Abstract: A segmented turbine shroud for radially encasing a rotatable turbine in a gas turbine engine comprising a carrier, a ceramic matrix composite (CMC) seal segment, and an elongated pin. The carrier defines a pin-receiving carrier bore and the CMC seal segment defines a pin-receiving seal segment bore. The elongated pin extends through the carrier bore and the seal segment bore. The pin-receiving carrier bore includes a cantilevered member such that the carrier bore has a length sufficient to effect radial flexion between the carrier bore and the pin received within the carrier bore during operation of the turbine.

Abstract: Flow path assemblies for gas turbine engines are provided. For example, a flow path assembly comprises an inner wall and an outer wall defining an opening therethrough, as well as a nozzle airfoil having an inner end radially opposite an outer end that is positioned within the opening defined in the outer wall such that the outer end projects radially outwardly from the outer wall. The flow path assembly also comprises a pin and a first support member. The pin extends from the first support member to the outer end of the nozzle airfoil to support the nozzle airfoil. In some embodiments, the outer end of the nozzle airfoil defines a flange configured to receive the pin and a radial stop configured to press against a radially outward external support, and the pin is configured to transfer a load on the nozzle airfoil to an external structure.

Abstract: A turbine shroud for a gas turbine engine includes an annular metallic carrier and a ceramic-containing blade track. A connection is formed between the annular metallic carrier and the ceramic-containing blade track by insert pins extending through metallic carrier into the ceramic-containing blade track.

Abstract: Flow path assemblies and gas turbine engines are provided. A flow path assembly may comprise a combustor dome positioned at a forward end of a combustor of a combustion section of a gas turbine engine, and a unitary outer wall including a combustor portion extending through the combustion section and a turbine portion extending through at least a first turbine stage of a turbine section of the gas turbine engine. The combustor portion and the turbine portion are integrally formed as a single unitary structure. The flow path assembly also comprises an inner wall extending from the forward end of the combustor through at least the combustion section. The combustor dome extends radially from the unitary outer wall to the inner wall and is configured to move axially with respect to the inner wall and the unitary outer wall. Other flow path assemblies and gas turbine engine configurations are provided.

Abstract: A gas turbine engine is disclosed with a turbine section having at least one turbine rotor with a plurality of turbine blades, a plurality of blade tracks positioned circumferentially around the turbine blades, at least one dovetail shaped connecting member extending radially outward from each blade track, and a hanger connected to a structural member of the gas turbine engine and configured to releasably couple with the at least one dovetail shaped connecting member of a corresponding blade track.

Abstract: The present invention relates to a turbomachine component (1) or collection of components comprising at least a first and a second blade (3I, 3E) and a platform (2) from which the blades (3I, 3E) extend, characterized in that the platform (2) has a non-axisymmetric surface (S) bounded by a first and a second end plane (PS, PR) and defined by at least two class C construction curves each one representing the value of a radius of said surface (S) as a function of a position between the pressure face of the first blade (3I) and the suction face of the second blade (3E) in a plane substantially parallel to the end planes (PS, PR), these including at least one upstream curve and one downstream curve; each construction curve being defined by at least one pressure face control end point and one suction face control end point such that: —the tangent to the downstream curve at the suction face control end point 20 is inclined by at most 5°; —any other tangent to a construction curve at a control end point is inclined

Abstract: A gas turbine engine includes a turbine section including a plurality of blade outer air seals disposed therein, the blade outer air seals each including a body including a raised material that extends beyond the outer surface of the body and the raised material includes an inlet hole formed through the raised material.

Abstract: A method of manufacturing a turbine shroud segment for a gas turbine engine comprises: metal injection molding (MIM) a shroud segment body with a groove defined in a first lateral side thereof and with a flow restrictor projecting integrally from an opposite second lateral side thereof. The groove is oversized relative to the flow restrictor to provide for a clearance fit between the flow restrictor and the groove of adjacent turbine shroud segments when assembled together in a ring formation. The shroud segment body with the integrated flow restrictor are then subjected to debinding and sintering operations.

