Abstract: A compressor for a turbocharger having a compressor wheel with compressor blades, and being mounted for rotation on a shaft. The compressor has a shroud mounted adjacent the wheel and defining a gas flow path between the shroud and the blades from a compressor inlet to a diffuser outlet. In cross-section the shroud has a surface in the flow path, the surface having a profile that includes a section with a smoothly curving surface and at least one relative discontinuity forming a downstream-facing face.

Abstract: An example turbine engine sealing arrangement includes a blade array rotatable about an axis. The blade array has a plurality of blades extending radially from the axis. A control ring is circumferentially disposed about the blade array. A plurality of tiles are secured relative to the control ring and configured to establish an axially extending seal with one of the blades.

Abstract: A stacked laminate component for a turbine engine that may be used as a replacement for one or more metal components is provided. The stacked laminate component can have a body formed by a process of stacking and laminating layers to define a radially inner surface along the hot gas path. The layers can be substantially orthogonal to the radially inner surface. The layers can be at least a first layer of a first material and a second layer of a second material. At least the first material is a ceramic matrix composite. The second material can have a higher thermal conductivity or a higher creep strength than the first material.

Abstract: A blade outer air seal for a gas turbine engine has a plurality of stacked panels extending for at least a part circumferential extent, and extending from a radially outer location to a radially inner location. The plurality of stacked panels are mounted together. A turbine section is also claimed.

Abstract: A propeller fan has a gap between a bellmouth and each blade tip. A leakage flow is thus produced from a positive pressure surface of a blade toward a negative pressure surface of the blade via the gap. The leakage flow develops while flowing from the leading edge of each blade toward the trailing edge, thus forming a blade tip vortex. This increases blowing noise and raises the input of the fan motor. In order to suppress the blade tip vortex caused by the leakage flow, the present invention includes recessed portions and projected portions, which are alternately formed in a portion of a blade surface of each blade corresponding to the blade tip.

Abstract: A system for supporting a shroud used in an engine has a shroud positioned radially outboard of a rotor, which shroud has a plurality of circumferentially spaced slots; a forward support ring for supporting the shroud; the forward support ring having a plurality of spaced apart first tabs on a first side for functioning as anti-rotation devices; the forward support ring having a plurality of spaced apart second tabs on a second side; and the second tabs engaging the slots in the shroud and circumferentially supporting the shroud.

Abstract: A gas turbine engine component includes first and second wall surfaces, an inlet located at the first wall surface, an outlet located at the second wall surface and a diffusing section positioned between the inlet and the outlet. The diffusing section includes a first lobe, a second lobe adjacent the first lobe and a third lobe adjacent the second lobe. The first lobe and the second lobe meet at a first ridge and the second lobe and the third lobe meet at a second ridge.

Abstract: A gas turbine article includes a substrate, a plurality of geometric surface features that protrude from the substrate, and a thermally insulating topcoat disposed over the plurality of geometric surface features. The thermally insulating topcoat includes segmented portions that are separated by faults that extend through the topcoat from the geometric surface features.

Abstract: A turbine rotor blade with a hollow open cavity formed between the airfoil walls, and radial flow near wall cooling channels formed within the walls that open on the tip surface to discharge the cooling air for sealing of the tip. The radial channels have a converging flow area to increase the cooling air flow velocity, and the radial cooling channels on the pressure side wall are slanted toward the pressure side so that the discharged cooling air will produce a smaller vena contractor and further reduce tip leakage flow.

Abstract: A sealing system and method is disclosed. The sealing system includes a brush sealing element and an accommodating element with an accommodating chamber. The brush sealing element is disposed in the accommodating chamber and a bendable strap is disposed on the accommodating element. The method includes fastening the brush sealing element in the accommodating chamber of the accommodating element. The accommodating element is arranged in a groove of a housing segment and a position of the accommodating element is secured on the housing element by bending the strap of the accommodating element to engage with a side wall of the housing segment.

