Abstract: A turbine rotor blade with a tip region cooling and sealing circuit that includes a squealer tip with a continuous tip rail, the pressure side tip rail being offset from the pressure side wall, the pressure ad suction side tip rails both include concave shaped deflector surfaces on the forward side walls and a row of air jet passages opening onto a top surface of the tip rails, and with radial near wall cooling channels formed within the pressure side and suction side walls to provide near wall cooling, and where the pressure side radial cooling channels supplies the cooling air to the tip rail deflectors to form a vortex flow on the forward side wall of the tip rails, and the suction side radial cooling channels supplies cooling air to the tip rail air jet passages to discharge cooling air to block the on-coming leakage flow across the blade tip.

Abstract: A method of detecting rubs during operation of a turbomachine comprising at least one rotating object having a tip and a shelf is provided. The method includes generating signals representative of a sensed parameter and processing the signals to generate height versus time data for the tip. The height of the tip corresponds to the distance between the tip and the shelf. The method further includes monitoring the height versus time data, in order to determine whether a change in the height data exceeds a threshold value, and detecting a rub of the rotating object(s) on a second object, when the change in the height data exceeds the threshold value. A rub detection system for a turbomachine and a turbine engine system with rub detection are also provided.

Abstract: A multiple of circumferential grooves within said arcuate engine casing and an abradable material located radial inboard of the multiple of circumferential grooves.

Abstract: External shroud for a turbomachine turbine wheel, this shroud being formed of ring sectors (120) positioned circumferentially end to end, and each comprising a circumferentially directed plate (136) bearing a block (138) of abradable material, the plate and the block of each ring sector being offset one with respect to the other in the circumferential direction so that a circumferential end part of the plate of each ring sector overlaps a corresponding end part of the block of an adjacent ring sector.

Abstract: In one embodiment, an abradable seal includes a soft material that may be worn away and a hard material that may provide mechanical strength for the abradable seal. The soft material and the hard material may include different compositions of a base material.

Abstract: A turbine blade is provided comprising: an airfoil including an airfoil outer wall extending radially outwardly from a blade root, a squealer tip section located at an end of the airfoil distal from the root, and cooling structure. The squealer tip section comprises a blade tip surface including pressure and suction edges joined together at chordally spaced-apart leading and trailing edges of the airfoil, and a squealer tip rail. At least a substantial portion of the squealer tip rail is located near the blade tip surface suction edge. The cooling structure directs cooling fluid toward the squealer tip rail to effect impingement cooling of the rail after the cooling fluid has convectively cooled at least a portion of the airfoil outer wall. Cooling fluid is also deflected by the squealer tip rail so as to yield a very small effective flow area above the squealer tip section through which hot working gases may flow.

Abstract: A method for assembling a rotor assembly and a rotor assembly are provided. The method comprises providing a rotor blade including a first sidewall, a second sidewall, where the first and second sidewalls are connected at a leading edge and a trailing edge and extend in span from a root portion to a tip portion, removing blade material from the tip portion to form a tip portion rake angle that enables the tip portion to extend obliquely between the first and second sidewalls, and coupling the rotor blade to a shaft such that during tip rubs the tip portion rake angle facilitates reducing radial loading induced to the blade during tip rubs.

Abstract: A gas turbine shroud includes a ceramic abradable coating superior in abradable property and durability. The gas turbine ceramic abradable coating of the present invention is configured by an abradable metal layer and a porous ceramic abradable layer (hardness RC15Y: 80±3), the porous ceramic abradable layer is provided with slit grooves by machining work, and a slit groove width is 0.5 to 5 mm. Thereby, the abradable property, and durability against a thermal cycle and high-temperature oxidation are improved.

Abstract: A turbine blade growth pocket provides a feature to measure blade growth due to engine operation while maintaining dynamic characteristics of the turbine blade within acceptable limits by providing a variable radius fillet between the pocket and a side rail of the pocket.

Abstract: An abrasive coating on a rotor shaft interacts with cantilevered vanes to form an abradable air seal in a turbine engine. The abrasive coating includes a metal bond coat, and an abrasive layer containing a plurality of abrasive CBN grit particles in a ceramic matrix.