Abstract: Various embodiments of the disclosure include a turbomachine component. and methods of forming such a component. Some embodiments include a turbomachine component including: a first portion including at least one of a stainless steel or an alloy steel; and a second portion joined with the first portion, the second portion including a nickel alloy including an arced cooling feature extending therethrough, the second portion having a thermal expansion coefficient substantially similar to a thermal expansion coefficient of the first portion, wherein the arced cooling feature is located within the second portion to direct a portion of a coolant to a leakage area of the turbomachine component.

Abstract: An actuation system according to various embodiments can include an actuation ring having a first end and a second end separated by a gap, the actuation ring being configured to be coupled to a blade outer air seal (BOAS). The actuation system can also include an actuator coupled to at least one of the first end or the second end and configured to adjust a size of the gap such that a tip clearance between the BOAS and a blade tip is reduced in response to the size of the gap being reduced.

Abstract: A shroud segment that includes a body including a leading edge, a trailing edge, a first side edge, a second side, and a pair of opposed lateral sides. A first lateral side is configured to interface with a cavity having a cooling fluid, and a second lateral side is oriented toward a hot gas flow path. The shroud segment includes at least one channel disposed within the body, wherein the at least one channel includes a first portion extending from upstream of the trailing edge towards the trailing edge in a first direction from the leading edge to the trailing edge, a second portion extending from the trailing edge to upstream of the trailing edge in a second direction from the trailing edge to the leading edge, and a third portion extending from upstream of the trailing edge towards the trailing edge in the first direction.

Abstract: A turbine engine casing flow-path segment that is locally diffusing, followed by a flow-path segment contracting in the vicinity of a fan blade. This contraction accelerates the fluid flow axially forward of the fan blade leading edge at the tip and converges with the linear flow-path aft of the fan blade leading edge but forward of the fan blade trailing edge. More diffused fluid flow results in increased flow capacity of the fan, and increased fan efficiency.

Abstract: A turbine blade is described herein, the turbine blade including a blade root, a blade tip, and an airfoil extending between the blade root and the blade tip. The airfoil has opposite pressure and suction sides extending between a forward leading edge and an aft trailing edge of the airfoil, and a maximum thickness located between the leading edge and the trailing edge. The blade tip includes a winglet extending laterally outward from at least one of the pressure side and the suction side from a leading point between the leading edge and the maximum thickness aftward to a trailing point between the maximum thickness and the trailing edge.

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 carrier segment of a carrier section for circumscribing an array of circumferentially spaced turbine blades of a gas turbine engine, the blades being disposed radially inwardly of a turbine casing, the carrier segment including a carrier wall disposed radially inwardly of the casing and radially outwardly of the turbine blades, and the carrier wall including one or more portions facing the casing, wherein at least one of the one or more portions of the carrier wall is provided with one or more impingement apertures therein for passage therethrough of air of a predetermined temperature, heating air, from a feed source into impingement onto the turbine casing. The application of heating air onto the casing during transient stages of enhanced engine power, e.g. during step climbs, results in better matching of radial expansion of the casing relative to the turbine blades, thereby enabling improved blade tip clearance control.

Abstract: A rotor disk arrangement having a rotor disk which has, distributed around the outer circumference, a plurality of axially extending blade holding grooves and an encircling fastening protrusion and/or a plurality of fastening protrusions arranged in a circumferentially distributed manner, and having a plurality of sealing elements arranged in a circumferentially distributed manner, the sealing elements covering the blade holding grooves at least in portions on the end side and having, on the side facing the rotor axis, at least one fastening portion that bears against the underside of the fastening protrusion, wherein the fastening protrusion is embodied in an undercut manner, wherein the distance thereof from the rotor axis at the free end facing away from the rotor disk is less than in a region within the fastening protrusion.

Abstract: A shroud hanger with integral retainer assembly comprises a ceramic matrix composite shroud hanger a first wall and a second wall, the hanger having a support wall extending between the first and second walls, the support wall having a shoulder near circumferential ends, a retainer depending from the support wall having a first lower leg and a second lower leg extending in the circumferential direction, a first shroud supported by the first lower leg and a second shroud supported by the second lower leg.