Abstract: A pump for chopping solid material in a liquid stream having a volute housing, a rotating cutter with at least one blade parallel to a rotational axis of the rotating cutter, an impeller arranged at an outer circumference of the rotating cutter, and a stationary cutter having at least one blade parallel to the rotational axis of the rotating cutter. The stationary cutter is concentric with the rotating cutter, and the stationary cutter, the rotating cutter and the impeller are all at least partially housed within the volute housing. Typically, the rotating cutter has a different number of blades than the stationary cutter. Optionally, the pump, further comprises an inspection cover that is removably attached to the pump, and has an opening large enough to remove the stationary cutter and the rotating cutter from the volute housing.

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: A turbine blade includes an airfoil tip with first and second ribs extending along the opposite pressure and suction sides. The ribs extend outwardly from a tip floor and are joined together at opposite leading and trailing edges. A cascade tip baffle transversely bridges the two ribs above the tip floor forward of the maximum width of the tip to partition the tip chordally into corresponding tip pockets on opposite sides of the baffle.

Abstract: A turbofan engine has a fan having a circumferential array of fan blades. A fan case encircles the fan. There is at least one compressor section, a combustor, and at least one turbine section. The fan case comprises a composite structural member and a metallic member encircled by the composite structural member. The metallic member is mounted to the composite structural member to permit differential thermal expansion proximate the 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: A sealing arrangement for a turbomachine for sealing a radial gap between a rotor and a stator, with a plurality of sealing segments each comprising a honeycomb element and a base body for holding the honeycomb element, and with at least one support for arranging the sealing segments on a rotor section or stator section, wherein the base bodies can be arranged on the at least one support by a radial movement and by a plastic deformation of at least one body section, a generative method for the integral production of the sealing segments, and a turbomachine are disclosed.

Abstract: A device for sealing a gas path in a turbine includes a first shroud segment and a slot in a surface of the first shroud segment. A barrier extends inside the slot, and a bypass channel in the slot provides fluid communication between the barrier and the slot to the gas path in the turbine. A method for sealing a gas path in a turbine includes placing a barrier between a first slot in a first shroud segment and a second slot in a second shroud segment and flowing a fluid between the barrier and the first slot to the gas path in the turbine, wherein the fluid flows through a first bypass channel in the first slot.

Abstract: A cooling system of a ring segment for a gas turbine, which includes first cooling passages disposed in an axial direction of the rotating shaft of a main body of the segment body, second cooling passages disposed at one end portion in a direction approximately perpendicular to the first cooling passages, and blowing a cooling air toward the end portion of a neighboring segment body, and third cooling passages connecting a first cavity, which is disposed approximately perpendicular to the axial direction of the rotating shaft at the upstream-end portion, with a cooling space, which is surrounded by the segment body and a collision plate having a plurality of small holes. The first cooling passages include cooling passages of a first region and a second region having a smaller passage cross-sectional area than the first-region cooling passages or a greater arrangement pitch than the first-region cooling passages.

Abstract: The present application provides a stage of a gas turbine engine. The stage may include a bucket, a shroud facing the bucket, and a forward step honeycomb seal on the shroud. The forward step honeycomb seal may include a forward step portion and one or more linear portions.

Abstract: A Blade Outer Air Seal (BOAS) includes a body manufactured of a metal alloy, the body includes a face opposite a forward interface and an aft interface, the face includes a cavity. A non-metallic insert within the cavity such that the insert is flush with the face.

Abstract: A ring segment for a turbine in an industrial gas turbine engine, the ring segment having an inner side with a number of pedestals extending radially inward, each pedestal having an inlet metering hole connected to a diffusion chamber having an opening flush with a TBC covering the pedestals. The pedestals form a larger surface area to secure the TBC to the ring segment so that the TBC can be formed thicker than the prior art without spalling.

Abstract: A BOAS with a blade tip shroud having a row of teeth formed by rows of grooves on an underside surface, and where rows of teeth include cooling air passages to provide cooling for the blade tip shroud and to discharge cooling air into the blade tip gap to reduce an effective hot gas leakage area to reduce the blade tip leakage flow. The cooling passages include inlet convergent channels to increase the cooling air flow velocity, followed by flow metering sections and then cooling air exit slots that open onto the bottom surface of the teeth.

