Abstract: The present invention relates to a gas turbine implemented for example at the interface between the combustor and the vane platform. An efficiency of a cooling film associated to the vane platform can be increased, hence reducing the quantity of the air needed.

Abstract: The invention concerns a turbine blade for a gas turbine having a platform part and an aerofoil part. The platform part includes a platform surface arranged to be attached to a corresponding aerofoil surface of the aerofoil part. The turbine blade further includes a cooling duct for cooling the platform and aerofoil surfaces, the cooling duct having at least one cavity in the platform surface and at least one cavity in the corresponding aerofoil surface, and the platform and aerofoil surface cavities are aligned such that when the platform surface and the aerofoil surface are touching, the cooling duct remains open. This provides a reliable cooling of both the platform and aerofoil surfaces.

Abstract: The disclosure pertains to a vane comprising a platform and airfoil extending form said platform and connected to the platform by a fillet. An impingement tube is inserted into said airfoil delimiting a cooling channel between the impingement tube and the side walls. The vane further comprises a baffle structure positioned adjacent the fillet and which follows the inside contour of the fillet; delimiting a first cooling passage between the fillet and the baffle structure. A first obstruction is arranged on the inside of the airfoil at the connection of the fillet to the side walls for separating the first cooling passage from the cooling channel in the airfoil and to guide the cooling gas from the first cooling passage into the impingement tube. The disclosure further refers to a method for cooling such a vane.

Abstract: A gas turbine including a combustion chamber and a first row of guide vanes, arranged essentially directly downstream thereof, of a turbine. The outer and/or inner limitation of the combustion chamber defined by at least one outer and/or inner heat shield, mounted on at least one combustion chamber structure arranged radially outside and/or inside. The hot gases flow path in the region of the guide vane row being restricted radially on the outside and/or inside by an outer and/or inner vane platform, mounted at least indirectly on at least one turbine carrier. A minimal gap size directly upstream of the first row of guide vanes is achieved by mounting at least indirectly on the turbine carrier at least one mini heat shield, arranged upstream of the first row of guide vanes and essentially adjacent the vane platform and in the flow direction between the heat shield and the vane platform.

Abstract: The invention refers to an arrangement for sealing an open cavity against hot gas entrainment. The open cavity is arranged between two axially adjacent stationary components limiting radially a hot gas path of a rotary flow machine, of which at least the downstream component carries at least one airfoil extending radially into said hot gas path. A multitude of air jets is directed into the open cavity such that hot gas flowing over the open cavity is prevented from penetrating into the open cavity completely. The invention has an arrangement of supply conduits for air jets which are directed into the open cavity with a jet momentum such that hot gas is induced not to enter the open cavity and forming at least one hot gas vortex close to the hot gas path within the open cavity and preventing hot gas from penetrating into the open cavity beyond the extension of the hot gas vortex.

Abstract: The invention refers to a rotor blade or guide vane airfoil for a gas turbine engine having a longitudinal axis and a source of cooling fluid. The airfoil has a pressure wall, a suction wall, a leading edge, a trailing edge and at least one cooling fluid flow passage. The cooling fluid flow passage is in fluid communication with the source of cooling fluid. Means for directing cooling fluid at least to the trailing edge are provided, whereas the cooling fluid flow passage including: a plurality of axially extending walls, each of the walls extending laterally between the pressure wall and suction wall. The plurality of walls are radially spaced within the cooling fluid flow passage such that adjacent pairs of walls define a channel. The axial spacing between the adjacent walls include in radial direction of the airfoil a pins and a ribs structure, wherein the pins holistic or approximately cover the axial height of the channel.

Abstract: A blade (10) for a gas turbine has a blade airfoil (11), the blade wall (18) of which encloses an interior space (17). For cooling the blade wall (18), the blade wall (18) includes a cooling arrangement (19) which has a radial passage (20) extending in the longitudinal direction of the blade and from which a multiplicity of cooling passages (21, 22), extending in the blade wall (18), branch in the transverse direction, and from which a multiplicity of film-cooling holes (23) are led to the outside in the transverse direction. Particularly efficient cooling is made possible by the distribution of the film-cooling holes (23) along the radial passage (20) being selected independently of the distribution of the cooling passages (21, 22) along the radial passage (20).

Abstract: The invention refers to an internally cooled casted airfoil for a rotary machine, preferably a gas turbine engine comprising a suction and a pressure side wall each extending in an axial direction, i.e. from a leading to a trailing edge region of said airfoil. At least one suction wall sided cooling channel extends in axial direction confined by the suction side wall and a first inner wall. At least one pressure wall sided cooling channel extends in axial direction confined by the pressure side wall and a second inner wall. At least one feed chamber is defined between the first and second inner wall for feeding said at least one suction and pressure sided cooling channel each by at least one through hole inside of the first and second inner wall. At least one suction wall sided cooling channel and the at least one pressure wall sided cooling channel extend into the trailing edge region separately.

Abstract: The invention refers to a fluid seal arrangement for constricting a leakage flow directed through a leakage gap bordered by a rotational and a stationary component including at least one nozzle opening in the rotating and/or stationary component facing towards the rotating or stationary component of an opposite side of the leakage gap respectively in order for injecting a liquid or gaseous fluid flow through the nozzle opening into the leakage gap. The at least one nozzle opening is fluidly connected to a cooling channel inside said rotating and/or stationary component, so that said fluid flow emanating at the nozzle opening consists of a cooling fluid of the rotating and/or stationary component exclusively.

Abstract: The invention relates to a component for a thermal machine, in particular a gas turbine, which includes a corner and/or edge subjected to a thermally high load. The cooling of the component is improved in a manner such that at least one cooling channel is countersunk into the surface of the component in the immediate vicinity of the corner and/or edge in order to cool the corner and/or edge.