Abstract: One exemplary embodiment of this disclosure relates to a gas turbine engine including an engine case, a retention block attached to the engine case, and a blade outer air seal (BOAS). The BOAS includes a plurality of layers formed of a ceramic matrix composite (CMC) material. At least one of the plurality of layers provides a slot receiving a portion of the retention block.

Abstract: A system includes a diffuser section comprising an outer barrel, an inner barrel, a seal interface, and an aft plate, where the outer barrel and the inner barrel are disposed about a turbine axis, and the diffuser section is configured to receive an exhaust gas from a gas turbine. The seal interface is disposed at a downstream end of the inner barrel between the inner barrel and the aft plate. The seal interface includes a first circumferential groove configured to receive the aft plate, where the first circumferential groove opens in a first direction away from the turbine axis.

Abstract: A compressor for a turbine engine, including: a casing, at least one compressor stage including an impeller having stationary blades and an impeller having moving blades positioned downstream of the stationary blade impeller, and cavities in a thickness of the casing that are disposed along a circumference of the casing opposite the moving blades. The cavities, which are elongate and extend along a main direction of orientation, are closed upstream and downstream by upstream and downstream faces respectively, and an upstream border and a downstream border are formed at the intersections between same and the casing. The cavities are offset in relation to the moving blades to overlap the moving blade impeller in the upstream portion, thereby covering the upstream end thereof. The downstream border of the cavities is oriented parallel to the chord at the head of the moving blade.

Abstract: A turbine includes blades and a structure rotating relative to the blades and having a fluid flowing thereto. The turbine includes step portions that are provided in either radial tip portions of the blades or areas of the structure that face the radial tip portions; sealing fins that extend from the other of the radial tip portions of the blades or areas of the structure that face the radial tip portions toward the step portions, and form minute gaps between the sealing fins and the step portions; a flow collision surface that is provided upstream of the sealing fins in a flow direction of the fluid and against which the fluid collides; a protrusion that protrudes toward an upstream side from the flow collision surface; and a facing surface that faces the flow collision surface.

Abstract: A blade of a compressor or of an axial-flow turbomachine. The blade comprises a joining airfoil, or main airfoil, extended by branches, or auxiliary branch airfoils. The branches are adjacent and extend the joining airfoil radially. The branches form a row over the circumference and comprise faces facing one another. The branches comprise joining edges connected to the joining airfoil and are coincident or joined so as to form a circumferential material continuation between the adjacent branches. The leading edges and the trailing edges of the branches are a continuation of the leading edges and the trailing edges of the joining airfoil. The branch airfoils offer more anchoring points for a blade. The overlap of the branch airfoils strengthens the edges of the joining airfoil, which can thus be formed thinner. The blade comprises a fixing platform connected to the branch airfoils.

Abstract: A device for axially arresting a sealing ring made of an abradable material in contact with a periphery of an aircraft turbomachine module rotor, the device having a first reference central axis, and including: a support including a support opening defining a circular track centered on a second axis at a distance from the axis; and an axial stop part including an extraction passage centered on a third axis at a distance from the axis, and defining a complementary circular track around its periphery, cooperating with the track to allow the part to rotate about the axis, between an axial stop position for the sealing ring, and a position to extract the ring through the extraction passage.

Abstract: A method of lining a turbofan casing (30) with an array (32) of liner modules (34), including the steps of: providing a liner module (34); and providing a coupling element (48) on the casing for interlocking engagement with the liner module to maintain the liner module in position relative to the casing; and interlockingly engaging the liner module (34) and coupling element (48) by progressive movement of the liner module relative to the coupling element.

Abstract: A turbine shroud for a gas turbine engine includes a plurality of shroud segments arranged around a central axis. Each shroud segment includes a carrier segment, a blade track segment, and a radial seal element that seals a radial interface between the carrier segment and the blade track segment. The turbine shroud further includes a plurality of circumferential seals that seal circumferential interfaces between the shroud segments.

