Abstract: A gas turbine (1) includes a combustion chamber (2) followed by a stator airfoil row (4) defining a plurality of guide vanes and separated by the combustion chamber (2) by a first gap (5), and a rotor airfoil row (6) separated by the stator airfoil row (4) by a second gap (7). The stator airfoils (15) of the stator airfoil row (4) are connected to guide vane boxes (17) collecting a cooling fluid (A) and injecting it through nozzles (20) in the second gap (7) to make it to enter rotor airfoil inlets (23). The guide vane boxes (23) are provided with passages (30) connecting a zone (31) upstream of the guide vane boxes (17) to a zone (32) of the second gap (7) downstream of the guide vane boxes (32). Moreover, the mouths (3) of the passages (30) facing the rotor airfoil row (6) are closer to a hot gases path than the nozzles (20).

Abstract: In a gas turbine, a plurality of burners, which are arranged concentrically to the rotational axis in a regular arrangement, each directing hot gas through an associated combustion chamber outlet into a turbine, at the inlet of which a second plurality of stator blades are arranged in a ring, uniformly spaced apart around the rotational axis. Cooling openings are provided, which are distributed over the circumference, through which cooling air is injected into the hot gas flow at the combustion chamber outlet. An improvement of the flow conditions in the hot gas is achieved by the cooling openings being divided into first cooling opening groups and second cooling opening groups. The arrangement of the first cooling opening groups corresponds to the arrangement of the stator blades, and in that the arrangement of the second cooling opening groups corresponds to the regular arrangement of the burners.

Abstract: A gas turbine is disclosed which includes an annular combustion chamber defined by an inner wall and an outer wall. A stator airfoil row can be defined by an annular inner stator wall and an annular outer stator wall housing a plurality of stator airfoils, and at least a rotor airfoil row defined by an annular inner rotor wall and an annular outer rotor wall housing a plurality of rotor airfoils. A gap is arranged, for example, between at least one of the inner stator wall and the inner combustion chamber wall, and the outer stator wall and the outer combustion chamber wall, upstream of said stator airfoil row. A border of at least one of the inner and outer stator wall facing the gap can be axisymmetric.

Abstract: A strip seal and method of configuration thereof for sealing two adjacent non-rotating gas turbine hot gas components exposed to pressure pulsations. The strip seal has at least two clamping projections distributed along discrete points of the strip seal length that extend out from a pressure face of the strip seal. The location of the projections are defined by two ratios. The first ratio, of less than 25, is the strip seal length extending free of clamping projections from any one of the ends of the strip seal to a clamping projection to the ratio of strip seal thickness. The second ratio, of less than 200, is the ratio of the strip seal length extending free of clamping projections between any two projections to the thickness of the strip seal. This seal arrangement ameliorates detrimental effects of induced resonance.

Abstract: A gas turbine with a stator blade is disclosed. The stator blade is fastened on a blade carrier and includes a blade airfoil which extends inwards in the radial direction from an outer platform into a hot gas passage. A cooling medium can flow through an interior of the blade airfoil, such as cooling air, which flows through an access in the blade carrier into a first plenum which is arranged above the outer platform, and from there, via an inlet which is provided in the outer platform, flows into the interior of the blade airfoil.

Abstract: A turbomachine (1), in particular a gas turbine, has a stator (2) and a rotor (3) and with at least one axial sealing device (6) arranged between them. On the stator side, the sealing device (6) has a radially stepped sealing contour (8) with regions (9, 10) projecting and retracting in the direction of the rotor (3), on the rotor side a plurality of sealing fins (11) projecting in the direction of the stator (2) being arranged, which engage in each case into adjacent retracting regions (10) of the stator-side sealing contour (8). On the rotor side, at least one additional fin (12) projecting in the direction of the stator (2) is provided, which is positioned between two adjacent sealing fins (11) arranged on the rotor side and which lies opposite a projecting region of the stator-side sealing contour (8).

Abstract: A gas turbine system includes a combustion chamber (2), a turbine (3), a radially inward and/or radially outward axial gap (4; 16, 17) between the combustion chamber (2) and the turbine (3), at which gap inner and/or outer combustion chamber walls (7, 8) and inner and/or outer turbine walls (11, 12) end, and a cooling gas supply (5), which via the gap (4; 16, 17) introduces a cooling gas into the turbine gas path (13) and/or into the combustion chamber gas path (9). An end section (21, 22, 23, 24) of the respective turbine wall (11, 12) and/or of the respective chamber wall (7, 8) adjoining the gap (4; 16, 17) is radially inwardly and/or radially outwardly, alternatingly formed.

Abstract: A gas turbine with a stator blade is disclosed. The stator blade is fastened on a blade carrier and includes a blade airfoil which extends inwards in the radial direction from an outer platform into a hot gas passage. A cooling medium can flow through an interior of the blade airfoil, such as cooling air, which flows through an access in the blade carrier into a first plenum which is arranged above the outer platform, and from there, via an inlet which is provided in the outer platform, flows into the interior of the blade airfoil.

