Abstract: A cooling device for cooling platforms of a guide vane ring of a gas turbine is arranged downstream inside a main flow channel of a combustion chamber. Cooling air passages are arranged in a wall of the platforms or of an intermediate piece that is connected therewith to guide cooling air for film cooling the surfaces of the platforms. At least in certain areas, the wall is configured with at least two layers having—as viewed from the main flow channel—an outer wall and a spaced apart inner wall forming a hollow space, wherein the hollow space can be impinged by cooling air through at least one cooling air blow-in opening inside the outer wall, and at least one cooling air blow-out opening is arranged inside the inner wall extending in the downstream direction to the surfaces of the platforms.

Abstract: A turbine blade has a root portion, a platform and an aerofoil, the aerofoil is mounted on the platform and is formed by a pressure side wall and a suction side wall and has an outer surface, the pressure side wall and the suction side wall meet at a leading edge and a trailing edge, the aerofoil has an axial chord length, the suction side wall defines part of the radially outward surface of the aerofoil, the suction side wall defines an overhang, the overhang has a maximum overhang length that is between 5% and 20% of the axial chord length of the blade and is located between 5% and 50% of the suction surface length from the leading edge.

Abstract: A method for determining temperature in a flow channel of a gas turbine positioned on a testing bench includes: arranging at least one rod-shaped element provided with a thermal paint coating inside the flow channel; operating the gas turbine on the testing bench in a defined operating mode, wherein the thermal paint coating of the at least one rod-shaped element changes its color depending on the temperature that the thermal paint coating is exposed to; detecting the color distribution of the thermal paint coating of the at least one rod-shaped element; and determining the temperature that has been present in the flow channel along the at least one rod-shaped element in the defined operating mode based on the detected color distribution. Another embodiment relates to a measuring device having a rod-shaped element provided with a thermal paint coating.

Abstract: A turbine, in particular a high-pressure turbine, for an aircraft engine, with a housing at which turbine guide vanes are circumferentially arranged, wherein the turbine guide vanes have at least one interior space through which cooling air flows during operation of the turbine. At least one turbine guide vane has a cooling air passage in the area of an inner wall with respect to the radial direction of the turbine, via which the interior space of the turbine guide vane can be supplied with cooling air.

Abstract: A method for determining temperature in a flow channel of a gas turbine positioned on a testing bench includes: arranging at least one rod-shaped element provided with a thermal paint coating inside the flow channel; operating the gas turbine on the testing bench in a defined operating mode, wherein the thermal paint coating of the at least one rod-shaped element changes its color depending on the temperature that the thermal paint coating is exposed to; detecting the color distribution of the thermal paint coating of the at least one rod-shaped element; and determining the temperature that has been present in the flow channel along the at least one rod-shaped element in the defined operating mode based on the detected color distribution. Another embodiment relates to a measuring device having a rod-shaped element provided with a thermal paint coating.

Abstract: A gas turbine has a combustion chamber and a turbine nozzle guide vane ring. The combustion chamber includes an outer and inner combustion chamber walls with heat shield tiles. The guide vane ring has airfoils, and outer and inner platforms. The outer and/or inner platforms form a radial step towards a main flow path such that the upstream end of the outer platform is arranged radially outside and upstream of the downstream end of the heat shield tiles of the outer combustion chamber wall, and/or the upstream end of the inner platform is arranged radially inside and upstream of the downstream end of the heat shield tiles of the inner combustion chamber wall, so that the outer platform and the heat shield tiles of the outer combustion chamber wall and/or the inner platform and heat shield tiles of the inner combustion chamber wall overlap in the axial direction.

Abstract: A turbine stator wheel of a gas turbine with a plurality of stator vanes spaced apart around the circumference is provided. Two adjacent stator vanes form a passage each between the suction side of the one stator vane and the pressure side of the other stator vane starting from the vane trailing edge, which includes a constant passage portion in which the passage has a substantially constant passage cross-section. The constant passage portion has an inlet area and an outlet area. Each stator vane forms on the pressure side a rear area that extends, starting from the vane trailing edge adjoining the constant passage portion as far as the inlet area of the passage portion, and on the pressure-side forms a front area that extends upstream of the rear area. Each stator vane has on the pressure side a convex pressure-side contour.

Abstract: A cooling device for cooling platforms of a guide vane ring of a gas turbine is arranged downstream inside a main flow channel of a combustion chamber. Cooling air passages are arranged in a wall of the platforms or of an intermediate piece that is connected therewith to guide cooling air for film cooling the surfaces of the platforms. At least in certain areas, the wall is configured with at least two layers having—as viewed from the main flow channel—an outer wall and a spaced apart inner wall forming a hollow space, wherein the hollow space can be impinged by cooling air through at least one cooling air blow-in opening inside the outer wall, and at least one cooling air blow-out opening is arranged inside the inner wall extending in the downstream direction to the surfaces of the platforms.