Abstract: Provided is a seal structure (50, 60) configured to seal a clearance (31, 32) with respect to a flow path defined by a rotary body (20) rotated about a central shaft (P) and configured to allow passage of a mainstream, and a stationary body (10) disposed at the rotary body (20) to form the clearance (31, 32), the seal structure including a seal member (51A, 51B) installed at the clearance (31, 32) and configured to divide a fluid introduced from the mainstream into the clearance (31, 32) into a high pressure side and a low pressure side, and a shifting member (52A, 52B) installed at the high pressure side of the fluid in the clearance (31, 32), and configured to allow passage of the fluid and reduce a velocity component in a rotational direction of the rotary body (20) of the velocity component of the fluid.

Abstract: A compressor stator blade 10 includes a seal mechanism 12 configured to hermetically seal a part between a radial inner end on the upstream side and an outer surface of a rotary body, and a remainder part on the downstream side of the seal mechanism includes a hub clearance 14 between the outer surface of the rotary body and itself. Additionally, a compressor rotor blade 20 includes a seal mechanism 22 configured to hermetically seal a part between a radial outer end on the upstream side and an inner surface of a stationary body, and a remainder part on the downstream side of the seal mechanism includes a tip clearance 24 between the inner surface of the stationary body and itself.

Abstract: Disclosed herein is a production method of a shroud segment that includes a forming process of molding a cylindrical fiber fabric (10) into a shroud segment shape by pressing a cylindrical surface of the fiber fabric, and a matrix forming process of impregnating the fiber fabric molded into the shroud segment shape with a matrix.

Abstract: A BOAS segment for a turbine of a gas turbine engine where the BOAS segment includes an arrangement of micro sized serpentine flow cooling channels that cover the entire BOAS surface, with each serpentine channel discharging into break out holes that open onto the four sides of the BOAS in diffusion openings. The serpentine channels on the leading edge side are shorter while the serpentine channels for the trailing edge are longer because of the different metal temperature.

Abstract: A BOAS with a blade tip shroud having rows of sinusoidal shaped grooves on a bottom side, each groove having a row of thin slots extending along the aft side of the groove and opening into the groove, and each thin slot connected to a plurality of metering holes that open into a cooling air supply cavity formed on the top surface of the tip shroud.

Abstract: A BOAS assembly for a gas turbine engine with a full annular impingement ring secured in place above BOAS segments between front and aft full retainer rings and the front and aft hooks on the BOAS segments. The full impingement ring uncouples the impingement ring from the BOAS segments and eliminates gaps for cooling air leakage so that all of the cooling air is used for impingement cooling of the BOAS segments. A long tongue extending from the aft retainer rings guides the impingement ring and the BOAS segments into place during assembly.

Abstract: A turbine engine blade outer air seal segment has a body having a base portion. The base portion has a transversely concave ID face, a forward end, an aft end, and first and second circumferential edges. The body has at least one mounting hook. At least one cover plate is secured to the body to define at least one cavity. The cover plate has a plurality of feed holes. A plurality of outlet holes extend through the base portion to the ID face. At least one of the base portion and cover plate comprises a protruding portion protruding into the cavity to form a partial restriction separating forward and aft cavity portions.

Abstract: A gas turbine engine component comprises a plurality of walls, a cooling channel, a plurality of ribs and a plurality of pedestals. The plurality of walls has a pair of major surfaces opposed to define an interior chamber. The cooling channel extends through the interior chamber of the plurality of walls between the major surfaces. The plurality of ribs extends through the cooling channel to form a plurality of wavy passages having bowed-out sections. The plurality of pedestals is positioned between adjacent ribs, each pedestal being positioned in a bowed-out section.

Abstract: A fiber preform for a turbine ring sector obtained by three-dimensional weaving and including a base-forming first portion, two tab-forming L-shaped portions each presenting two branches, two ends of the first portion being extended by respective ones of the first branches, and a second portion connecting together the two tabs, first and second strips woven together forming the first branches, a first fraction of the thicknesses of the second branches, and the first portion, there being a non-linked zone between them that is situated in the first branches and in the first portion, and a third woven strip forming the second portion and a second fraction of the thickness of the second branches of each of the tabs.

Abstract: A fan casing (2) for a jet engine has a burst protection arrangement (7) made up of layers at least in an area of fan blades (5). To contain detached fan blades (5) within a fan casing (2) of a jet engine by use of a low-weight and easily manufacturable burst protection arrangement (7), the fan casing (2) is made of a fiber-composite material which itself forms at least one layer provided with at least one inorganic, non-metallic protective layer (8).

