Abstract: The present invention relates to a blade (100) or vane for a turbomachine, having at least one impulse element housing (1.1, 1.2) with a first impact cavity (10), in which an impulse element (11) is arranged with play of movement, wherein the impulse element housing has at least one second impact cavity (20), which is in alignment with the first impact cavity in a first matrix direction (A) and in which an impulse element (21) is arranged with play of movement, and has at least one third impact cavity (30), which is in alignment with the first impact cavity in a second matrix direction (B) crosswise to the first matrix direction and in which an impulse element (31) is arranged with play of movement.

Abstract: A turbofan aircraft engine has at least one stage pressure ratio is at least 1.5, and a quotient of the total blade count divided by 110 is less than a difference ([(p1/p2)?1]) of the total pressure ratio minus one, and the total pressure ratio is greater than 4.5, and the turbine has at least two and no more than five turbine stages; and/or a product (An2) of an exit area (AL) of the second turbine and a square of a rotational speed of the second turbine at the design point is at least 4.5·1010 [in2·rpm2], and a blade tip velocity (uTIP) of at least one turbine stage of the second turbine at the design point is at least 400 meters per second. A jet and method are also provided.

Abstract: A turbofan aircraft engine having a primary duct (C), including a combustion chamber (BK), a first turbine (HT) disposed downstream of the combustion chamber, a compressor (HC) disposed upstream of the combustion chamber and coupled (W1) to the first turbine, and a second turbine (L) disposed downstream of the first turbine and coupled (W2) via a speed reduction mechanism (G) to a fan (F) for feeding a secondary duct (B) of the turbofan aircraft engine. A square of a ratio of a maximum blade diameter (DF) of the fan to a maximum blade diameter (DL) of the second turbine is at least 3.5, in particular at least 4.

Abstract: A gas turbine jet engine having a main flow channel and a housing structure which radially surrounds this main flow channel. A housing gas flow flows in the same direction as the main flow in the main flow channel, the housing structure having a flow conducting assembly. The flow of the housing gas flow to and/or along the main flow channel may be adjusted by adjusting a flow conducting sheet.

Abstract: The present invention relates to a gas turbine casing arrangement having a gas turbine casing element (10), a guide-vane ring with an outer ring (20), and a coated ring (30), which lies radially opposite a rotor grid (40) adjacent to the guide-vane ring, whereby an intermediate ring (50) by a downstream front face (51) engages behind an upstream stop (61; 61?) fixed in place on the casing element, and by an upstream front face (52) of a radial flange (53) of the intermediate ring engages behind a downstream stop (22) fixed in place on the outer ring, in order to secure the guide-vane ring axially at the gas turbine casing element in the direction of through-flow.

Abstract: A rotor (10) has a basic rotor body (12) and a plurality of rotating blades (14) mounted on the basic rotor body (12). The rotating blades (14) in this case are mounted rigidly or non-detachably, in particular, cohesively, on the basic rotor body (12). Thus, at least one rotating blade (14) has at least one integral sealing element (24), by means of which a root intermediate space (22) is sealed in the region of a blade root (20) radially underneath a blade platform (18) of the rotating blade (14). In addition, the invention relates to a method for producing a rotor (10), in particular, for an aircraft engine, in which a plurality of rotating blades (14) is mounted on a basic rotor body (12), wherein at least one rotating blade (14) having at least one integral sealing element (24) is mounted on the basic rotor body (12).

Abstract: The present invention relates to a sealing element (100) of an axial turbomachine for the sealing of regions at or in at least two static components of the turbomachine, wherein the components can be moved axially relative to one another. The sealing element (100) has a first radial region (1) with a first stiffness or rigidity and a second radial region (5) with a second stiffness or rigidity, wherein the first stiffness and the second stiffness are different from one another, and the sealing element (100) is configured as segmented in the peripheral direction of the turbomachine.

Abstract: A component system of a turbine engine including a first component segment and a second component segment configurable in a ring segment shape, so that at least one abutment surface of the first component segment and an abutment surface of the second component segment abut against each other; together, the first component segment and the second component segment including at least three overlapping elements for sealing a gap between the abutment surfaces. In the case of mutually abutting abutment surfaces, each overlapping element overlapping radially with the respective other component segment. At least two of the overlapping elements are configured on the first component segment, while at least one of the overlapping elements is configured on the second component segment. In the case of mutually abutting abutment surfaces, the overlapping element of the second component segment is axially configured between the overlapping elements of the first component segment.

Abstract: A sealing device (1) is disclosed for sealing a radially inner gas channel (2) between a guide vane ring (4) and a rotor (6) of a turbomachine, wherein the sealing device (1) has a sealing ring (8) for forming a sealed space (10) with a rear segment, with an inner wall structure (30) oriented in the opposite direction, which are joined to each other via an annular arch (32), wherein the radial flange (18) transitions into the outer wall structure (28) and the cylinder (26) forms the sealing ring (8), wherein the inner wall structure (30) transitions, via an annular web (36), into at least one inner body segment (38, 50), wherein the sealing device (1) has a uniform, preferably relatively reduced wall thickness over its individual segments integrally formed with one another, so that the sealing device (1) is resilient within certain limits.

