Abstract: A rotor (10) for a turbomachine, in particular for an aircraft engine, having a rotor base body (12), on which at least one sealing fin (14), which is disposed on a base (16), is provided for cooperating with an associated sealing element (20) of the turbomachine; relative to an axial direction of the rotor (10), the base (16) having a base portion (16a) disposed upstream of the sealing fin (14) and a base portion (16b) disposed downstream thereof, for supporting masks during the coating of sealing fins; the upstream base portion (16a) and the downstream base portion (16b) having different radial distances (A1, A2) to a radially outer sealing tip (18) of the sealing fin (14). Also, a turbomachine having at least one such rotor (10).

Abstract: A gas turbine component for a gas turbine is provided. The gas turbine component has a shell portion having an abradable portion (2) to allow rotor blade ends (1) of a rotor blade array of the gas turbine to rub into the same. The abradable portion adjoins a neighboring portion (3) of the shell portion axially at both sides and is additively manufactured together therewith.

Abstract: A turbomachine modification apparatus includes a machining device, and a machining device mount attached to the machining device. The machining device mount is configured to radially and axially adjust a position of the machining device. A blade mount is configured to be attached to one or more turbomachine blades. The blade mount is attached to the machining device mount.

Abstract: A method of installing a retaining ring assembly in a turbine case is provided. The method may include installing a blade outer air seal into a case before installation of a retaining ring assembly. The method may further include installing a key system on a retaining ring to create the retaining ring assembly. The method may also include installing the retaining ring assembly in a case. The method may further include rotating the retaining ring assembly in the case until key system is aligned with a case slot. The method may also include tightening the key system to the retaining ring and cause key system to engage the case slot.

Abstract: The invention relates to a sealing assembly for a fluid kinetic machine, in particular for an aircraft engine, for sealing a radial gap between a rotor and a stator, including at least one sealing support for retaining and/or fixing at least one sealing element, wherein the sealing support includes a first and a second radial web extending in a radial extension direction as well as an axial web extending in an axial extension direction, firmly bonded to the radial webs, and the radial webs form a receptacle for receiving an element of the stator. Therein, the radial webs are formed as sheet elements formed elongated in radial direction, wherein a radially interior end of the radial webs is respectively firmly bonded, in particular welded, to a radially exterior surface of the axial web.

Abstract: In some examples, a system includes a blade including a blade tip and a blade track or blade shroud segment including a substrate and an abradable coating layer on the substrate. The substrate defines a leading edge and a trailing edge. The abradable coating layer includes a first tapered portion that substantially continuously tapers from a center portion of the substrate toward the leading edge of the substrate, a second tapered portion that substantially continuously tapers from the center portion of the substrate toward the trailing edge of the substrate, and a blade rub portion that extends between the first tapered portion and the second tapered portion. The abradable coating extends from the leading edge to the trailing edge, and the blade tip is configured to contact at least a portion of the blade rub portion upon rotation of the blade.

Abstract: Leakage of a working fluid from a shaft sealing device of a turbomachine which handles a liquid as the working liquid is reduced and also unstable fluid force which acts on a rotor is reduced. For this purpose, the turbomachine has a rotor having an impeller, a bearing which rotatably supports the rotor, a casing which contains the rotor therein and forms a stationary flow passage, and the shaft sealing device which reduces the leakage of the working fluid through a gap part between the casing and the rotor. The working fluid is a liquid. A plurality of axial grooves which are intermittently arranged in axial direction and circumferential direction of the rotor with a land part being interposed are provided in an inner circumferential surface of the shaft sealing device.

Abstract: A labyrinth seal for a gas turbine engine is provided. The labyrinth seal includes a stator and a rotor spaced apart from the stator. The rotor includes a first tooth at a first side having a first radial height, and a second tooth between the first tooth and a second side. The second tooth has a second radial height that is different than the first radial height. The rotor includes a third tooth between the first tooth and the second side having a third radial height that is substantially the same as the first radial height. A first clearance between a first tip of the first tooth and the stator is different than a second clearance between a second tip of the second tooth and the stator, and a third clearance between a third tip of the third tooth and the stator is substantially the same as the first clearance.

Abstract: A turbine engine that includes a stationary assembly, and a rotor assembly configured to rotate relative to the stationary assembly. The rotor assembly includes a plurality of unitary turbine and fan blades. Each unitary turbine and fan blade includes a single turbine airfoil, a single fan airfoil positioned radially outward from the single turbine airfoil, and a midspan shroud segment defined between the single turbine airfoil and the single fan airfoil.