Abstract: A turbine blade includes an airfoil having first and second tip ribs extending along the opposite pressure and suction sides thereof. The tip ribs are spaced apart between the leading and trailing edges of the airfoil to include a trifurcate tip baffle defining a triforial tip cavity.

Abstract: A turbine shroud ring sector for a turbomachine supported at the upstream end by a downstream support of a turbine casing with circular sliding including a first stop able to collaborate with a second stop borne by an element of the turbomachine adjacent to the shroud ring in order to immobilize the shroud ring circularly is disclosed. The sector includes, on an end facing the element, a recess able to allow the second stop to pass to come into contact with the first stop. The recess is cut substantially in the form of a rectangle having, at the bottom of the recess, corners that are rounded in a circular arc of radius r. The bottom of the recess has a convex shape tangential to the circular arcs of the rounded corners and with a curvature that evolves between a radius of curvature r where it meets the rounded corners and a radius of R, greater than the radius r, at a point situated between the two rounded corners.

Abstract: A shroud for a gas turbine engine includes a leading portion having a leading edge and a first set of circumferentially spaced slots at the leading edge that break up the leading portion into circumferentially spaced segments separated by the first set of slots, and a trailing portion adjacent to the leading portion. The trailing portion has a trailing edge.

Abstract: A bucket assembly is disclosed. The bucket assembly includes an airfoil having a generally aerodynamic contour and defining a tip, and a lower body portion extending generally radially inward from the airfoil. The bucket assembly further includes a tip shroud disposed on the tip of the airfoil and comprising a main body and a rail. The rail includes an exterior surface. The exterior surface defines a microchannel. The bucket assembly further includes a cover layer configured on the exterior surface.

Abstract: A seal structure includes shrouds that are provided at tips of a plurality of blades on which a main flow of fluid impinges and that come into contact with adjacent shrouds to form a cylindrical shape; a sealing part connected to a structure body facing the shrouds; and a space that is defined by downstream sides of the shrouds, a downstream side of the sealing part, and a wall on a downstream side of the structure body, and that communicates with the main flow of fluid that has impinged on the blades. An end of the wall on the downstream side of the structure body is formed so as to be disposed a predetermined distance away from a shroud extension and away from a sealing-part extension and so as to be located on the opposite side of the shroud extension with respect to the sealing-part extension.

Abstract: A seal assembly that limits gas leakage from a hot gas path to one or more disc cavities in a turbine engine. The seal assembly includes a seal apparatus that limits gas leakage from the hot gas path to a respective one of the disc cavities. The seal apparatus comprises a plurality of blade members rotatable with a blade structure. The blade members are associated with the blade structure and extend toward adjacent stationary components. Each blade member includes a leading edge and a trailing edge, the leading edge of each blade member being located circumferentially in front of the blade member's corresponding trailing edge in a direction of rotation of the turbine rotor. The blade members are arranged such that a space having a component in a circumferential direction is defined between adjacent circumferentially spaced blade members.

Abstract: A fluid flow machine has a flow path provided by a casing (1) and a rotating shaft (2), in which rows of rotor blades (3) and stator blades (4) are arranged, and includes at least one annular groove-type recess (5) being disposed in a blade (3, 4) tip area in an annulus duct wall of the casing (1) and/or the shaft (2). An upstream end point (E) of the recess (5) in a flow direction is set at a distance (e)>0 forward of a forward blade tip point (A), and a downstream end point (F) of the recess (5) is set at a distance (f) rearward of point (A), where: 0.5 L>(f)>0, and L is a distance between point (A) and a rearward blade tip point (B).

Abstract: A sealing assembly for a turbine engine has a brush configured to provide an air seal between an outer air seal segment and a blade. A support is provided for the brush and is configured for attachment to the outer air seal segment. The support is a flexible planar carrier having a first side and a second side. The brush extends from the first side while the second side is for attachment to the outer air seal segment. The flexible planar carrier has a stiffness less than a stiffness of the outer air seal segment.