Abstract: A blade includes a blade body extending from a blade root to an opposed blade tip surface along a longitudinal axis. The blade body defines a pressure side and a suction side. The blade body includes a cutting edge defined where the tip surface of the blade body meets the pressure side of the blade body. The cutting edge is configured to abrade a seal section of an engine case. A method for manufacturing a blade includes forming an airfoil with a root and an opposed tip surface along a longitudinal axis, wherein the airfoil defines a pressure side and a suction side. The method also includes forming a cutting edge where the tip surface of the airfoil meets the pressure side of the airfoil.

Abstract: An apparatus and method for controlling turbine blade tip clearance is disclosed herein. A blade track assembly can include a blade track carrier having a plurality of slots and rails defined by paths that vary in radial position as a function of circumferential location. An expansion ring can be operably coupled with the slots of the blade track carrier. A plurality of blade track segments can be operably coupled with the expansion ring and engageable with the rails of the blade track carrier such that expansion and contraction of the expansion ring causes radially outward movement and radially inward movement, respectively, of the blade track segments.

Abstract: A gas turbine engine component is provided. The gas turbine engine component has an internal cooling passage having first endwall and a second endwall. A primary inlet aperture is in fluid communication with a coolant supply and the first end of the internal cooling passage. An elongated rib is inside the internal cooling passage and proximate the primary inlet aperture. The elongated rib is configured to direct at least a portion of the coolant from the primary inlet aperture against a proximate one of the first or second endwall and then substantially reverse direction to flow downstream toward a second end of the internal cooling passage. A blade outer air seal (“BOAS”) assembly and a BOAS segment thereof are also provided.

Abstract: An air conditioning apparatus includes a casing having intake and blow-out ports, a partition member dividing an interior of the casing, a heat exchanger and a centrifugal fan. The centrifugal fan includes a hub and a bladed wheel having a plurality of rearward blades and sucks air existing in the heat exchanger compartment into the fan compartment through the fan entrance. The bladed wheel is mounted in the fan compartment such that a rotary shaft of the bladed wheel is oriented along an opening direction of the fan entrance and an opening direction of the blow-out port. The hub connects blow-out port side ends of the plural rearward blades and is configured to be rotated about the rotary shaft. The hub has an outer diameter smaller than an outer diameter of the rearward blades.

Abstract: A device (1) for transmitting electrical current and/or signals in a motor vehicle is proposed. The device (1) has a wind up spring cassette (2) composed of a stator (4) and a rotor (5) in which at least one line (7) extending in windings is connected with one end thereof to the stator (4) and with the other end to the rotor (5), and at least one fixing element (3) through which a predetermined relative position for the mounting of the wind up spring cassette (2) into the motor vehicle is fixed. The fixing element is a foil like label (3) which is attached to the outside of the wind up spring cassette (2) and extends from the rotor (5) to the stator (4). The label (3) is inseparably mechanically secured at the stator (4) as well as at the rotor (5).

Abstract: A structural assembly for a fluid-flow machine includes: a main flow path boundary and at least one row of relatively rotating blades with a gap existing between blade ends of the at least one row of blades and the main flow path boundary. At least one secondary flow duct has in the main flow path boundary one opening each at ends spaced apart in the flow direction, such that the secondary flow duct is connected to the main flow path via the two openings. The structural assembly has at least two components connected to one another, i.e. at least one support component and at least one connecting component, where the support component at least partially forms the main flow path boundary and where the connecting component forms or surrounds at least one part-section of the secondary flow duct.

Abstract: A shroud apparatus for a gas turbine engine includes: an annular shroud segment having an arcuate bottom wall defining an arcuate inner flowpath surface, spaced-apart forward and aft walls extending radially outward from the bottom wall, and spaced-apart side walls extending radially outward from the bottom wall and between the forward and aft walls, each side wall defining an end face which includes: an axial slot extending in a generally axial direction along the end face; a first radial slot extending in a generally radial direction along the end face, and intersecting the axial slot; an axial spline seal received in the axial slot; and a first radial spline seal having an L-shape with radial and axial legs, the radial leg being substantially longer than the axial leg, wherein the radial leg is received in the first radial slot, and the axial leg is received in the axial slot.