Abstract: A gas turbine (1) includes a combustion chamber (2) followed by a stator airfoil row (4) defining a plurality of guide vanes and separated by the combustion chamber (2) by a first gap (5), and a rotor airfoil row (6) separated by the stator airfoil row (4) by a second gap (7). The stator airfoils (15) of the stator airfoil row (4) are connected to guide vane boxes (17) collecting a cooling fluid (A) and injecting it through nozzles (20) in the second gap (7) to make it to enter rotor airfoil inlets (23). The guide vane boxes (23) are provided with passages (30) connecting a zone (31) upstream of the guide vane boxes (17) to a zone (32) of the second gap (7) downstream of the guide vane boxes (32). Moreover, the mouths (3) of the passages (30) facing the rotor airfoil row (6) are closer to a hot gases path than the nozzles (20).

Abstract: In a gas turbine, a plurality of burners, which are arranged concentrically to the rotational axis in a regular arrangement, each directing hot gas through an associated combustion chamber outlet into a turbine, at the inlet of which a second plurality of stator blades are arranged in a ring, uniformly spaced apart around the rotational axis. Cooling openings are provided, which are distributed over the circumference, through which cooling air is injected into the hot gas flow at the combustion chamber outlet. An improvement of the flow conditions in the hot gas is achieved by the cooling openings being divided into first cooling opening groups and second cooling opening groups. The arrangement of the first cooling opening groups corresponds to the arrangement of the stator blades, and in that the arrangement of the second cooling opening groups corresponds to the regular arrangement of the burners.

Abstract: A gas turbine is disclosed which includes an annular combustion chamber defined by an inner wall and an outer wall. A stator airfoil row can be defined by an annular inner stator wall and an annular outer stator wall housing a plurality of stator airfoils, and at least a rotor airfoil row defined by an annular inner rotor wall and an annular outer rotor wall housing a plurality of rotor airfoils. A gap is arranged, for example, between at least one of the inner stator wall and the inner combustion chamber wall, and the outer stator wall and the outer combustion chamber wall, upstream of said stator airfoil row. A border of at least one of the inner and outer stator wall facing the gap can be axisymmetric.

Abstract: A strip seal and method of configuration thereof for sealing two adjacent non-rotating gas turbine hot gas components exposed to pressure pulsations. The strip seal has at least two clamping projections distributed along discrete points of the strip seal length that extend out from a pressure face of the strip seal. The location of the projections are defined by two ratios. The first ratio, of less than 25, is the strip seal length extending free of clamping projections from any one of the ends of the strip seal to a clamping projection to the ratio of strip seal thickness. The second ratio, of less than 200, is the ratio of the strip seal length extending free of clamping projections between any two projections to the thickness of the strip seal. This seal arrangement ameliorates detrimental effects of induced resonance.

Abstract: A gas turbine system, in particular for a utility power plant, includes a combustion chamber (2), which at least in an annular outlet region (6) having a chamber inner wall (7) and a chamber outer wall (8), radially delimits a combustion chamber gas path (9), a turbine (3), which at least in a stationary annular inlet region (10) having a turbine inner wall (11) and a turbine outer wall (12), radially delimits a turbine gas path (13), a radially inward and/or radially outward axial gap (4; 16, 17) between the combustion chamber (2) and the turbine (3), at which gap the inner and/or outer chamber wall (7, 8) and the inner and/or outer turbine wall (11, 12) end, and a cooling gas supply (5), which via the gap (4; 16, 17) introduces a cooling gas into the turbine gas path (13) and/or into the combustion chamber gas path (9).

Abstract: A turbomachine (1), in particular a gas turbine, has a stator (2) and a rotor (3) and with at least one axial sealing device (6) arranged between them. On the stator side, the sealing device (6) has a radially stepped sealing contour (8) with regions (9, 10) projecting and retracting in the direction of the rotor (3), on the rotor side a plurality of sealing fins (11) projecting in the direction of the stator (2) being arranged, which engage in each case into adjacent retracting regions (10) of the stator-side sealing contour (8). On the rotor side, at least one additional fin (12) projecting in the direction of the stator (2) is provided, which is positioned between two adjacent sealing fins (11) arranged on the rotor side and which lies opposite a projecting region of the stator-side sealing contour (8).

Abstract: In a turbomachine blade/vane for subsonic conditions with a convex curvature profile contour in the region of the leading edge and over the major part of the suction surface (S) and with a concave curvature profile contour in the region of the pressure surface (D), the profile contour in the suction-surface trailing edge region is designed to be concave (K). The momentum loss thickness mainly responsible for the profile loss is reduced in this region.