Abstract: A blade tip of a turbine rotor blade for a gas turbine engine, the turbine rotor blade including an airfoil and a root portion for mounting the airfoil along a radial axis to a rotor disk inboard of a turbine shroud, a pressure sidewall and a suction sidewall that join together at a leading edge and a trailing edge, the pressure sidewall and suction sidewall extending from the root portion to the blade tip, and a squealer tip cavity formed at the blade tip, the blade tip comprising: a trailing edge trench originating at the squealer tip cavity, wherein the trailing edge trench generally extends toward the trailing edge of the blade tip.

Abstract: A turbine blade is provided comprising: an airfoil including an airfoil outer wall extending radially outwardly from a blade root, a squealer tip section located at an end of the airfoil distal from the root, and cooling structure. The squealer tip section comprises a blade tip surface including pressure and suction edges joined together at chordally spaced-apart leading and trailing edges of the airfoil, and a squealer tip rail. At least a substantial portion of the squealer tip rail is located near the blade tip surface suction edge. The cooling structure directs cooling fluid toward the squealer tip rail to effect impingement cooling of the rail after the cooling fluid has convectively cooled at least a portion of the airfoil outer wall. Cooling fluid is also deflected by the squealer tip rail so as to yield a very small effective flow area above the squealer tip section through which hot working gases may flow.

Abstract: The present application provides an abradable bucket shroud for use with a bucket tip so as to limit a leakage flow therethrough and reduce heat loads thereon. The abradable bucket shroud may include a base and a number of ridges positioned thereon. The ridges may be made from an abradable material. The ridges may form a pattern. The ridges may have a number of curves with at least a first curve and a second curve and with the second curve having a reverse camber shape.

Abstract: An abradable coating for interaction with tips of airfoils (such as vanes or blades of a compressor of a gas turbine engine) includes a metal bond coat thermal, a ceramic layer, and an abradable layer. The ceramic layer on the metal bond coat provides insulation and acts as a fuse that is adapted to spall off upon high rub interaction. The abradable coating on the ceramic layer contacts the tips of the airfoils during operation of the compressor. The abradable coating is sufficiently abradable to roundup the coating by contact with the airfoil tips.

Abstract: A seal structure for a steam turbine is provided, which is capable of suppressing transfer of heat generated by a friction between a rotating portion and a fixed portion to the rotating portion and of suppressing an increase in the temperature of the rotating portion. A labyrinth seal device serves to suppress the amount of steam leaking from a clearance present between each stator blade (fixed portion) of the steam turbine and a rotor (rotating portion). The seal structure for the steam turbine is formed to ensure that permeable spacers made of a permeable metal are provided on the rotor and a seal stationary body and face respective seal fins provided in the labyrinth seal device.

Abstract: The invention relates to a turbomachine thermomechanical part forming a body of revolution about a longitudinal axis, and including at least one abradable ring for a labyrinth seal. In characteristic manner, the abradable ring is made up of angular sectors that present different stiffnesses between adjacent pairs of sectors. The invention is applicable to a compressor, a turbine, or to a rotor and stator assembly.

Abstract: A steam turbine has: a casing; a rotor rotatably arranged in the casing; a nozzle diaphragm concentrically arranged with respect to the rotor, the nozzle diaphragm being engaged with the casing; moving blades arranged in circumferential direction on outer circumference of the rotor at positions adjacent to the nozzle diaphragm; seal strips circumferentially extending on tips of the moving blades, the seal strips protruding in radial outward direction; and an abradable structure rigidly connected to the nozzle diaphragm. The abradable structure faces the seal strips in radial direction at a facing surface, and has an abradable part made of an abradable material arranged at the facing surface.

Abstract: A blade outer air seal for a gas turbine has an abradable coating with a substantially uniform thickness during service use. Prior to being placed in service, the abradable coating has a central portion with a greater thickness than leading and trailing portions. When placed in service, the blade outer air seal is positioned so that the central portion of the abradable coating with the greater thickness is located in a blade rub path of the turbine blades. As a result of abrasion in the blade rub path, the thickness of the central portion of the abradable coating is reduced so that it becomes substantially similar to the thickness of the leading and trailing portions.