Abstract: A method is provided for calibrating an active clearance control system for a plurality of turbine engines. During this method, a squeeze test is performed between a tip of a rotor blade and a shroud. Results of the squeeze test are applied to adjust a gap between the tip and the shroud. The performance of the squeeze test and the application of the results may be individually performed for each of the turbine engines.

Abstract: A bladed rotor assembly includes a bladed rotor and a rotor housing. The bladed rotor has a plurality of blades arranged in a circumferential array around an axis of rotation, and the rotor housing encircles the bladed rotor and has a radially inwardly facing surface. Each of the blades comprises a first, radially extending, edge, a second opposite, radially extending, edge, and a radially distal edge. The radially distal edge has a first end, a second end, and a center portion, with the first end adjoining the first, radially extending, edge, and the second end adjoining the second, radially extending, edge. A clearance between the radially distal edge and the radially inwardly facing surface is a first clearance at the first end, and decreases to a second clearance across the center portion.

Abstract: A blade outer air seal (BOAS) includes a body defining a core having a forward cavity portion and a plurality of forward hooks extending radially from the body, the hooks extending to a hook height. The BOAS includes a seal surface defined between the forward hooks and which includes a forward side, an aft side, and a seal surface height. The seal surface is aft of the forward cavity portion. The BOAS includes a thermal regulation channel defining an inlet aft of the sealing surface and an outlet defined in fluid communication with the forward cavity portion.

Abstract: An assembly is provided for a turbine engine with an axial centerline. This turbine engine assembly includes a blade outer air seal segment, a linkage, a rocker arm and an actuation device. The linkage is attached to the blade outer air seal segment. The rocker arm includes a first arm and a second arm engaged with the linkage. The actuation device is engaged with the first arm. The actuation device is configured to pivot the rocker arm and thereby radially move the blade outer air seal segment.

Abstract: A multi-stage axial compressor with an inner wall including a step portion for each of the compressor stages. Each step portion is defined along a respective stage. Each step portion may extend over at least a majority of an axial length of the stage. Each step portion may optionally include a point aligned with a maximum thickness of the airfoil portions of the rotor blades and a point aligned with a maximum thickness of the stator vanes. Adjacent step portions are connected by a transition portion converging toward a central axis of the compressor from the upstream step to the downstream step. Each transition portion has a steeper slope than that of the adjacent step portions. A method of directing flow through a multi-stage axial flow compressor is also discussed.

Abstract: Various embodiments of the disclosure include a turbomachine component. and methods of forming such a component. Some embodiments include a turbomachine component including: a first portion including at least one of a stainless steel or an alloy steel; and a second portion joined with the first portion, the second portion including a nickel alloy including an arced cooling feature extending therethrough, the second portion having a thermal expansion coefficient substantially similar to a thermal expansion coefficient of the first portion, wherein the arced cooling feature is located within the second portion to direct a portion of a coolant to a leakage area of the turbomachine component.

Abstract: An assembly for a turbine engine includes a turbine engine first component, a turbine engine second component and a flexible seal that is attached to the first component. The flexible seal at least partially seals a gap between the first component and the second component. The flexible seal includes a mount and a finger seal that sealingly engages the second component. The mount includes a boss that sealingly engages the first component.

Abstract: The present application provides a turbine component for use in a hot gas path of a gas turbine. The turbine component may include an outer surface, an internal cooling circuit, a number of cooling pathways in communication with the internal cooling circuit and extending through the outer surface, and a number of adaptive cooling pathways in communication with the internal cooling circuit and extending through the outer surface. The adaptive cooling pathways may include a high temperature compound therein.

Abstract: A seal segment for bounding a hot gas flow path within a gas turbine engine, including: a plate having an inboard side which bounds the hot gas flow path in use, an outboard side and fore and aft cooling circuits, wherein the fore and aft cooling circuits are fluidically separated from one another within the plate and each has at least one tortuous path between an inlet on the outboard side of the plate and an exhaust.