Abstract: The invention relates to a sealing arrangement 15, a guide vane arrangement, and a turbomachine 11 with such a sealing arrangement 15, wherein the sealing arrangement 15 is designed for a guide vane ring 60 of a turbomachine 11, wherein the sealing arrangement 15 comprises a thin-walled annular structure 80 that is substantially closed on all sides, and wherein the annular structure 80 delimits an annular interior space 105, wherein a hollow cell structure 109, which is designed so as to mechanically support the annular structure 80, is provided in the annular interior space 105.

Abstract: A rotor blade for a turbomachine, including an airfoil (1) for flow deflection, a blade root (2) for attachment to a rotor of the turbomachine, an inner platform (3) between the airfoil and the blade root, and at least one pocket (4, 5) defined by two axially spaced walls (4.1, 4.2; 5.1, 5.2) extending from the side of the platform opposite the airfoil toward the blade root. A first (4.1, 5.1) of these walls has an outer side (4.1A, 5.1A) that faces away from the pocket and, in at least one cross section perpendicular to a radial longitudinal axis of the rotor blade, slopes outwardly.

Abstract: A stator vane segment (100) of a fluid flow machine, having an upstream casing-side attachment (1) and/or a downstream casing-side attachment (3), and a platform (5). The attachments are provided on the platform (5). The platform (5) has a first sealing section (17) and extends in the direction of flow beyond a connection region (15) that connects the rearward attachment (3) to the platform (5). The first sealing section (17) that is located in a section (18) downstream of the connection region (15) has a first slot (19) which is continuous in the circumferential direction. The present invention also relates to a turbine having at least one stator vane segment (100).

Abstract: A turbine casing (100) for a gas turbine, having a casing section (5a, 5b, 5c, 5d, 5e), the casing section (5a, 5b, 5c, 5d, 5e) having a containment section (16) for containing blade fragments expelled radially outward. The casing section (5a, 5b, 5c, 5d, 5e) has a first reinforcing element (1a, 1b, 1c, 1d, 1e) which is joined in face-to-face contact with the radially inner and/or outer surface of the casing section (5a, 5b, 5c, 5d, 5e).

Abstract: The present invention relates to a balancing body (1) for fastening to a ring (4, 5) of a continuous blade arrangement of a compressor or turbine stage of a gas turbine, wherein the balancing body has a first stop (11) for the form-fitting attachment of the balancing body in one peripheral direction (R) to a first axial shoulder (6; 7) of the ring.

Abstract: A turbomachine in the form of a stationary gas turbine or an aircraft engine, respectively a housing structure therefor; the housing structure including an outer housing wall (1) and an inner wall (2) defining the flow channel; and a hollow space (4) being formed between the inner wall and the outer housing wall. The hollow space is separable into at least two regions (5, 6); a movable wire element (slide ring seal) (10, 10?), which is adapted to rest against the contact faces (8, 9), being configured in the hollow space for purposes of the separation.

Abstract: A turbomachine is disclosed. The turbomachine includes a rotor, a plurality of blades arranged side by side on the rotor in a circumferential direction, and a plurality of sealing apparatuses each disposed on a respective blade shaft of the plurality of blades. Each of the plurality of sealing apparatuses includes first and second bottom plates protruding in a respective axial direction on the respective blade shaft and first and second bulkheads which extend in a radial direction along the respective blade shaft and form a double-walled seal in the axial direction. A respective rounded or parabolic transitional cross-section from the bottom plates to the bulkheads is provided and at least one of the bulkheads is offset with respect to a blade root edge of the respective blade shaft in a direction toward a radial longitudinal axis of the respective blade shaft.

Abstract: A turbine blade (100) having an airfoil (1). In its radially outer region, the turbine blade (100) has at least one shroud (200) or at least one shroud segment, as well as at least one cutting tooth (300) which is intended to engage at least once into portions of a turbine casing during use of the turbine blade (100). The cutting tooth (300) is connected to the shroud (200) or the shroud segment.

Abstract: A housing structure for a turbomachine having a stationary blade ring annularly surrounding a flow channel, and having a seal carrier ring which also annularly surrounds the flow channel, the stationary blade ring and the seal carrier ring being connected to each other, and either the stationary blade ring or the seal carrier ring having a one-sidedly open receptacle in the radial direction in relation to the flow channel and a radially projecting engagement area which corresponds to the connected seal carrier ring or stationary blade ring and which is held in a form-fitting manner in the axial direction and in the radial direction in the direction of the closed side of the receptacle but is freely movable with respect to the connected receptacle and the seal carrier ring or stationary blade ring having this receptacle in the radially opposite direction in the direction of the receptacle opening.

Abstract: A component combination for a shell of a flow channel (10) of a turbomachine, in particular for a housing of a gas turbine or an aircraft engine, including at least two components, a first component (1) forming a closed ring. The closed ring is axially and/or radially held in a form-fitting manner against at least one second component (2) in relation to the flow channel and is able to be removed from or placed into this position only by elastically deforming the ring without changing the second component; the present invention also includes a method for manufacturing a component combination of this type and a corresponding turbomachine having a component combination of this type.

Abstract: A housing structure for a turbomachine, the housing structure annularly surrounding, at least partially, a flow channel (14) in which guide vanes and rotor vanes are arranged. The housing structure has an inner wall (2) that delimits the flow channel and an outer wall (1) that is closed off vis-à-vis the environment. A heat-protection plate (3) is arranged on the radial outside of the inner wall, and insulation (4) is arranged between the heat-protection plate and the outer wall. At least one cooling-air channel (6) is formed between the insulation and the radial inside of the outer wall, the cooling-air channel (6) having at least one air-guiding element (5) resting against the insulation.