Abstract: A seal in a gas turbine for sealing a radial gap defined between rotating and stationary structure. The rotating structure may include a row of rotor blades. The seal may include a pad attached to the stationary structure. The pad may include an abradable structure. The pad may further include a liner attached to and covering an outer surface of the abradable structure. The stationary structure to which the pad is attached may define an axial section of the outer radial boundary of the annular flowpath, the axial section coinciding axially with an axial position of the row of rotor blades.

Abstract: The present invention relates to a seal arrangement for a turbomachine, in particular a gas turbine, having a plurality of rows, arranged in succession in the axial direction (A), of shells (1-3) connected to one another in the circumferential direction (U), wherein shells adjacent in the axial direction have cross sections opened counter to a throughflow direction (A) and/or a thread axis inclined counter to the throughflow direction.

Abstract: One exemplary embodiment according to this disclosure relates to a system includes a blade outer air seal (BOAS), and a meter plate. A portion of the meter plate is provided radially outward of a radially outermost surface of the BOAS.

Abstract: An air cooled component for a turbine stage of a gas turbine engine, comprising: a main body having radially inner main gas path wall and a cooling chamber, the main gas path wall separating the main gas path of the turbine stage and the cooling chamber; at least one flange extending from the main body; a cooling cavity enclosed within the flange; and, an inlet conduit extending between and fluidically connecting the cavity and cooling chamber.

Abstract: A blade outer air seal (BOAS) for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion. At least one cooling fin is disposed on the radially outer face between the leading edge portion and the trailing edge portion.

Abstract: Disclosed is a turbomachine having an annular flow duct and a housing structure surrounding the flow duct and a multiplicity of guide vanes and rotor blades which are arranged in the flow duct. The rotor blades are rotatably accommodated in the housing structure whereas the guide vanes are fixed in the housing structure, a plurality of guide vanes forming an annular guide vane ring. The housing structure has a seal in the region of the radially inner flow duct boundary in order to prevent hot gas escaping from the flow duct, which seal is arranged on guide vane roots of the guide vanes of the guide vane ring via a seal support and forms a seal against a rotatable seal surface, the seal support being formed from an intermetallic material, in particular a TiAl material.

Abstract: A turbofan engine has a fan portion in fluid communication with a core stream and a bypass stream of air separated by splitters disposed both upstream and downstream of the fan portion. A fluid passage is defined between the splitters. The turbofan engine has a plurality of surface interruptions on a core side surface of the upstream splitter and a plurality of vortices originating from the surface interruptions, thereby restricting the migration of the core stream into the bypass stream through the fluid passage.

Abstract: A rotating machine includes a casing (10) having a cavity (12) that a tip of a rotor blade enters, a plurality of sealing fins (17) extending from an inner circumferential surface of the cavity (12) of the casing (10) toward the tip of the rotor blade (50) and configured to seal a space between the casing (10) and the rotor blade (50), and swirl breakers (2) disposed between the plurality of sealing fins, extending from the inner circumferential surface of the cavity (12) of the casing (10) inward in the radial direction, and having swirl flow collision surface (3) with which a swirl flow collides and swirl flow transmission parts (n) formed at at least parts of the swirl flow collision surfaces (3) and through which the swirl flow passes in a circumferential direction.

Abstract: A seal arrangement for effecting a seal between first and second components arranged for relative rotation about a common axis. Seal arrangement includes a first seal surface associated with the first component, and a second seal surface associated with the second component. First and second seal surfaces are presented for air-riding interaction with one another during relative rotation between components such that, in use, a pressure drop is maintained across seal arrangement. First seal surface incorporates a plurality of spaced-apart hydrodynamic lift-generating features and plurality of ventilation recesses, each ventilation recess being interposed between respective pair of lift-generating features, and wherein each ventilation recess has depth measured normal to first seal surface which is greater than maximum dimension of each lift-generating feature measured normal to first seal surface. Seal arrangement may be embodied in compliant filament seal such as a leaf seal or a brush seal.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines vary localized porosity or abradability through use of holes or dimple depressions of desired polygonal profiles that are formed into the surface of otherwise monolithic abradable surfaces or rib structures. Abradable porosity within a rib is varied locally by changing any one or more of hole/depression depth, diameter, array pitch density, and/or volume. In various embodiments, localized porosity increases and corresponding abradability increases axially from the upstream or forward axial end of the abradable surface to the downstream or aft end of the surface. In this way, the forward axial end of the abradable surface has less porosity to counter hot working gas erosion of the surface, while the more aft portions of the abradable surface accommodate blade cutting and incursion with lower likelihood of blade tip wear.