Abstract: An air-cooled heat exchange system includes a heat exchanger and an axial fan assembly. The heat exchanger has a heat exchange surface area. The axial fan assembly has a propeller-type impeller rotatably supported within an annular housing. The annular housing defines an air passageway from an air inlet end of the housing, across the impeller, and to an air outlet end of the housing. Rotation of the impeller within the annular housing causes air to flow into the air inlet end of the housing, along the air passageway, and out the air outlet of the housing. The axial fan assembly also has a wind-turning vane assembly extending beyond the air inlet end of the annular housing.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an engine case, a rotor stage including a plurality of rotor blades, a plurality of carriers for supporting a plurality of blade outer air seals and an interlock formed between circumferential ends of a first adjacent carrier and a second adjacent carrier of the plurality of carriers.

Abstract: A seal segment is positioned, in use, radially adjacent the rotor. The seal segment has first and second circumferentially spaced passageways each of which extends in the fore and aft direction. In use, a first support bar is contained within the first passageway, and a second support bar is contained within the second passageway. The first and second support bars being mountable to complementary formations provided by the casing of the engine. The first passageway is configured such that the seal segment is fixed relative to the first support bar in the radial and circumferential directions. The second passageway is configured such that the seal segment is fixed relative to the second support bar in the radial direction but allows relative movement of the seal segment and the second support bar in the circumferential direction to accommodate differential thermal expansion of the seal segment and the casing.

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 method for machining a workpiece. An electrode is moved linearly in the direction of the workpiece to cause material to be removed from the workpiece, at least one end of a surface of the workpiece running obliquely to a guide edge of the electrode machining this surface. The electrode is moved at least partially with the electrode surface parallel to the surface, so that during the approach to the workpiece, areas of the workpiece having an irregular edge machined at a different intensity are formed, the difference in intensity of machining on the edge of the surface to be machined being compensated in that the surface to be machined is provided with a height profile adapted to the shape of the end of the surface to be machined. A blade ring segment and blade ring is also disclosed.

Abstract: A component of a gas turbine engine is formed from a continuous fiber reinforced ceramic matrix composite (CMC). The component has a sealing portion which, in use, makes sealing contact with an adjacent component of the engine. The sealing portion comprises a recess formed in the CMC and filled with a finer grade ceramic relative to the surrounding CMC, a metal or an intermetallic.

Abstract: A method of characterizing a turbine engine, the method comprising: steps of using measurements from an accelerometer monitoring a particular turbine engine while it is starting to detect the energy released by any contact during said starting between the rotor and the stator of the turbine engine, and of associating any such detection of contact with thermodynamic data measured on the particular turbine engine; and then a recognition training step, based on the associations, to enable a thermal context of the turbine engine to be used to recognize the presence of a rotor thermal unbalance to be taken into account in order to avoid contacts between the rotor and the stator on starting.

Abstract: An airfoil for a gas turbine engine includes pressure and suction walls spaced apart from one another and joined at leading and trailing edges to provide an airfoil having an exterior surface that extends in a radial direction to a tip. A camber line at the tip extends from the leading edge to the trailing edge. Pressure and suction side shelves are arranged in the exterior surface on opposing sides of the camber line respectively in the pressure and suction side walls. A plateau is proud of and separates the pressure and suction side shelves. The plateau is arranged along the camber line and extends to the leading edge.

Abstract: A turbine airfoil usable in a turbine engine and having at least one snubber with a snubber cooling system positioned therein and in communication with an airfoil cooling system is disclosed. The snubber may extend from the outer housing of the airfoil toward an adjacent turbine airfoil positioned within a row of airfoils. The snubber cooling system may include an inner cooling channel separated from an outer cooling channel by an inner wall. The inner wall may include a plurality of impingement cooling orifices that direct impingement fluid against an outer wall defining the outer cooling channel. In one embodiment, the cooling fluids may be exhausted from the snubber, and in another embodiment, the cooling fluids may be returned to the airfoil cooling system. Flow guides may be positioned in the outer cooling channel, which may reduce cross-flow by the impingement orifices, thereby increasing effectiveness.