Abstract: A gas turbine includes a turbine housing; a rotatable turbine rotor with turbine blades; and at least one heat shield segment for a stator. The heat shield segment is disposed radially between the rotor and the housing and attached to the housing and has a profile with a curved section in at least one region in an axial direction of the gas turbine and a radial outer surface. The heat shield segment includes a rib, a first boss, and a second boss disposed on the radial outer surface.

Abstract: A guide vane for a gas turbine includes a vane body having a leading edge and a trailing edge and a shroud extending at least between the leading edge and the trailing edge. The shroud has a first side wall and a second side wall extending essentially radially and in a longitudinal direction of the gas turbine. The first side wall has a groove disposed in a region of the trailing edge extending in a longitudinal direction of the shroud and is configured to receive a sealing plate. The first side wall also has a clearance extending from the groove in a region of the trailing edge, wherein a depth of the clearance in a circumferential direction of the gas turbine is equal to a depth of the groove, and wherein a width of the clearance in the longitudinal direction of the shroud is between one to three times the depth of the clearance.

Abstract: A cast turbine blade is disclosed and includes an internal cooling passage that passes (e.g., zig-zags or meanders) through the blade from an inlet in the blade root to an outlet in the blade tip. The cooling passage can have a zone at a bend that is at a distance which is remote from the inlet of the cooling passage when the distance from the inlet is measured around the passage, but that is closer to the inlet when the distance from the inlet is measured in a straight line. During casting of the blade, the cooling passage can be defined by a core or cores having a leachable material, the cores being removed after casting by a chemical leaching process. A supplementary passage is also provided for connecting the remote zone to the inlet during the leaching process. The supplementary passage can likewise be defined by a leachable core, or it may be machined into the blade after casting.

Abstract: A seal assembly for a gas turbine is arranged in grooves of a rotor heat shield having several bends. The assembly comprises four seal portions overlapping one another and extending in the axial, radial, and circumferential direction with respect to the turbine rotor. A holding means retains the radial sections of one seal portion allowing a limited movement of said seal portion independent of another seal portion. The independent movement assures contact of the individual seal portions with all mating surfaces of the rotor heat shield and improved sealing function regardless of displacements of the rotor heat shield and tolerances of turbine parts.

Abstract: A gas turbine (1) includes a rotor (2) which has two rotor blade rows (5) with a plurality of rotor blades (6), and also a rotor heat shield (7), which is arranged between them, with a plurality of heat shield elements (12), and with a stator (3) which has a stator blade row (8), with a plurality of stator blades (9), which is arranged axially between the two adjacent rotor blade rows (5). The stator blades (9) have a stator sealing structure (10) radially on the inside. The heat shield elements (12) have a rotor sealing structure (13) radially on the outside which interacts with the stator sealing structure (10) for forming an axial seal (14). Furthermore, a blade radial seal (15) is formed between two adjacent rotor blades (6), and also a heat shield radial seal (16) is formed between two adjacent heat shield elements (12), and in each case separates a gas path (17) from the rotor (2).

Abstract: A fluid flow engine (1), in particular a turbo engine, has at least one guide vane row (5) with a plurality of guide vanes (6) and at least one rotor blade row (7) with a plurality of rotor blades (8). One guide vane row (5) and one rotor blade row (7) that follows the former directly downstream with respect to a working gas flow (11), together form a stage (10) of the fluid flow engine (1).

Abstract: A stator heat shield (6) for a gas turbine (1) includes an outside (11) which in the installed state faces a hot gas path (9) of the gas turbine (1), and an inside (12) facing away from the outside (11). A plurality of ribs (13) are formed on the inside (12) and extend axially in the installed state with regard to an axis of rotation (8) of a rotor (3) of the gas turbine (1) and, in the circumferential direction, are spaced a distance apart from one another. At least one baffle plate (14) is arranged on the inside (12) and is in contact with the ribs (13). At least one groove (16) is designed in an end face (15) bordering the stator heat shield (6) in the circumferential direction, into which at least one sealing element (18) can be inserted. A plurality of bores (19) are included, each opening at a distance from the groove (16) in the direction of the outside (11) at the inside (12) at one end and on the end face (15) at the other end and arranged a distance apart from one another in the axial direction.