Abstract: The invention relates to a seal (10) for sealing a gap between a stationary component and a moving component, in particular for sealing a radial gap between a rotor and a stator of a turbomachine, comprising at least one sealing segment (12) with an edge zone (14) facing the gap, whereby the seal (10) is produced layer-by-layer by a free-forming method, in particular a generative or additive method. A plurality of pre-defined weak regions (16) is formed in the edge zone (14) of the sealing segment (12). In addition, the invention relates to a method for producing a seal (10) as well as a turbomachine.

Abstract: A corrosion-resistant abradable covering to adhesively bond to a wear portion of a structural part, the covering including at least a polyester-based matrix and an aluminum alloy filler that are molded together.

Abstract: A brush seal system for sealing a clearance between components of a turbo engine that are movable in relation to one another, in particular of a thermal gas turbine, is disclosed. The brush seal system includes a brush seal housing, which accommodates at least one brush head of a brush seal, where the brush seal housing includes a first component having a cover plate section and a second component having a support plate section. The first component includes an axial flange forming a fish mouth seal on an end opposite the cover plate section. A thermal gas turbine having such a brush seal system is also disclosed.

Abstract: The invention pertains to a seal system for a passage between a turbine stator and a turbine rotor, including: a first arm extending radially outwards from the turbine rotor and toward the first seal arranged on the stator, and terminating short of the first seal thereby creating a first gap between the first seal and the first arm. The seal system further includes a second seal arranged on the turbine stator, and a second arm extending axially from the turbine rotor towards the second seal base, and terminating short of the second seal thereby creating a second gap between the second seal and the second arm. The invention further refers to a gas turbine including such a seal system.

Abstract: A blade outer air seal member includes a distinct body that has two circumferential sides, a leading edge and a trailing edge, and a gas path side and a radially outer side opposite the gas path side. A ceramic coating is initially disposed on a portion of the gas path side. The ceramic coating includes a forward coating portion and an aft coating portion. The gas path side has a bare area axially separating the forward coating portion and the aft coating portion. The bare area excludes any of the ceramic coating. One or more cooling passages have an outlet that opens at the bare area. The cooling passage extends in the body in an axial direction under the ceramic coating.

Abstract: Systems are provided for a reinforcement element coupled to a sheet metal turbine housing that imparts desirable thermal-protective and structurally strengthening characteristics to the housing layers. In one example, a system may include a turbine comprising a housing surrounding a turbine rotor, the housing having an outer layer surrounding an inner layer at a distance to form an intermediate space between the inner and outer layers. Moreover, disposed in the intermediate space is a reinforcement element coupled to the inner and outer layers, providing strength and consistent rigidity without a significant increase in weight to the housing.

Abstract: A turbofan engine has a fan portion in fluid communication with a core stream and a bypass stream of air separated by splitters disposed both upstream and downstream of the fan portion. A fluid passage is defined between the splitters. The turbofan engine has a plurality of high pressure fluid jets originating from the low pressure side of the fan blades, the jets restricting the migration of the core stream into the bypass stream through the fluid passage.

Abstract: A damper system may include an inner damper having a first annular geometry and a textured surface. An outer damper may also have an annular geometry and be disposed about the inner damper. The outer damper may have a textured surface. The textured surfaces of the inner damper and outer damper may define a cavity that is configured to contain a viscous fluid.

Abstract: A fish mouth seal carrier for a guide vane arrangement of a gas turbine comprises a first half-shell element and a second half-shell element bonded to it, which together form a box profile with two axial arms and two radial arms. A sealing element is arranged on one of the axial arms of the box profile. At least one of the two half-shell elements has an integrally formed axial flange for forming, with a guide vane platform, a fish mouth seal accommodating an axial flange of an adjoining moving blade radially between the guide vane platform and the axial flange of the seal carrier.

