Ansaldo Energia Switzerland Patente

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A method for manufacturing a burner assembly for a gas turbine combustor, which is provided with a pilot burner (2) extending along a longitudinal axis (A) and a premix burner (3) surrounding the pilot burner (2), comprises the step of manufacturing at least one portion of the pilot burner (2) by means of an additive manufacturing technique; the step of manufacturing at least one portion of the pilot burner (2) comprises manufacturing at least one first thermal bridge (20; 20a, 20b, 20c) connecting parts (6, 7; 6, 8; 7, 10) of the pilot burner (2) having a temperature difference (T) not greater than a threshold value.

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The invention refers to a blade (1) of a rotary flow machine comprising an airfoil (2) having a suction surface (3) and a pressure surface joining each other along a trailing and a leading edge (4, 5), a radially outward directed airfoil tip and a radially inward directed end joining an inner platform (6) connecting the airfoil to a shank (7) at a radial end of the airfoil and providing, at least one shank pocket (8) radially encircled by an axially extending portion of the platform and at least one radially extending rim (9) that extends from the trailing edge side of the shank and has an essentially radially orientated first slot (10) for receiving a seal (11), and a mount (12) extending radially inwardly from said shank pocket. Said first slot has a first aperture (13) on a shank surface orientated in an axial direction.

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An assembly (1) of at least two members. One of the members (10; 20) supports the other member (20; 10), the assembly defining an axial direction (x), a radial direction (r), and a circumferential direction, an inner member (20) of said at least two members being received radially inside an outer member (10) of said at least two members, wherein the inner member and the outer member are attached to each other by a support arrangement, the support arrangement comprising at least one floating support assembly (40), wherein the floating support assembly provides a displaceable coupling between an inner member support point provided at the inner member and an outer member support point provided at the outer member. The floating support assembly (40) is arranged and configured to provide for a guided relative displacement between the inner member support point and the outer member support point along a displacement path (43), wherein the displacement path (43) of the relative displacement is inclined with respect to the axial direction (x) as well as to the radial direction (r) at a nonzero angle (), such that a relative displacement of the support points in the radial direction results in an interrelated relative displacement of the support points in the axial direction and vice versa.

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In a turbo engine, the rotor comprises a shaft, blades, and a shaft-blade connection means. The shaft blade connection means is provided in particular a fir tree connection, comprising a fir tree blade root and fir tree shaft grooves receiving the blade roots. The blade roots comprise lobes (260), while the grooves comprise undercut recesses receiving the lobes. The inclined mating bearing surfaces (161, 261) provided on a shaft post (150) or on a lobe (260), respectively, are offset with respect to each other in the circumferential direction (U) while partially overlapping. Each bearing surface is tangentially adjoined by a convexly curved end of bedding surface (162, 262).

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A turbine blade (12) of a gas turbine comprises a radially extending airfoil (29) with a suction side (26) and pressure side (27), which extend each in axial direction between a leading edge (24) and a trailing edge of said airfoil (29), whereby said leading edge (24) is cooled by means of impingement cooling with rows of radially distributed jets of a cooling medium impinging on the inner side of said leading edge (24), and whereby said row of radially distributed jets is generated at an internal web (16), which divides the hollow interior of the airfoil (29) into first and second cavities (15, 17), with the second cavity (17) being arranged at said leading edge (24). An enhanced cooling is achieved by said internal web (16) comprising two rows of radially distributed cooling medium supply holes (18, 19), through which cooling medium enters said second cavity (17) in form of impinging jets, whereby said cooling medium supply holes (18, 19) are oriented such that the directions of said jets of one row cross the directions of said jets of the other row.

A design for an effectively cooling a liner (7) of a gas turbine combustor by means of convective cooling is disclosed.

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A sealing arrangement at the interface between a combustor and a turbine of a gas turbine, said turbine comprising deflecting vanes (2) at its inlet, which deflecting vanes (2) are each mounted within said turbine so as to define an inner or outer diameter platform (9) and are in sealing engagement by means of an inner or outer diameter vane tooth (9a) with a seal (10) arranged at the corresponding inner or outer diameter part of the outlet of said combustor, wherein the seal (10) is movable and is pressed on said inner or outer diameter vane tooth (9a) by a differential pressure such that the pressure of the mainstream hot gas flow is a lower pressure (P2).

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The invention concerns a sequential liner (10) for a gas turbine combustor, comprising a sequential liner outer wall (12) spaced apart from a sequential liner inner wall (22) to define a sequential liner cooling channel between the sequential liner outer wall (12) and the sequential liner inner wall (22). The sequential liner outer wall (12) comprises a first face (14), a first adjacent face (16) and a second adjacent face (16), the first (16) and second (16) adjacent faces each being adjacent to the first face (14), the first face (14) of the sequential liner outer wall (12) comprising a first convective cooling hole (18) adjacent to the first adjacent face (16) and a second convective cooling hole (18) adjacent to the second adjacent face (16), each convective cooling hole (18) being arranged to direct a convective cooling flow into the sequential liner cooling channel adjacent to each adjacent face (16). The invention also concerns a method of cooling using the sequential liner (10) and a method of retrofitting a gas turbine.

Turbine part replacing apparatus for moving a combustor (1) of an industrial gas turbine from and to a mounting port of the gas turbine, the apparatus comprising a mobile frame (101) with a moving mechanism (112, 114) for substantially horizontal translation motion on the ground, a vertically extendable arm (120) attached to and extending from said mobile frame (101), and a carriage (15) attached to said extendable arm (120) such that the carriage (15) may be moved vertically with respect to the mobile frame (101). Moreover, the turbine part replacing apparatus is adapted for receiving said combustor (1) in said carriage (15) and for vertically moving said received combustor (1) by means of said extendable arm (120) from and to the mounting port.

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