Abstract: An example turbine blade includes a blade having an airfoil profile extending radially toward a blade tip. A shelf is established in the blade tip. A sealing portion of the blade tip extends radially past a floor of the shelf. The sealing portion extends from a blade tip leading edge to a blade tip trailing edge. A groove is established in the blade tip. The groove extends from adjacent the shelf to adjacent the blade tip trailing edge. The groove is configured to communicate a fluid from a position adjacent the shelf to a position adjacent the blade tip trailing edge.

Abstract: A turbomachine includes a compressor provided with a casing having an inner surface defining a flow path. At least one blade is rotatable within the flow path. The at least one blade includes a tip portion that is spaced from the inner surface. At least one patch ring segment is mounted to the inner surface of the casing. The at least one patch ring segment includes a body having a first end that extends to a second end through a surface portion. A first wall member extends from the first end and a second wall member extends from the second end. The first and second wall members include in-turned sections that define a C-shaped cross-section of the at least one patch ring segment.

Abstract: Disclosed is an annular flow duct for a turbomachine, the flow duct arranged concentrically about a machine axis running in the axial direction and defined by a boundary wall of circular cross section for directing a main flow, wherein the boundary wall has a plurality of return flow passages distributed over its circumference and through which in each case a partial flow is detached from the main flow at a bleed position and returned to the main flow at a feed position situated upstream of the bleed position, and having aerofoils, arranged radically in the flow duct, of a blade ring, the aerofoil tips of which lie opposite the boundary wall, with a gap formed in each case, wherein the aerofoils are movable in a predetermined rotation direction along the circumference of the boundary wall. Furthermore, a compressor which is insensitive to pumping and flow separations, as viewed in rotation direction, the bleed position of each return flow passage lies upstream of the corresponding feed position.

Abstract: A turbofan engine comprising an annular inner wall surrounding tips of the fan blades, a layer of insulating material surrounding the inner wall, and an outer casing including an annular outer wall surrounding the insulating material and concentric to the inner wall and at least two annular rub elements extending radially inwardly from the outer wall through only part of a radial thickness of the layer of insulating material, at least two of the rub elements being in axial alignment with the blade tips at every point around a circumference of the fan, each rub element having a radially inner end spaced apart from the inner wall and made of a material harder than that of the blades, and a containment fabric layer wrapped around a support structure of the outer wall.

Abstract: A ring structure with a metal design for a moving blade of axially flowed through compressor and turbine stages, particularly in gas turbine engines is provided. The ring structure includes a circular ring-shaped outer wall, a circular ring-shaped inner wall, which is located at a short radial distance from the moving blade tips, and includes a joining structure provided in the form of hollow chamber structure between the outer wall and the inner wall. The outer wall is provided in the form of a closed, mechanically stable housing wall of the compressor or turbine stage, and the joining structure provided in the form of hollow chamber structure is joined in a fixed manner to the outer wall and to the inner wall, for example, by soldering, diffusion welding or other joining techniques. The inner wall is provided in the form of a closed, mechanically stable structure, which serves as a run-in lining for the moving blade tips and which is made of a metal woven fabric and/or of a metal felt.

Abstract: Systems and methods involving air seals are provided. In this regard, a representative air seal for a gas turbine engine comprises a layer of ceramic exhibiting patterned surface features, wherein a substrate supporting the layer of ceramic lacks the patterned surface features.

Abstract: A support apparatus for a gas turbine shroud is disclosed. The apparatus includes an outer shroud block having a coupling connectable to a casing of the gas turbine and a shroud component having a forward flange and an aft flange. The shroud component is attached to the outer shroud block via the forward flange and the aft flange. The apparatus further includes a damper disposed between the outer shroud block and the shroud component and a biasing element disposed within the outer shroud block. A translational degree of freedom between the damper and the outer shroud block defines a direction of motion of the damper. The biasing element is in operable connection between the outer shroud block and the shroud component via the damper, a bias force of the biasing element directed along the direction of motion of the damper.

Abstract: An abradable seal for a gas turbine engine includes a metal alloy and a plurality of pores in the metal alloy. The plurality of pores have a diameter of approximately 1 to 10 microns.

