Abstract: An airfoil for a working fluid path of a turboengine extends along a spanwidth direction from a base to a tip. An aerodynamic body thereof includes a suction side surface, a pressure side surface, a leading edge, a trailing edge and a tip, the tip of the aerodynamic body having a tip cross section and a cross-sectional contour circumscribing the tip cross section. A rim extends to the tip of the airfoil and follows the cross-sectional contour on the pressure side, the suction side and extends over the leading edge of the airfoil, the rim delimiting a tip cavity which is open at the tip of the airfoil. The rim is further open at the trailing edge of the airfoil such that the tip cavity is open at the trailing edge of the airfoil.

Abstract: The present invention relates to an airfoil for a gas turbine, including an improved turbulator arrangement formed on an inner cooling channel of the airfoil. According to preferred embodiments of the invention, in order to ensure a constant angle of the cooling flow inside the channel relative to each turbulator, the angle formed between the turbulator and the vertical axis is advantageously adapted, in the curved area, for every single turbulator. Furthermore, the same principle may be applied to all the cooling channels present within the airfoil.

Abstract: An airfoil for a working fluid path of a turboengine extends along a spanwidth direction from a base to a tip. An aerodynamic body thereof includes a suction side surface, a pressure side surface, a leading edge, a trailing edge and a tip, the tip of the aerodynamic body having a tip cross section and a cross-sectional contour circumscribing the tip cross section. A rim extends to the tip of the airfoil and follows the cross-sectional contour on the pressure side, the suction side and extends over the leading edge of the airfoil, the rim delimiting a tip cavity which is open at the tip of the airfoil. The rim is further open at the trailing edge of the airfoil such that the tip cavity is open at the trailing edge of the airfoil.

Abstract: The blade for a gas turbine includes a root, a platform and an airfoil. The blade further has a cooling channel with an inlet located at the root or platform and outlets. The outlets are located at the platform.

Abstract: A cooled blade is disclosed for a gas turbine that includes a radially extending aerofoil with a leading edge, a trailing edge, a suction side and a pressure side, and wherein a lip overhang is provided on the suction side of the trailing edge The blade also includes a plurality of radial internal flow channels connected via flow bends to form a multi-pass serpentine for a coolant flow, whereby a trailing edge ejection region is provided for cooling the trailing edge, the trailing edge ejection region comprising a trailing edge passage of the multi-pass serpentine running essentially parallel to the trailing edge and being connected over its entire length with a pressure side bleed. An optimized cooling is achieved by also determining the cooling flow from the trailing edge passage to the pressure side bleed by means of a staggered field of pins.

Abstract: A gas turbine includes a rotor having a rotor groove and a rotor bore extending through the rotor, the rotor bore having a diffuser-shaped rotor bore exit. A blade is attached to the rotor and includes a blade tip having at least one dust hole. An airfoil has a leading edge and a trailing edge extending along a longitudinal axis of the blade between a lower end of the airfoil and the blade tip. A blade root is disposed at the lower end of the airfoil and is configured to be removably disposed in the rotor groove. The blade root includes a blade inlet having a cross sectional area that exceeds a cross sectional area of the rotor bore in at least one direction. A hollow blade core is disposed in the airfoil and extends along the longitudinal axis of the blade between the blade root and the blade tip.

Abstract: The present invention relates to an airfoil for a gas turbine, including an improved turbulator arrangement formed on an inner cooling channel of the airfoil. According to preferred embodiments of the invention, in order to ensure a constant angle of the cooling flow inside the channel relative to each turbulator, the angle formed between the turbulator and the vertical axis is advantageously adapted, in the curved area, for every single turbulator. Furthermore, the same principle may be applied to all the cooling channels present within the airfoil.

