Ansaldo Energia Switzerland Patente

The invention concerns a turbine for a gas turbine comprising a blade (10), a vane (20) and an abradable lip (40) attached to the blade (10) or to the vane (20), wherein the blade (10) and the vane (20) are separated by a gap (30) and the abradable lip (40) extends part of the distance across the gap (30). Embodiments include the addition of an abrasive layer (50, 52) attached on the other side of the gap (30) from the abradable lip (40). A method of manufacturing is also described.

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An apparatus for controlled delivery of cooling air to turbine blades (5) in a gas turbine (1) including a rotor unit (4), at least a first and second row of turbine blades (5), a plurality of rotor heat shield elements (23) disposed in a circumferential row between the first and second rows of turbine blades (5) and configured to define with said rotor unit (4) a rotor heat shield passage (33, 34, 35, 36) that is fed with cooling air, and a plurality of lock plates (41, 42, 43) disposed in a circumferential row and configured to secure at least one of said first and second rows (6, 7, 8) of turbine blades (5) to the rotor unit (4), wherein a radial portion (28) of the rotor shield elements (23) and/or the lock plates (41, 42, 43) include at least one opening (49) for controlled flow of cooling air.

Disclosed is a turboengine. The turboengine comprises an outer wall structure (11) and an inner wall structure (12), wherein the inner wall structure is provided at a radially inner position with respect to the outer wall structure, and each of the wall structures (11, 12) comprises a surface, said surfaces being arranged facing each other in the radial direction. A duct (14) is provided between the inner wall structure and the outer wall structure and having a downstream end with respect to a main working fluid flow direction of the turboengine. At least one guide vane member (20) is arranged at the downstream end of the duct, the guide vane member (20) comprising at least one airfoil, a radially inner end and a radially outer end. The inner end of the guide vane member is supported at the inner wall structure and the outer end of the guide vane member is supported at the outer wall structure. The inner wall structure (12) and the outer wall structure (11) are jointly provided as a vane carrier unit (1), wherein the inner wall structure and the outer wall structure are fixedly connected to each other by at least one bridging member (13) extending between the inner wall structure and the outer wall structure, and wherein the bridging member is provided upstream the downstream end of the duct with respect to the main working fluid flow direction.

A rotating blade for a gas turbine (5) is provided with: ¢ a root (17) configured to be coupled to a disc of the gas turbine (5) which rotates about an axis (A) of the gas turbine (5); a platform (20) coupled to the root (17); ¢ an airfoil (18) extending from the platform (20) and provided with a tip (23) radially opposite to the root (17) or the platform (20); the airfoil (18) comprising a leading edge (27), a trailing edge (28), a pressure side (25) and a suction side (26); ¢ at least one snubber (22; 22a, 22b) projecting from the suction side (26) and/or the pressure side (25); the snubber (22; 22a, 22b) being shaped so as to create a centripetal force towards the root (17) opposing the centrifugal force generated during the rotation of the disc to which the blade (15) is coupled.

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A fuel injector device (1) comprises a body, said body comprising a leading edge (11) and a trailing edge (12) and defining a streamwise direction (3) from the leading edge to the trailing edge. The fuel injector device body further comprises a first surface (101) and a second surface (102) opposite the first surface, each of said surfaces extending between and comprising the leading edge (11) and the trailing edge (12), and said surfaces conjoining each other at the leading edge and the trailing edge. The trailing edge (12), when seen in the streamwise direction, undulates along a trailing edge mean line (13) and, along its extent, deviates in opposite directions from said mean line and further comprises at least one inflection point (14) along its extent. The fuel injector device further comprises at least one first internal fuel supply means and at least one first fuel discharge duct fluidly connecting the at least one first internal fuel supply means with a first fuel discharge means (15) opening out to the exterior of the injector device, wherein at least one first fuel discharge duct terminates at a first fuel discharge means arranged at the trailing edge. At least one first fuel discharge means (15) is arranged at the trailing edge (12) between two trailing edge inflection points (14).

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Blade assembly (1) for gas turbine engines including a turbine blade (2) and a pyrometer device (3) at least partly housed inside the turbine blade (2); the pyrometer device (3) comprising: an oblong supporting body (11) which is accommodated inside a longitudinal inner cavity (4) that extends inside the turbine blade (2) behind a leading edge (6) of the turbine blade (2); and a number of transversal oblong ceramic caps (12) which jut out from said oblong supporting body (11) spaced side by side to one another so as to form a rake structure (13), and each of which engages in a pass-through manner a corresponding opening (14) formed in the turbine blade (2), at the leading edge (6) thereof, so that a segment of the oblong ceramic cap (12) protrudes outside of the turbine blade (2) in cantilever manner.

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Gas turbine unit having an axis (A) parallel to the main gas flow (M), the gas turbine unit comprising: - a blade (1) having a tip (2); - a stator heat shield (3) having an inner surface (4) facing the blade tip (2); wherein the inner surface (4) of the stator heat shield (3) and the blade tip (2) define a variable clearance depending on the applied thermal condition; wherein the blade tip (2) is configured to have a cylindrical shape along the axial direction in a hot running condition starting from a conical shape along the axial direction in a cold starting condition.

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An acoustic damping device is suggested comprising at least one opening (27) for a sealing gas (29) to deflect a grazing gas flow (19, 19.2) away from the mouth (17) of an acoustic damper (11).

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A combustor arrangement (10) for a gas turbine comprises a first burner (20), a first combustion chamber (21), a mixer (30) for admixing a dilution gas to the gases leaving the first combustion chamber (21) during operation, a second burner (60), and a second combustion chamber (40) arranged sequentially in a fluid flow connection. These elements of the combustor arrangement (10) are arranged in a row to form a flow path (27) extending between the first combustion chamber (21) and the second burner (60). The combustor arrangement (10) comprises acentral lance body (50) arranged inside the flow path and extending from the first burner (20) through the first combustion chamber (20) into the mixer (30) and into the second burner (60), wherein the lance body (50) comprises a fuel duct (28, 128, 62, 162) for providing fuel for the first burner (20) and/or for the second burner (60).

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