Abstract: An air seal for use with rotating parts in a gas turbine engine has a matrix of agglomerated fine hBN (hexagonal boron nitride) powder, the particles of which having a first dimension, and of a fine metallic alloy powder, the particles of which having a second dimension. An hBN (hexagonal boron nitride) powder, the particles of which have a third dimension that is greater than the first dimension, is mixed with the matrix.

Abstract: A gas turbine fan blade containment assembly includes a fan case having an inner surface surrounding a jet engine fan and an outer surface. Mounted on the inner surface and across a blade containing region of the fan case is a load spreader layer for initially receiving a point load from a fan blade release (a “blade-out event”). A band layer is mounted to an outer surface of the fan case for carrying at least a portion of a hoop tensile load on the fan case resulting from the blade-out event, and separator film layer is mounted between the outer surface of the fan case and the band layer to retard the formation of stress concentrations in the band layer. In one embodiment, the load spreader layer includes a plurality of circumferentially-arrayed load spreader layer segments.

Abstract: A method for controlling blade tip clearance within a gas turbine engine includes a compressor having at least one first rotor assembly, wherein the blade tips in each of at least one first rotor assembly has a mating geometry with at least one of at least one blade seal surfaces, and a clearance distance extending between the blade tips and the blade seal surfaces; a turbine having at least one second rotor assembly, wherein the blade tips in each of at least one second rotor assembly has a mating geometry with at least one of at least one blade seal surfaces, and a clearance distance extending between the blade tips and the blade seal surfaces; providing an actuator selectively operable to move at least one rotor assembly relative to the shroud; providing an electronic engine controller having a control logic for operating the actuator; and moving at least one first rotor assemblies and at least one second rotor assembly relative to the shroud using the actuator at a response rate according to the control l

Abstract: A blade tip shroud segment for a gas turbine engine in which the shroud segment is cooled and sealed using a combination of vortex cooling air flow and cooling air discharge film cooling. The shroud segment includes a row of vortex chambers formed within the shroud segment and a row of circumferential circular grooves formed on the hot gas flow surface of the shroud segment, where each groove is connected to a vortex chamber through a slot. Cooling air is to the vortex chambers from individual cooling air feed slots or through inter-linking channels connecting adjacent vortex chambers.

Abstract: There is provided a gas turbine shroud including a ceramic abradable coating used as a gap adjusting component that can reduce a fluid leakage from a gap and increase turbine efficiency. A gas turbine ceramic abradable coating includes a bond layer, a thermal barrier ceramic layer, and a porous ceramic abradable layer (hardness may be RC15Y: 80±3). A slit groove is provided in the porous ceramic abradable layer by machining. A width of a rectangular section of the ceramic abradable layer divided by the slit groove may be set to a range of 2 to 7 mm.

Abstract: The invention relates to gas turbine engine seal systems having a rotating member with an abrasive tip surface disposed in rub relationship to a stationary seal member with an abradable surface. The abrasive tip surface is coated with a metallic alloy matrix having ceramic abrasive particles embedded in and projecting from the matrix. The abradable seal surface is coated with a ceramic coating. The rub relationship affords a tight operating clearance between the rotating member and the stationary seal member, thereby improving engine efficiency, reducing fuel consumption and minimizing overhaul downtime.

Abstract: A method for fabricating a fan assembly is provided, wherein the method includes providing a rotor having a plurality of rotor blades. At least one rotor blade is fabricated from a frangible material. The method also includes coupling a casing substantially circumferentially about the rotor, and positioning the casing a distance from the plurality of rotor blades. The distance is selected to facilitate increasing an amount of initial resistance created between the casing and the plurality of rotor blades after a bladeout.

Abstract: A method of attaching a labyrinth seal support to a turbine engine shaft is disclosed. The method includes inserting axial bolts to secure the labyrinth seal support, the shaft and a device for driving the turbine module with the shaft, and screwing tightening screws in extracting holes, for extracting the drive device with respect to the shaft, provided on the shaft to tighten the labyrinth seal support onto the shaft.

Abstract: A seal (48) between first and second relatively rotatable members (26,28). The seal comprises an abradable material on at least one of the relatively rotatable members (26,28). The abradable material comprises an open cell metallic foam (60) and polymeric films (64) closing the cells (62) of the metallic foam (60). The seal (48) is used between compressor rotor blades (26) and a stator casing (28) of a gas turbine engine (10). The seal (48) combines the sealing properties of a closed cell metallic foam with the abradability and manufacturing control of open cell metallic foam.