Abstract: A turbine blade of an axial turbine includes internal cooling fluid passages with radially outwardly extending passages connected to holes in the blade root. The holes are generally core printouts providing stability to the core during the casting process, but are not needed and need to be closed to guarantee the functioning of the cooling system. This is achieved by at least one covering plate. The plate is held by at least two slots located at the root of the turbine blade. Thus, the supply holes for cooling fluid located at the root section are closed by a simple mechanical device, e.g., a plate that does not require any subsequent brazing/welding operations. In addition, the plate is removable to facilitate inspection/cleaning, or further processing of the blade at service intervals.

Abstract: A stator blade for a gas turbine with sequential combustion, has a blade airfoil which extends in the radial direction between a blade tip and a shroud, with cooling passages extending inside the blade airfoil, through which a cooling medium can flow for cooling the blade and can then discharge from the stator blade into the hot gas flow flowing through the turbine. The blade airfoil has a sharply curved shape in space in the radial direction, and three cooling passages, which extend in the radial direction, arranged inside the blade airfoil in series in the hot gas flow direction and are interconnected by deflection regions, which are arranged at ends of the blade airfoil, so that the cooling medium flows through the cooling passages one after the other, with change of direction. The cooling passages follow the curvature of the blade airfoil in space in the radial direction.

Abstract: A turbine blade or vane includes at least one internal radial channel for the circulation of cooling medium bordered on a pressure side by a pressure side wall and on a suction side by a suction side wall joined at a upstream side at a leading edge and at and downstream side at the trailing edge. At least one exit hole extends through at least one of the pressure side wall or the suction side wall for blowing out of cooling medium from the internal radial channel to a medium surrounding the blade or vane. At least one trailing edge exit hole along the trailing edge has a surfacial exit opening disposed at the pressure side of the trailing edge.

Abstract: The invention relates to a cooled blade for a gas turbine that includes a radially extending aerofoil with a leading edge, a trailing edge, a suction side and a pressure side, and wherein a lip overhang is provided on the suction side of the trailing edge The blade also includes a plurality of radial internal flow channels connected via flow bends to form a multi-pass serpentine for a coolant flow, whereby a trailing edge ejection region is provided for cooling said trailing edge, said trailing edge ejection region comprising a trailing edge passage of said multi-pass serpentine running essentially parallel to said trailing edge and being connected over its entire length with a pressure side bleed.

Abstract: A blade for a gas turbine includes an airfoil extending in radial direction of the turbine or longitudinal direction of the blade, respectively, between a platform and a blade tip. The airfoil is bordered across the airfoil by a leading edge and a trailing edge and has a suction side and a pressure side. At the trailing edge a first cooling passage runs parallel to the trailing edge from the platform to the blade tip in the interior of the airfoil. The cooling passage is supplied with a cooling air flow from the platform side, and from which cooling air is discharged through a plurality of cooling holes arranged all over the blade. For such a blade the cooling is optimized by providing a first cooling passage, the passage area of which is tapered in radial direction by between 35% and 59%.

Abstract: A turbine blade or vane includes at least one internal radial channel for the circulation of cooling medium bordered on a pressure side by a pressure side wall and on a suction side by a suction side wall joined at a upstream side at a leading edge and at and downstream side at the trailing edge. At least one exit hole extends through at least one of the pressure side wall or the suction side wall for blowing out of cooling medium from the internal radial channel to a medium surrounding the blade or vane. At least one trailing edge exit hole along the trailing edge has a surfacial exit opening disposed at the pressure side of the trailing edge.

Abstract: A turbine component (25) includes a root (21), a tip (22), and an airfoil portion (7) having a leading and a trailing edge (8,9), an external suction side and pressure side wall (13, 14) between the leading and trailing edge. The walls enclose a central cavity (1-6) for the passage of cooling air, the cavity being partitioned into a leading edge- and a trailing edge region (7a, 7b) by at least one longitudinally extending first web (15) connecting the suction side wall with the pressure side wall and a second longitudinally extending web (16), connecting the first web with the suction side wall, thereby defining a first and second entry chamber (2, 3). The first web (15) is provided with at least one cross-over hole (H1) between the first entry chamber and the leading edge chamber (1), whereas the second web has no openings.