Abstract: A component, in particular for a gas turbine, is disclosed. The component has at least two component segments which are arranged relative to one another leaving a gap and are sealed against each other by a sealing device. The sealing device includes at least one brush seal. Also disclosed is a gas turbine, in particular an aircraft engine, having such a component.

Abstract: A rotor blade 2 for a gas turbine engine, comprising a blade tip 4. At least a portion of the blade tip 4 is provided with a wear-resistant coating 16, 18 comprising magnesium aluminate spinel.

Abstract: An air seal assembly for a gas turbine engine the air seal comprises a first assembly and a second assembly. One of the first assembly and the second assembly is rotatable relative to the other of the first assembly and the second assembly. The second assembly is aligned annularly with the first assembly and includes a circumferential surface with an abradable coating disposed annularly adjacent to the first and second airfoil tips. The first assembly includes at least one first airfoil with a first tip having an abrasive coating, and at least one second airfoil with a second tip absent the abrasive coating, the at least one first airfoil co-aligned axially and intermingled with a respective at least one second airfoil around a periphery of the first assembly.

Abstract: A turbine stage in a turbomachine including ring sectors arranged about a turbine impeller, and an annular support supporting the ring sectors and attached to a turbine casing is disclosed. The annular support is able to be deformed elastically in the radial direction to cushion at least a portion of the deformations of the turbine casing in operation.

Abstract: A thermal control system for a turbine casing (2) mounted radially outwardly of and circumferentially surrounding blade tips of the turbine blades, the system comprising: one or more circumferential chamber walls (6), each disposed around the casing (2), each to form a circumferential flow chamber (7) radially outward of said surface of the turbine casing (2), each to form a circumferential gap (12) between the circumferential chamber wall (6) concerned and said surface of the turbine casing (2), and the one or more circumferential chamber walls (6) define at least one opening (9) on the radially inward side of the chamber (7) adjacent the surface of the turbine casing (2) such that a thermal fluid injected into the flow chamber (7) concerned may leak out through the circumferential gap (12) via the opening (9), and means to induce circumferential flow on the surface of the casing (2) bounded by the circumferential flow chamber (7) comprising one or more ducts (8) radially outwardly connected to the or each c

Abstract: Composite containment casings including a body having an interior and a fragment catcher integrally joined to the interior of the body of the containment casing.

Abstract: An example blade outer air seal assembly may consist of a blade outer air seal containing channels that communicate cooling air through at least some of the blade outer air seal. The blade outer air seal has at least one circumferentially extending barrier separating the leading channel into a forward portion and an aft portion. Cooling air outlets from the leading channel are exclusively coupled to the aft portion.

Abstract: Systems and devices for substantially fluidly isolating portions of a rotary machine are disclosed. In one embodiment, a seal assembly for a rotary machine includes: a complementary set of seal segments configured relative one another so as to form a substantially continuous body with butt gaps between the seal segments, the segments including: a head flange; and an axial neck connected to the head flange, wherein the axial neck extends radially inboard and includes an axial slot formed substantially there through; and at least one butt gap seal element, wherein the at least one butt gap seal element includes an axial flange and a radial flange, the axial flange disposed within the axial slots of adjacent seal segments; and the radial flange extending radially inboard from the axial slots to a proximity of the rotary machine substantially forming a barrier to axial flow through the butt gap.

Abstract: A seal carrier for a seal used between rotating and non-rotating components includes an arcuate sheet metal seal carrier body formed to include a stepped, conical configuration wherein an inner diameter at a forward end is larger than the diameter at an aft end, with plural stepped sections defined by alternating radial and axial portions between the forward end and the aft end. Each axial portion is adapted to carry a seal element, preferably a honeycomb seal. Mounting flanges are provided at the forward end and the aft end of the arcuate sheet metal seal carrier body to enable attachment to the non-rotating component.

Abstract: A ring segment for a turbine in a gas turbine engine, the ring segment having a backside impingement cavity to collect spent impingement cooling air, and a number of separate impingement chambers spaced around all four edges of the ring segment to provide impingement cooling for the edges. Each impingement chamber is connected to the central impingement cavity by a number of metering and impingement holes to supply cooling air and a number of cooling air exit holes to discharge the cooling air out the sides of the four edges. The trailing edge of the ring segment includes two rows of impingement chambers connected in series.