Abstract: A cast turbine blade is disclosed and includes an internal cooling passage that passes (e.g., zig-zags or meanders) through the blade from an inlet in the blade root to an outlet in the blade tip. The cooling passage can have a zone at a bend that is at a distance which is remote from the inlet of the cooling passage when the distance from the inlet is measured around the passage, but that is closer to the inlet when the distance from the inlet is measured in a straight line. During casting of the blade, the cooling passage can be defined by a core or cores having a leachable material, the cores being removed after casting by a chemical leaching process. A supplementary passage is also provided for connecting the remote zone to the inlet during the leaching process. The supplementary passage can likewise be defined by a leachable core, or it may be machined into the blade after casting.

Abstract: A guide vane for a gas turbine includes a vane body having a leading edge and a trailing edge and a shroud extending at least between the leading edge and the trailing edge. The shroud has a first side wall and a second side wall extending essentially radially and in a longitudinal direction of the gas turbine. The first side wall has a groove disposed in a region of the trailing edge extending in a longitudinal direction of the shroud and is configured to receive a sealing plate. The first side wall also has a clearance extending from the groove in a region of the trailing edge, wherein a depth of the clearance in a circumferential direction of the gas turbine is equal to a depth of the groove, and wherein a width of the clearance in the longitudinal direction of the shroud is between one to three times the depth of the clearance.

Abstract: A gas turbine includes a turbine housing; a rotatable turbine rotor with turbine blades; and at least one heat shield segment for a stator. The heat shield segment is disposed radially between the rotor and the housing and attached to the housing and has a profile with a curved section in at least one region in an axial direction of the gas turbine and a radial outer surface. The heat shield segment includes a rib, a first boss, and a second boss disposed on the radial outer surface.

Abstract: A fluid flow engine (1), in particular a turbo engine, has at least one guide vane row (5) with a plurality of guide vanes (6) and at least one rotor blade row (7) with a plurality of rotor blades (8). One guide vane row (5) and one rotor blade row (7) that follows the former directly downstream with respect to a working gas flow (11), together form a stage (10) of the fluid flow engine (1).

Abstract: A guide vane arrangement of a turbomachine includes a casing, at least one guide vane carrier connected to the casing and having a first carrying portion and a second carrying portion spaced apart axially from the first carrying portion, a plurality of guide vanes connected to the guide vane carrier and arranged next to each other in a circumferential direction, and at least one securing element connected to the guide vane carrier and configured to provide an axial fixing of at least one of the guide vanes. Each guide vane has a platform including a first locking portion and a second locking portion spaced apart axially from the first locking portion, wherein each of the first and second locking portions and the respective first and second carrying portion are adapted to provide an axially pluggable and a radially positive fastening between the guide vane carrier and the respective guide vane.

Abstract: A vane arrangement (1) of a turbo machine (4), in particular of a gas turbine, includes at least one vane carrier (2) attached to a housing (5) of a turbo machine (4), and several vanes (3) attached to the vane carrier (2) and circumferentially arranged side by side, wherein the attachment between the vane carrier (2) and respective vane (3) is adapted for being axially pluggable. In order to simplify dismounting of a single vane (3) at least one of the vanes (3) is provided with an extraction device (35) having at least one axially extending extraction through-hole (36). The respective extraction through hole (36) is provided with an inner thread (37), and the vane carrier (2) is provided with an abutment zone (38) axially opposing the extraction through hole (36).

Abstract: A stator heat shield (6) for a gas turbine (1) includes an outside (11) which in the installed state faces a hot gas path (9) of the gas turbine (1), and an inside (12) facing away from the outside (11). A plurality of ribs (13) are formed on the inside (12) and extend axially in the installed state with regard to an axis of rotation (8) of a rotor (3) of the gas turbine (1) and, in the circumferential direction, are spaced a distance apart from one another. At least one baffle plate (14) is arranged on the inside (12) and is in contact with the ribs (13). At least one groove (16) is designed in an end face (15) bordering the stator heat shield (6) in the circumferential direction, into which at least one sealing element (18) can be inserted. A plurality of bores (19) are included, each opening at a distance from the groove (16) in the direction of the outside (11) at the inside (12) at one end and on the end face (15) at the other end and arranged a distance apart from one another in the axial direction.

