Abstract: A flow body for a gas turbine includes an airfoil extending along a radial direction between a platform end and a tip which has a tip surface. The airfoil is formed of a first metal material and comprises an inner cavity for receiving a gaseous cooling fluid. The flow body further includes a squealer tip protruding from the tip surface. The squealer tip is formed from a second metal material and includes an internal cooling structure which is in fluid communication with the inner cavity. The squealer tip is joined to a contact surface of the tip by a transition layer that connects the contact surface and a remaining portion of the squealer tip. The transition layer, compared to at least one of the remaining portion of the squealer tip and the airfoil, has one of a reduced stiffness and an increased ductility in combination with reduced yield strength.

Abstract: A flow body for a gas turbine includes an airfoil extending along a radial direction between a platform end and a tip which has a tip surface. The airfoil is formed of a first metal material and comprises an inner cavity for receiving a gaseous cooling fluid. The flow body further includes a squealer tip protruding from the tip surface of the tip and extending along a circumference of the tip so that the squealer tip at least partially surrounds the tip surface. The squealer tip is formed from a second metal material and includes a plurality of internal cooling cavities that are separated from each other within the squealer tip, wherein each of the internal cooling cavities is in fluid communication with the inner cavity via one or more fluid passages.

Abstract: A method for manufacturing a blade for a gas turbine includes forming a blade body including an airfoil, forming a groove adjacent to a trailing edge of the airfoil in an outer surface of the airfoil, positioning a cover on the blade body such that it covers the groove and such that an outer surface of the cover forms a continuous surface with the outer surface of the airfoil, and joining the cover to the airfoil so that the cover and the groove define a cooling channel.

Abstract: A method for manufacturing a blade for a gas turbine includes forming a blade body, forming a groove in an outer surface of the blade body, positioning a cover on the blade body such that it covers the groove and such that an outer surface of the cover forms a continuous surface with the outer surface of the blade body, and joining the cover to the blade body so that the cover and the groove define a cooling channel.

Abstract: A blade for turbo machine is provided. The blade for a turbo machine includes an airfoil body extending in a radial direction between a root end and a tip end and including an inner void extending from the root end in the radial direction, a root body integrally formed with the airfoil body, extending from the root end of the air foil body to a bottom end in the radial direction, and including a receiving slot extending from the bottom end in the radial direction and opening into the inner void of the air foil body, and an insert positioned in the receiving slot of the root body and including a plurality of through holes extending in the radial direction to form a fluid connection to the inner void of the air foil body.

Abstract: The present technique presents a gas turbine blade for re-using cooling air, a turbomachine assembly having the blade, and a gas turbine having the turbomachine assembly. The blade includes a platform and an airfoil extending from the platform. The airfoil includes a pressure surface, a suction surface, a leading edge and a trailing edge. The platform includes a pressure side, a suction side, a leading-edge side and a trailing-edge side, disposed towards the pressure surface, the suction surface, the leading edge and the trailing edge of the airfoil, respectively. The suction side of the platform includes a part of the upper surface and a suction-side lateral surface of the platform. At least a part of an edge between the suction-side lateral surface and the upper surface of the platform comprises a chamfer part.

Abstract: A turbine blade is provided. The turbine blade may include an airfoil having an airfoil tip, a leading edge, a trailing edge, and a pressure side and a suction side extending from the leading edge to the trailing edge and defining an airfoil cavity, a squealer tip arranged at the airfoil tip part and comprising a trailing edge tip portion disposed at the trailing edge of the airfoil and a pressure side rail and a suction side rail meeting at the trailing edge tip portion and defining a squealer tip pocket at the airfoil tip, and at least one tip cooling hole disposed at the squealer tip pocket to provide cooling air from the airfoil cavity to the squealer tip pocket, wherein the trailing edge tip portion of the squealer tip includes a chamfer disposed towards the pressure side of the airfoil and a groove extending from the squealer tip pocket to the chamfer to provide cooling air from the squealer tip pocket to the chamfer.

Abstract: A blade for a turbo machine is provided. The blade for a turbo machine includes an airfoil body extending in a radial direction and including a suction side surface and a pressure side surface opposite to the suction side surface with respect to a circumferential direction extending across the radial direction, and a snubber structure including a first snubber element protruding in the circumferential direction from the suction side surface of the airfoil body and a second snubber element protruding in the circumferential direction from the pressure side surface of the airfoil body, wherein the first snubber element is connected to the suction side surface of the airfoil body by a concave curved first transition portion having a first radius, and the second snubber element is connected to the pressure side surface of the airfoil body by a concave curved second transition portion having a second radius, the first radius being smaller than the second radius.

Abstract: The present invention provides an airfoil 110 with the squealer tip cooling system 50 for a turbine blade 100 at the blade tip 113, wherein the squealer tip cooling system 50 comprises a cooling passage 170 arranged within a squealer tip 117, wherein the cooling passage 170 at least partly extends toward a terminal end 74 of the squealer tip 117, and a pocket 172 at a lateral surface 75, 76 of the squealer tip 117, open externally and extending inwardly at least partly across the cooling passage 170. The pocket 172 intersects the cooling passage 170 and the pocket 172 comprises an impingement surface 70 facing the cooling passage 170, on which a cooling medium expelled through the cooling passage 170 impinges before being discharged externally through the pocket 172.