Abstract: A turbine blade of an axial turbine includes internal cooling fluid passages with radially outwardly extending passages connected to holes in the blade root. The holes are generally core printouts providing stability to the core during the casting process, but are not needed and need to be closed to guarantee the functioning of the cooling system. This is achieved by at least one covering plate. The plate is held by at least two slots located at the root of the turbine blade. Thus, the supply holes for cooling fluid located at the root section are closed by a simple mechanical device, e.g., a plate that does not require any subsequent brazing/welding operations. In addition, the plate is removable to facilitate inspection/cleaning, or further processing of the blade at service intervals.

Abstract: A gas turbine includes a rotor having a rotor groove and a rotor bore extending through the rotor, the rotor bore having a diffuser-shaped rotor bore exit. A blade is attached to the rotor and includes a blade tip having at least one dust hole. An airfoil has a leading edge and a trailing edge extending along a longitudinal axis of the blade between a lower end of the airfoil and the blade tip. A blade root is disposed at the lower end of the airfoil and is configured to be removably disposed in the rotor groove. The blade root includes a blade inlet having a cross sectional area that exceeds a cross sectional area of the rotor bore in at least one direction. A hollow blade core is disposed in the airfoil and extends along the longitudinal axis of the blade between the blade root and the blade tip.

Abstract: A cooled blade for a gas turbine includes an airfoil section which extends in the radial direction of the turbine or in the longitudinal direction of the blade between a platform and a blade tip which is provided with a cap. The airfoil section is bounded transversely with respect to the longitudinal direction by a leading edge and a trailing edge and has a pressure face and a suction face. Cooling channels extend in a radial direction between the platform and the blade tip in an interior of the airfoil section. The cooling channels can be acted upon by a cooling air flow from the platform. The blade tip is cooled by first cooling holes for convection cooling provided on the pressure face of the blade, and second cooling holes for film cooling provided on the suction side of the blade, through the cap of the blade, in the blade tip from the cooling channels, and distributed over the blade width.

Abstract: A blade for a gas turbine includes an airfoil extending in radial direction of the turbine or longitudinal direction of the blade, respectively, between a platform and a blade tip. The airfoil is bordered across the airfoil by a leading edge and a trailing edge and has a suction side and a pressure side. At the trailing edge a first cooling passage runs parallel to the trailing edge from the platform to the blade tip in the interior of the airfoil. The cooling passage is supplied with a cooling air flow from the platform side, and from which cooling air is discharged through a plurality of cooling holes arranged all over the blade. For such a blade the cooling is optimized by providing a first cooling passage, the passage area of which is tapered in radial direction by between 35% and 59%.

Abstract: A gas turbine airfoil (1) includes a pressure sidewall (15) and a suction sidewall (16), extending from a root to a tip and from a leading edge region to a trailing edge and having at least one cooling passage between the pressure sidewall (15) and the suction sidewall (16) for cooling air to pass through and cool the airfoil from within. One or several of the cooling passages (3) extend along the leading edge of the airfoil (1) and several film cooling holes (1,2) extend from the internal cooling passages (3) along the leading edge region to the outer surface of the leading edge region. The film cooling holes (1,2) each have a shape that is diffused in a radial outward direction of the leading edge of the airfoil (1) at least over a part of the length of the film cooling hole (1,2).

Abstract: A stator blade for a gas turbine with sequential combustion, has a blade airfoil which extends in the radial direction between a blade tip and a shroud, with cooling passages extending inside the blade airfoil, through which a cooling medium can flow for cooling the blade and can then discharge from the stator blade into the hot gas flow flowing through the turbine. The blade airfoil has a sharply curved shape in space in the radial direction, and three cooling passages, which extend in the radial direction, arranged inside the blade airfoil in series in the hot gas flow direction and are interconnected by deflection regions, which are arranged at ends of the blade airfoil, so that the cooling medium flows through the cooling passages one after the other, with change of direction. The cooling passages follow the curvature of the blade airfoil in space in the radial direction.