled blending element extends between the base portion and a leading edge of the body portion. The blending element includes a deflection region that advantageously disrupts horseshoe vortex formation. The fluid guide component includes a supplemental cooling channel between the blending element and the intersection of the body portion leading edge and the base portion. In one embodiment, the supplemental cooling channel captures cooling fluid used to cool the blending member and provides additional cooling to the body portion leading edge/base interface. The component includes a strategically-perforated impingement panel or member that distributes cooling fluid to selected portions of the blending member, thereby ensuring that key regions within the blending member are efficiently cooled in accordance with temperature distribution.

Abstract: A vortex-reducing fluid guide component is disclosed. The fluid guide component includes a base portion and an upstanding body portion. An internally-cooled blending element extends between the base portion and a leading edge of the body portion. The blending element includes a deflection region that advantageously disrupts horseshoe vortex formation. The fluid guide component includes a supplemental cooling channel between the blending element and the intersection of the body portion leading edge and the base portion. In one embodiment, the supplemental cooling channel captures cooling fluid used to cool the blending member and provides additional cooling to the body portion leading edge/base interface. The component includes a strategically-perforated impingement panel or member that distributes cooling fluid to selected portions of the blending member, thereby ensuring that key regions within the blending member are efficiently cooled in accordance with temperature distribution.