Abstract: The present technique presents a blade 1 for a gas turbine 10. The blade 1 includes an airfoil 100 having an airfoil tip part 100a and a pressure side 102 and a suction side 104 meeting at a leading edge 106 and a trailing edge 108 and defining an internal space 100s of the airfoil 100. A squealer tip 80, 90 is arranged at the airfoil tip part 100a. The squealer tip 80, 90 comprises a suction side rail 90. The suction side rail 90 comprises a chamfer part 90x and at least one squealer tip cooling hole 99. The chamfer part 90x comprises a chamfer surface 9. An outlet 99a of the at least one squealer tip cooling hole 99 is disposed at the chamfer surface 9.

Abstract: A blade for a turbo machine is provided. The blade for a turbo machine includes an airfoil body extending in a radial direction and including a suction side surface and a pressure side surface opposite to the suction side surface with respect to a circumferential direction extending across the radial direction, and a snubber structure including a first snubber element protruding in the circumferential direction from the suction side surface of the airfoil body and a second snubber element protruding in the circumferential direction from the pressure side surface of the airfoil body, wherein the first snubber element is connected to the suction side surface of the airfoil body by a concave curved first transition portion having a first radius, and the second snubber element is connected to the pressure side surface of the airfoil body by a concave curved second transition portion having a second radius, the first radius being smaller than the second radius.

Abstract: A blade for turbo machine is provided. The blade for a turbo machine includes an airfoil body extending in a radial direction between a root end and a tip end and including an inner void extending from the root end in the radial direction, a root body integrally formed with the airfoil body, extending from the root end of the air foil body to a bottom end in the radial direction, and including a receiving slot extending from the bottom end in the radial direction and opening into the inner void of the air foil body, and an insert positioned in the receiving slot of the root body and including a plurality of through holes extending in the radial direction to form a fluid connection to the inner void of the air foil body.

Abstract: The present invention provides an airfoil 110 with the squealer tip cooling system 50 for a turbine blade 100 at the blade tip 113, wherein the squealer tip cooling system 50 comprises a cooling passage 170 arranged within a squealer tip 117, wherein the cooling passage 170 at least partly extends toward a terminal end 74 of the squealer tip 117, and a pocket 172 at a lateral surface 75, 76 of the squealer tip 117, open externally and extending inwardly at least partly across the cooling passage 170. The pocket 172 intersects the cooling passage 170 and the pocket 172 comprises an impingement surface 70 facing the cooling passage 170, on which a cooling medium expelled through the cooling passage 170 impinges before being discharged externally through the pocket 172.

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 present technique presents a blade 1 for a gas turbine 10. The blade 1 includes an airfoil 100 having an airfoil tip part 100a and a pressure side 102 and a suction side 104 meeting at a leading edge 106 and a trailing edge 108 and defining an internal space 100s of the airfoil 100. A squealer tip 80, 90 is arranged at the airfoil tip part 100a. The squealer tip 80, 90 comprises a suction side rail 90. The suction side rail 90 comprises a chamfer part 90x and at least one squealer tip cooling hole 99. The chamfer part 90x comprises a chamfer surface 9. An outlet 99a of the at least one squealer tip cooling hole 99 is disposed at the chamfer surface 9.

Abstract: The present technique presents a gas turbine blade for re-using cooling air, a turbomachine assembly having the blade, and a gas turbine having the turbomachine assembly. The blade includes a platform and an airfoil extending from the platform. The airfoil includes a pressure surface, a suction surface, a leading edge and a trailing edge. The platform includes a pressure side, a suction side, a leading-edge side and a trailing-edge side, disposed towards the pressure surface, the suction surface, the leading edge and the trailing edge of the airfoil, respectively. The suction side of the platform includes a part of the upper surface and a suction-side lateral surface of the platform. At least a part of an edge between the suction-side lateral surface and the upper surface of the platform comprises a chamfer part.

Abstract: Disclosed is a turbo-engine component comprising a wall, the wall comprising a hot gas side surface and a coolant side surface. At least one coolant discharge duct is provided in said wall and opening out onto the hot gas side surface at a coolant discharge opening. A coolant flow direction is defined from the interior of the coolant discharge duct towards the discharge opening, the coolant discharge duct further being delimited by a delimiting surface thereof provided inside the wall. The coolant discharge duct has a first cross sectional direction and a second cross sectional direction. The coolant discharge duct is a blind cavity and is closed towards the coolant side surface, and further a dimension of the coolant discharge duct measured in the first cross sectional direction decreases in the coolant flow direction.

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: 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: An internally cooled airfoil for a rotary machine, for example, a gas turbine engine includes a suction and pressure side wall each extending in an axial direction, i.e. from a leading to a trailing edge of the airfoil. A suction wall sided cooling channel and a pressure wall sided cooling channel extend in the axial direction. A feed chamber is defined between a first and second inner wall for feeding the suction wall and pressure wall sided cooling channel each by at least one through hole inside of the first and second inner wall. The suction wall sided cooling channel and the pressure wall sided cooling channel extend into the trailing edge region separately. The suction wall sided cooling channel and the pressure wall sided cooling channel join before discharging at the trailing edge.