Abstract: Aspects of the invention are directed to a ring seal segment having a base portion and a frame portion. The base portion and the frame portion are a unitary structure. The frame portion is made of four unitary side walls: a forward wall, an aft wall and a pair of transverse side walls. The side walls can be arranged in a rectangular configuration. At least a part of the base portion can be coated with a thermal insulating material. A ring seal segment according to aspects of the invention is well suited to withstand the expected operational loads in the turbine section. The transverse side walls can provide bending strength to the base and can strengthen the forward and aft side walls. Transition regions between the side walls and the base portion are strengthened by three perpendicular planes that cooperate to provide structural strength to the ring seal segment.

Abstract: An abradable coating system (2-8) for a turbine or a compressor is proposed has at least one layer of woven or non-woven tissue (15) or foam made of: ceramic, glass, glass-ceramic, ceramic-metal composite and combinations thereof. A free-standing preformed element (18) has at least one layer of such a woven or non-woven tissue or foam is filled with fillers and/or binder. A method prepares such a system. Finally, a method for the repair of an abradable coating portion of a turbine or the compressor component (9) includes in a first optional step that the damaged portion is at least partly removed and/or cleaned and/or surface treated, and in a second step at least one free-standing preformed abradable coating element (18) or a complete abradable coating system is attached to the component (9), preferably by using a matrix, wherein for example only a center part (2) or a fraction of a center part (2) is replaced.

Abstract: A turbomachine includes a housing having an inner surface, a compressor, a turbine and a rotary member including a plurality of blade members configured as part of one of the compressor and the turbine. Each of the plurality of blade members includes a base portion and a tip portion. The turbomachine also includes a honeycomb seal member mounted to the inner surface of the housing adjacent the rotary member. The honeycomb seal member includes a contoured surface having formed therein a deformation zone. The deformation zone includes an inlet zone and an outlet zone. The inlet zone is spaced a first distance from the tip portion of each of the plurality of blade members and the outlet zone is spaced a second distance from the tip portion of each of the plurality of blade members. The second distance being substantially equal to or less than the first distance.

Abstract: A turbine blade for use in an industrial gas turbine engine, the blade having a squealer pocket with a plurality of discrete curved cooling channels to cool the blade tip and to reduce the hot gas flow leakage across the tip. The curved cooling channels include a side wall cooling channel, a tip rail crown cooling channel and an inner tip rail wall cooling channel all discharging cooling air from a cooling supply channel within the airfoil. Both the pressure side and suction side tip rails include this arrangement of cooling channels.

Abstract: A ceramic hybrid structure (207, 502, 602, 608) that includes a wavy ceramic matrix composite (CMC) wall (214, 532, 603, 609) bonded with a ceramic insulating layer (230, 538, 604, 610) having a distal surface (242) that may define a hot gas passage (250, 550, 650) or otherwise be in proximity to a source of elevated temperature. In various embodiments, the waves (216, 537, 637) of the CMC wall (214, 532, 603, 609) may conform to the following parameters: a thickness (222) between 1 and 10 millimeters; an amplitude (224) between one and 2.5 times the thickness; and a period (226) between one and 20 times the amplitude. The uninsulated backside surface (218) of the CMC wall (214) provides a desired stiffness and strength and enhanced cooling surface area. In various embodiments the amplitude (224), excluding the thickness (222), may be at least 2 mm.

Abstract: A turbine engine ring seal for sealing gaps between turbine engine outer seal segments and turbine blade tips. The turbine engine ring segment may have an inner radial surface that defines a portion of a gap gas flow path where the inner radial surface may be formed of an abradable ceramic coating and includes a plurality of gas flow protrusions that are oriented transverse to the gap gas flow path. The gas flow protrusions may induce vortices in the gas flow in the gap gas flow path. Additionally, the gas flow protrusions may be series of peaks and depressions between two adjacent peaks, where the depressions have an approximate semicircular shape. The distance between two adjacent peaks may be equal or greater than a width of the depression and the height of a single peak may be six percent or greater than the distance between two adjacent peaks.