Abstract: The invention refers to an arrangement for sealing an open cavity against hot gas entrainment. The open cavity is arranged between two axially adjacent stationary components limiting radially a hot gas path of a rotary flow machine, of which at least the downstream component carries at least one airfoil extending radially into said hot gas path. A multitude of air jets is directed into the open cavity such that hot gas flowing over the open cavity is prevented from penetrating into the open cavity completely. The invention has an arrangement of supply conduits for air jets which are directed into the open cavity with a jet momentum such that hot gas is induced not to enter the open cavity and forming at least one hot gas vortex close to the hot gas path within the open cavity and preventing hot gas from penetrating into the open cavity beyond the extension of the hot gas vortex.

Abstract: A cooled blade for a gas turbine includes a blade airfoil extending between leading and trailing edges in a flow direction and on suction and pressure sides is delimited by a wall, which include an interior space in which cooling air flows towards the trailing edge in the flow direction and discharges to the outside in the region of the trailing edge. The pressure-side wall terminates at a distance in front of the trailing edge in the flow direction, forming a pressure-side lip, such that the cooling air discharges from the interior space on the pressure side. Multiple ribs subdivides the interior space, parallel to the flow direction, into a plurality of parallel cooling passages which create a high pressure drop and in which turbulators are arranged for increasing cooling. Before the outlet, multiple flow barriers are provided in the cooling air flow path, distributed transversely to the flow direction.

Abstract: A blade for a gas turbine has a leading edge and a trailing edge, and an interior cavity, which is delimited by internal surfaces, for guiding cooling air therethrough. A multiplicity of turbulators or pins, which are formed on the wall, are arranged in a distributed manner in the region of the trailing edge and project from the internal surfaces into the cavity, to improve the transfer of heat between the wall of the blade and the cooling air. An improvement of the internal cooling is achieved by the turbulators or pins extending into the cavity in a direction which can be freely selected within an angular range.

Abstract: A gas turbine is provided including a compressor, a combustion chamber, a guide vane row and a rotor airfoil row. The guide vane row includes a plurality of guide vane airfoils having a blade and an inner platform. The ratio between the pitch and the leading edge diameter of the guide vane airfoils is between 6.3-7.6 and the ratio between the platform length and the leading edge diameter of the guide vane airfoils is between 4.0-5.5.

Abstract: A cooled constructional element for a gas turbine includes a wall having a front and a rear side. The front side is configured to be thermally loaded during operation of the turbine, and the rear side has a plurality of pins projecting therefrom in a two-dimensional distribution, the two-dimensional distribution including a higher density distribution of pins in a critical zone of the cooled constructional element than in the remaining regions of the cooled constructional element. A device is configured to create jets of a cooling medium that are directed onto the rear side of the wall in a region of the plurality of pins so as to cool the rear side of the wall by impingement.

Abstract: A blade (10) for a gas turbine has a blade airfoil (11), the blade wall (18) of which encloses an interior space (17). For cooling the blade wall (18), the blade wall (18) includes a cooling arrangement (19) which has a radial passage (20) extending in the longitudinal direction of the blade and from which a multiplicity of cooling passages (21, 22), extending in the blade wall (18), branch in the transverse direction, and from which a multiplicity of film-cooling holes (23) are led to the outside in the transverse direction. Particularly efficient cooling is made possible by the distribution of the film-cooling holes (23) along the radial passage (20) being selected independently of the distribution of the cooling passages (21, 22) along the radial passage (20).

Abstract: A blade for a gas turbine includes a leading edge running in a longitudinal direction substantially radially to an axis of the turbine; a trailing edge running in the longitudinal direction; a blade body disposed between the leading edge and the trailing edge so as to define a pressure side and a suction side; an exit slot disposed in the blade body in an area of the trailing edge and running in the longitudinal direction and configured to discharge a cooling medium from an interior of the blade body; and a row of first and second control elements disposed in the exit slot in a distributed manner in the longitudinal direction and configured to control a mass flow of the cooling medium exiting through the exit slot, the first control elements having a first configuration and the second control elements having a second configuration different from the first configuration.

Abstract: A gas turbine including a combustion chamber and a first row of guide vanes, arranged essentially directly downstream thereof, of a turbine. The outer and/or inner limitation of the combustion chamber defined by at least one outer and/or inner heat shield, mounted on at least one combustion chamber structure arranged radially outside and/or inside. The hot gases flow path in the region of the guide vane row being restricted radially on the outside and/or inside by an outer and/or inner vane platform, mounted at least indirectly on at least one turbine carrier. A minimal gap size directly upstream of the first row of guide vanes is achieved by mounting at least indirectly on the turbine carrier at least one mini heat shield, arranged upstream of the first row of guide vanes and essentially adjacent the vane platform and in the flow direction between the heat shield and the vane platform.

Abstract: A cooled component for a gas turbine is disclosed, which by an outer side of a wall delimits hot gas passage of the gas turbine and on an inner side has a device for impingement cooling. The impingement cooling device can include a multiplicity of impingement cooling chambers which are arranged next to each other, operate in parallel, are covered by impingement cooling plates which are equipped with impingement cooling holes, and are impinged upon by cooling air during operation.

Abstract: A cooled constructional element for a gas turbine includes a wall having a front and a rear side. The front side is configured to be thermally loaded during operation of the turbine, and the rear side has a plurality of pins projecting therefrom in a two-dimensional distribution, the two-dimensional distribution including a higher density distribution of pins in a critical zone of the cooled constructional element than in the remaining regions of the cooled constructional element. A device is configured to create jets of a cooling medium that are directed onto the rear side of the wall in a region of the plurality of pins so as to cool the rear side of the wall by impingement.