Abstract: A device and method for precise measurement of erosion of compressors, without removing an engine, and with easy positioning. A centrifugal compressor of a gas turbine with a radial air inlet includes an impeller including blades and a casing for an air stream to flow in the blades of the impeller. The casing, covered with an abradable coating, includes an annular elbow zone in a substantially median part. Marking depressions of predetermined depths, preferably in groups, are machined in the abradable coating of the zone. Examinations by endoscopy are successively performed to provide an image signal of the markers. Processing the endoscopic signal supplies a number of remaining markers and a criterion for decision on removing the engine is applied thereto. Erosion occurs in the elbow of the casing and evolution thereof may enable monitoring of erosion of other components of the compressor, in particular the blades of the impeller.

Abstract: A seal arrangement for a gas turbine engine according to an example of the present disclosure includes, among other things, a component including a body having a cold side surface adjacent to a mate face, and a seal member including a leading edge region and a trailing edge region spaced by sidewalls. The seal member defines one or more grooves. The one or more grooves abut the cold side surface to define one or more cooling passages, with at least one of the one or more cooling passages having a flared inlet defined by a corresponding one of the one or more grooves.

Abstract: A gas turbine includes a compressor, a combustor, and a turbine. The turbine includes a honeycomb seal disposed so as to be secured to a casing side in a clearance between the casing and turbine blades rotating around a rotating shaft and a seal fin that is provided on an end face of each of the turbine blades facing the honeycomb seal. The seal fin extends in a direction perpendicular to the rotating shaft. The honeycomb seal is formed by a plurality of corrugated sheet metals overlapped with each other at walls of nodes thereof and the walls of the nodes are blazed with each other. Each of the corrugated sheet metals has trapezoids formed in alternating fashion. A longer direction of each wall of the nodes of the honeycomb seal is angled with respect to the rotational direction of the turbine blades.

Abstract: A turbine nozzle or a compressor diffuser includes sectors made of CMC material, each having an inner platform, an outer platform, and airfoils. An abradable material support ring is made up of sectors, each presenting upstream and downstream attachment tabs. Each inner platform presents on the inside an upstream hook and a downstream hook, and the abradable material support ring is supported by the nozzle or the diffuser by mutually engaging without fastening together firstly the end portions of the upstream attachment tabs and the upstream hooks and secondly the end portions of the downstream attachment tabs and the downstream hooks.

Abstract: A blade outer air seal for use in a gas turbine engine having an axis of rotation includes a main body having a mating face configured to face, be positioned radially outward from, and be positioned adjacent to a rotor blade of the gas turbine engine. The blade outer air seal also includes an axial member extending aft from the main body, having a first radial face configured to face a second radial face of an outer diameter platform of a stator of the gas turbine engine, and having a first abradable material coupled to the first radial face.

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: An abradable turbine component, a method of creating a turbine component with an abradable mesh structure, and a gas turbine engine are provided. The abradable turbine component includes a turbine component surface for coupling to a turbine casing, and a deposited abradable mesh structure coupled to the turbine component surface. The abradable mesh structure includes interlacing strands of material, each strand including a height relative to the turbine component surface. At least two of the plurality of interlacing strands include a height different from each other. The method includes applying a bond coat layer followed by a thermal barrier coating layer. An abradable mesh structure is deposited on top of thermal barrier coating wherein the abradable mesh structure includes interlacing strands of material wherein at least two of the interlacing strands include a height different from each other. A gas turbine engine including the abradable turbine component is also provided.

Abstract: An inner ring of a fluid flow machine includes fixing rings and seal carriers split into at least two fixing ring segments and into at least two seal carrier segments, respectively. The fixing ring segments and the seal carrier segments are arranged with the respective end faces facing each other. A first fixing ring segment has a first ring segment shoulder and a second fixing ring segment has an offset second ring segment shoulder. A first seal carrier segment has a first carrier segment shoulder, and a second seal carrier segment has a second carrier segment shoulder, which is offset with respect to the first carrier segment shoulder. In a separation plane of the inner ring, the first carrier segment shoulder abuts the second ring segment shoulder at the end face. Also in the separation plane, the second carrier segment shoulder abuts the first ring segment shoulder at the end face.

Abstract: A blade outer air seal (BOAS) for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion. At least one cooling fin is disposed on the radially outer face between the leading edge portion and the trailing edge portion.

Abstract: A bearing for a gas turbine engine may comprise a squeeze film damper. A bearing housing may comprise a roughened surface. The roughened surface may increase friction between the bearing housing and the oil in the squeeze film damper. The increased friction may increase the damping force in the squeeze film damper. The increased damping force may prevent oil inertia, and prevent resonant vibrations.