Abstract: A system is provided for controlling rotor blade over-tip leakage of a working fluid in rotating machine. The system has a circumferential row of rotor blades, and a circumferential row of seal segments for sealing with the radially outer tips of the rotor blades to reduce over-tip leakage of the working fluid. The seal segments have an inboard abradable layer which is adapted to be abraded by the blade tips to form a wear track for the blade tips. The system further has one or more tip wearing bodies which are adapted to wear down the blade tips. The thickness of the abradable layer and the cold build radial positions of the blade tips, seal segments and tip wearing bodies are arranged such that, during running-in of the machine, a wear track having a uniform radius is formed in the abradable layer by the blade tips, and the tip wearing bodies also wear down the blade tips to provide a uniform blade tip radius.

Abstract: A turbo-machine (1), including a stator (2), internally coated with a running-in layer (6), a rotor (4) within the stator (2), wherein the turbo-machine (1) supplementally includes a device for parallel displacement and rotation of the rotation axis of the rotor (10) about the axis of symmetry of the stator (2). By means of this device the gap width between stator (2) and rotor (4) is minimized and therewith the economy of the turbo-machine (1) is increased.

Abstract: A method of forming an interstage seal including removing a diaphragm seal box (14) from a gas turbine compressor assembly (10) and removing a labyrinth sealing member (12) from the diaphragm seal box (14). An abradable material layer (34) may be deposited on the diaphragm seal box (14). A spray gun may be mounted in relation to an engine disk (16) of the gas turbine compressor assembly (10) for cold-spraying a quantity of particles toward the engine disk (16). The particles may be sprayed at a velocity sufficiently high to cause at least a portion of the quantity of particles to adhere to the engine disk (16). The spray gun may be controlled to deposit a quantity of particles on the compressor disk (16) to form a geometry (32) that will abrade the abradable material layer (34) during operation of the gas turbine compressor assembly (10). The geometry (32) abrading the abradable material layer (34) forms an interstage seal.

Abstract: In one embodiment, an abradable seal includes a soft material that may be worn away and a hard material that may provide mechanical strength for the abradable seal. The soft material and the hard material may include different compositions of a base material.

Abstract: What disclosed is to form an abrasive coating having abrasiveness at a blade pressure side of a deposition layer using a second electrode composed of a second molded body molded from a mixed powder including a powder of a metal and a powder of a ceramic or the second electrode processed with a heat treatment, generating pulsing electric discharges between the blade pressure side of the deposition layer and the second electrode in an electrically insulating liquid or gas, and welding a material of the second electrode or a reacting substance of the material of the second electrode on the blade pressure side of the deposition layer by means of energy of the electric discharges.

Abstract: A pattern of depressions (36) in a sealing surface (34) on a CMC wall (32) of gas turbine ring segment (30) allows minimum clearance against turbine blades tips, and thus maximizes working gas sealing. An array of depressions (36) on the surface (34) increases abradability of the surface (34) by blade tip contact during zero clearance conditions and reduces blade tip damage. The depressions (36) are unconnected, preventing bypass of the working gas around the blade tips. A desired abradable surface geometry may be formed in a stacked laminate wall construction (40-43, 52) by staggered laminate edge profiles (50, 52) or by machining of depressions (36, 54) after construction.

Abstract: A fan rotor capable of avoiding coincidence will include a fan hub comprising one or more slots each designed to receive a fan blade; and one or more fan blades disposed within the slots. One or more slots will comprise an RZ? baseline and have one or more second fan blades disposed therein while another set of slots will have an RZ baseline having one or more first fan blades disposed therein. The tips of the second fan blades will be positioned at a distance farther from the fan hub than the tips of the first fan blades disposed, which when implemented in a turbofan engine will avoid coincidence occurrences.

Abstract: A multiple of circumferential grooves within said arcuate engine casing and an abradable material located radial inboard of the multiple of circumferential grooves.

Abstract: A device for attaching ring sectors around a turbine rotor in a turbomachine is disclosed. Each of the ring sectors includes at its upstream end a circumferential rim being able to be held on an annular casing rail by an annular locking member, and at its downstream end a pressing part that can press axially against a fixed element of the turbine to prevent the upstream rim of the ring sector from disengaging from the locking member in the event of considerable wear of the rail.