Abstract: A heat accumulation segment for local separation of a flow duct inside a turbo engine, from a stator housing that radially surrounds the flow duct is provided. The heat accumulation segment includes two axially opposed joining contoured elements that are engagable with two components that are axially adjacent along the flow duct. A first one of the two joining contoured elements has a radially oriented recess with a contoured surface against which a securing pin having an external contour complementary to the contoured surface acts radially under force action from a component that adjoins the first joining contoured element. The first joining contoured element has a collar portion having radially upper and lower collar surfaces, and the collar portion is connected within a counter-contoured receiving contoured element in the axially adjacent component by a joining force that acts between the securing pin and the conical contoured surface.

Abstract: A heat shield arrangement for local separation of a flow channel within a turbine engine, with respect to a stator housing radially surrounding the flow channel is provided. The heat shield includes two axially opposite joining contours which are each engageable with two components which are axially adjacent along the flow channel. Each provides a complementary reception contour for the joining contours. At least one of the reception contours has an axial clearance, in which the associated joining contour is axially displaceably mounted. At least one seal is provided between the axially displaceable joining contour and the reception contour. The seal is mounted movably within the reception contour or the joining contour in such a way that the seal is deflectable against a surface region of the reception contour or of the joining contour.

Abstract: A heat accumulation segment for local separation of a flow duct inside a turbo engine, from a stator housing that radially surrounds the flow duct is provided. The heat accumulation segment includes two axially opposed joining contoured elements that are engageable with two components that are axially adjacent along the flow duct. A first one of the two joining contoured elements has a radially oriented recess with a frustoconical contoured surface against which a securing pin having a frustoconical external contour that acts radially under force action from a component that adjoins the first joining contoured element, and the first joining contoured element has a collar portion having a radially upper collar surface and a radially lower collar surface. The collar portion is connected within a counter-contoured receiving contoured element in the axially adjacent component by a joining force that acts between the securing pin and the frustoconical contoured surface.

Abstract: What is described is a turbine guide vane having a guide vane leaf and with a radially outer platform connected to the guide vane leaf and having a platform surface which faces radially away from the guide vane leaf and on which a connecting structure is provided for fastening the guide vane to a supporting structure which has side wall portions which project radially beyond the platform surface and delimit an inner cavity and on which is provided at least in portions a joining contour which can be inserted into a reception shape countercontoured within the supporting structure. At least one seal is provided between the connecting structure and the supporting structure.

Abstract: A gas turbine (1) includes a rotor (2) which has two rotor blade rows (5) with a plurality of rotor blades (6), and also a rotor heat shield (7), which is arranged between them, with a plurality of heat shield elements (12), and with a stator (3) which has a stator blade row (8), with a plurality of stator blades (9), which is arranged axially between the two adjacent rotor blade rows (5). The stator blades (9) have a stator sealing structure (10) radially on the inside. The heat shield elements (12) have a rotor sealing structure (13) radially on the outside which interacts with the stator sealing structure (10) for forming an axial seal (14). Furthermore, a blade radial seal (15) is formed between two adjacent rotor blades (6), and also a heat shield radial seal (16) is formed between two adjacent heat shield elements (12), and in each case separates a gas path (17) from the rotor (2).

Abstract: A seal assembly for a gas turbine is arranged in grooves of a rotor heat shield having several bends. The assembly comprises four seal portions overlapping one another and extending in the axial, radial, and circumferential direction with respect to the turbine rotor. A holding means retains the radial sections of one seal portion allowing a limited movement of said seal portion independent of another seal portion. The independent movement assures contact of the individual seal portions with all mating surfaces of the rotor heat shield and improved sealing function regardless of displacements of the rotor heat shield and tolerances of turbine parts.