Abstract: A seal for a turbo machine is provided. The seal includes a plurality of honeycomb cells, at least one circumferential groove within the plurality of honeycomb cells, and at least one swirl brake within the at least one groove.

Abstract: A seal for a gas turbine engine is shown. The engine may have a first vane for imparting a first twist to air passing thereby and/or a second vane for imparting a second, different twist to air passing thereby, each of the first and second vanes having a mounting area for mounting the seal thereto. The seal includes a top, a bottom, a left side, a right side a back, and a front. The back is parallel to the front and the left side is parallel to the right side such that a non-square shape is formed by the intersection of the back, the left side, the front and the right side so that the seal may mate with the mounting area of the first vane and, if the seal is inverted, may mate with the mounting area of the second vane.

Abstract: A seal for a turbine engine is provided that includes a seal body disposed at a base of a turbine engine blade and a wing portion extending axially from the seal body. The wing portion has a first portion substantially parallel with a centerline of the engine and an angled upturn portion. The first portion extends between the seal body and the angled upturn portion. An angle between the engine centerline and the angled upturn portion is between about 0 degrees and about 90 degrees.

Abstract: A shroud for a turbine and a method of assembling is provided. The turbine includes a housing, a rotatable shaft and a bucket extending outward therefrom. The shroud includes an alignment member which is coupled to the housing, wherein the alignment member includes a first end, a second end and a body extending between the first and second ends. The second end includes an arcuate portion to facilitate fluid flow downstream from the bucket. The shroud further includes a seal coupled to the body to facilitate sealing a gap defined between the bucket and the body.

Abstract: A knife edge seal assembly includes at least one knife edge and at least one honeycomb ring seal. The at least one knife edge has material removed to a first knife edge radial height and the least one honeycomb ring seal has material removed to a finished radial thickness for sealable engagement with the at least one knife edge. The finished radial thickness of the replacement honeycomb ring seal establishes a first radial gap between an inner annular surface of the replacement honeycomb ring seal and the knife edge before the knife edge seal assembly begins rotating in the gas turbine.

Abstract: According to one embodiment of the disclosure, an annular seal generally includes a hollow member having two ends and an inner surface. The hollow member has a number of depressions formed in and extending around its inner surface. Each of the depressions having a depth that is a function of its distance from one of the two ends.

Abstract: A MCrAlY alloy, methods to produce a MCrAlY layer and a honeycomb seal are provided. The MCrAlY alloy includes chromium, aluminum, yttrium and iron and optionally titanium, hafnium or silicon. The honeycomb seal includes a substrate, honeycomb cells and a protective coating on side walls of the honeycomb cells or a diffusion area inside side walls of the honeycomb cells, the protective coating or the diffusion area including the MCrAlY alloy.

Abstract: In one aspect, the present subject matter discloses a seal including a substrate material. The substrate material has a first portion having a first abradability and a second portion having a second abradability, the first abradability being different from the second abradability.

Abstract: A seal assembly for a rotary machine is provided. The seal assembly includes multiple sealing device segments disposed intermediate to a stationary housing and a rotor. Each of the segments includes at least one plate and a sealing element. The seal assembly also includes multiple inter-segment gaps formed between the multiple sealing device segments. Further, for each pair of adjacent sealing device segments, at least one pair of adjacent plates includes one plate overlapping another plate at an inter-segment gap.

Abstract: A seal structure includes a thick abradable film, thereby improving sealing performance. A seal structure (1) is provided with at least one fin (5) that projects in a ring-shape from a circumferential surface of a rotating shaft (3) and that is provided along an axial direction (L) and a seal member (9) having a ring-shaped sealing surface (19) opposed to the fin (5), and on which an abradable layer (25) is formed by thermally spraying abradable material onto the sealing surface (19), the seal structure (1) is provided with tapered portions (21) inclined in a radial direction (K) at end portions of the sealing surface (19) in the axial direction (L).

Abstract: A honeycomb seal is provided. The honeycomb seal is in one part with a substrate so that there is no brazing is needed. Furthermore, a better corrosion resistance is gained by the substrate having a coating comprising iron, chromium, aluminum and/or yttrium, especially applied by vapour deposition. A method for producing a honeycomb seal is also provided.

Abstract: According to one embodiment of the disclosure, an annular seal generally includes a hollow member having two ends and an inner surface. The hollow member has a number of depressions formed in and extending around its inner surface. Each of the depressions having a depth that is a function of its distance from one of the two ends.