Abstract: The present invention relates to a turbine rotor blade of a gas turbine with a blade tip, said blade tip having at least on its suction side, extending from a stagnation point on the blade leading edge to an intersection point of the suction-side profile line of the blade with a trailing-edge circle, an overhang which is substantially zero at the stagnation point and at the intersection point and which has a maximum value at around 40% of the running length of the suction-side overhang.

Abstract: A turbine, in particular a high-pressure turbine, for an aircraft engine, with a housing at which turbine guide vanes are circumferentially arranged, wherein the turbine guide vanes have at least one interior space through which cooling air flows during operation of the turbine. At least one turbine guide vane has a cooling air passage in the area of an inner wall with respect to the radial direction of the turbine, via which the interior space of the turbine guide vane can be supplied with cooling air.

Abstract: The present invention relates to a turbine blade of a gas turbine with an airfoil arranged on a blade root and having at least one cooling air duct running in the longitudinal direction of the turbine blade, arranged inside the turbine blade and extending through the blade root, characterized in that at least one swirl-generating element is arranged in the transitional area between blade root and airfoil in the cooling air duct, with the swirl-generating element including an outer ring and several swirl-generating stator vanes arranged thereon, which are connected to a centric area, as well as to a method for its manufacture.

Abstract: The present invention relates to a turbine rotor blade of a gas turbine with a blade tip, on which means for duct-type guidance of cooling air extending from a front suction-side area of the blade tip to a rear area of the blade tip are provided, and to a method for cooling a blade tip of a turbine rotor blade of a gas turbine, where air from a hot gas flow is guided from a front suction-side area of a blade tip to a rear area of the blade tip through a duct-type guidance.

Abstract: The present invention relates to a cooled turbine blade for a gas-turbine engine having at least one cooling duct (14) extending radially, relative to a rotary axis of the gas-turbine engine, inside the airfoil and air-supply ducts (12) issuing into said cooling duct, characterized in that the cooling duct (14) extends into the blade root (6) in order to generate close to the wall a cooling airflow moved at high circumferential velocity and radially in helical form and that in the area of the blade root (6) at least one nozzle-shaped air-supply duct (12) issues into the cooling duct (14) tangentially or with a tangential velocity component.

Abstract: A turbine stator wheel of a gas turbine with a plurality of stator vanes spaced apart around the circumference is provided. Two adjacent stator vanes form a passage each between the suction side of the one stator vane and the pressure side of the other stator vane starting from the vane trailing edge, which includes a constant passage portion in which the passage has a substantially constant passage cross-section. The constant passage portion has an inlet area and an outlet area. Each stator vane forms on the pressure side a rear area that extends, starting from the vane trailing edge adjoining the constant passage portion as far as the inlet area of the passage portion, and on the pressure-side forms a front area that extends upstream of the rear area. Each stator vane has on the pressure side a convex pressure-side contour.

Abstract: The present invention relates to a turbine rotor blade of a gas turbine with a blade tip, said blade tip having at least on its suction side, extending from a stagnation point on the blade leading edge to an intersection point of the suction-side profile line of the blade with a trailing-edge circle, an overhang which is substantially zero at the stagnation point and at the intersection point and which has a maximum value at around 40% of the running length of the suction-side overhang.

Abstract: The present invention relates to a turbine blade of a gas turbine with an airfoil arranged on a blade root and having at least one cooling air duct running in the longitudinal direction of the turbine blade, arranged inside the turbine blade and extending through the blade root, characterized in that at least one swirl-generating element is arranged in the transitional area between blade root and airfoil in the cooling air duct, with the swirl-generating element including an outer ring and several swirl-generating stator vanes arranged thereon, which are connected to a centric area, as well as to a method for its manufacture.

Abstract: The present invention relates to a cooled turbine blade for a gas-turbine engine having at least one cooling duct (14) extending radially, relative to a rotary axis of the gas-turbine engine, inside the airfoil and air-supply ducts (12) issuing into said cooling duct, characterized in that the cooling duct (14) extends into the blade root (6) in order to generate close to the wall a cooling airflow moved at high circumferential velocity and radially in helical form and that in the area of the blade root (6) at least one nozzle-shaped air-supply duct (12) issues into the cooling duct (14) tangentially or with a tangential velocity component.