Abstract: A tangential on-board injector (TOBI) having: a body defining an annular passageway, discharge nozzles; a rotating component configured to be mounted for rotation relative to the body about an axis of rotation; a seal extending between the body and the rotating component; and flow passages circumferentially distributed about the axis of rotation and in fluid communication with the nozzles and the seal, the flow passages located upstream of the seal relative to a flow of the cooling air circulating toward the seal from the plurality of discharge nozzles, each of the flow passages extending along a respective passage axis, the passage axis of at least one of the flow passages having a tangential component at an outlet of the at least one of the flow passages that is different than a tangential component of an exit flow axis of at least one of the plurality of discharge nozzles.

Abstract: A ring segment, a turbine, and a gas turbine having improved cooling performance are provided. The ring segment may include a shielding wall mounted to a turbine casing which accommodates a turbine blade and configured to face an inner circumferential surface of the turbine casing, a first hook part and a second hook part configured to protrude from the shielding wall toward the turbine casing to be coupled to the turbine casing, a main cavity formed between the first hook part and the second hook part, a plurality of first cooling passages configured to connect the main cavity and first side surfaces facing each other of the shielding wall, a plurality of second cooling passages configured to extend in a direction crossing the first cooling passage and connect the main cavity and second side surfaces facing each other of the shielding wall, and a chamber configured to be connected to the first cooling passages.

Abstract: A gas turbine engine component having a cooling passage includes a first wall defining an inlet of the cooling passage, a second wall generally opposite the first wall and defining an outlet of the cooling passage, a metering section extending downstream from the inlet, and a diffusing section extending from the metering section to the outlet. The metering section includes an upstream side and a downstream side generally opposite the upstream side. At least one of the upstream and downstream sides includes a first passage wall and a second passage wall where the first and second passage walls intersect to form a V-shape.

Abstract: A turbine airfoil includes a trailing edge coolant cavity between a pressure sidewall and a suction sidewall. The trailing edge coolant cavity is positioned adjacent to and extending out to a trailing edge of the turbine airfoil. The interior includes an internal arrangement comprising an array of discrete fins formed aft of the trailing edge coolant cavity along the trailing edge.

Abstract: A gas turbine nozzle assembly of a gas turbine is provided. The turbine nozzle assembly may include a turbine nozzle extending from an inner platform to an outer platform and having an airfoil-shaped cross section having a leading edge and a trailing edge, and a pressure side and a suction side each of which extends from the leading edge to the trailing edge, wherein the turbine nozzle may include a plurality of vanes attached to the inner and outer platforms and the inner platform having an attached first and second endfaces and a flow surface surrounding opposing ends of a vane of the plurality of vanes, the flow surface terminating circumferentially at the first and second endfaces and terminating axially at forward and aft edges, and the inner platform may include a platform corner portion comprising the flow surface attached to the first endface at the forward edge and attached to the second endface at the aft edge.

Abstract: A gas turbine engine includes a primary flowpath connecting a compressor section, a combustor section and a turbine section. The turbine section includes a stage vane having a radially outward platform and a vane extending into the primary flowpath. The platform includes a cooling plenum. At least one retaining feature extends radially outward from the platform. At least one vectored cooling hole is disposed in the retaining feature and is configured to direct cooling air from the plenum to an adjacent gaspath component.

Abstract: A component for a gas turbine engine. The component includes a body. The body has an exterior surface abutting a flowpath for the flow of a hot combustion gas through the gas turbine engine. Further, the body defines a cooling passageway within the body to supply cool air to the component. The component includes a leading face and a trailing face defining a trench therebetween on the exterior surface. The body defines a plurality of cooling holes extending between the cooling passageway and a plurality of outlets defined in the trench such that the trench is fluidly coupled to the cooling passageway. Additionally, the leading face and trailing face are each tangent to at least one of the plurality of outlets. The trench directs the cool air along a contour of the component.

Abstract: During operation of a gas turbine engine, the turbine blades are subjected to extremely high temperatures. While internal cooling passages may be provided within the turbine blades to cool the airfoils, the platforms of the turbine blades also experience high temperatures. In an embodiment, cooling holes are provided in each platform to cool distressed locations in the platform and improve the durability of each turbine blade.

Abstract: A blade outer air seal for a gas turbine engine includes a platform having a leading edge and a trailing edge, and at least one cooling cavity disposed within the platform. A pair of circumferential edges connects the leading edge and the trailing edge. An end wall protrudes radially outward from the platform at the trailing edge. A support hook protrudes radially outward from a leading edge of the platform. A boss structure protrudes radially outward from the platform and defining a solid kinetic energy path from an inner diameter of the platform to a radially outward facing surface of the boss structure, wherein the boss structure extends less than a circumferential length of the platform.

Abstract: A cooling structure on a trailing-edge cutback of a turbine blade, including a plurality of lands, a trailing edge cutback and a dimple group. Adjacent lands are arranged on wall surfaces at two sides of the trailing edge cutback. The wall surfaces are each provided with the dimple group including multiple dimples. An extension direction of at least one dimple forms an inclined angle with the land on one side, and/or an extension direction of at least one dimple forms an inclined angle with the land on the opposite side. The cooling air enters the trailing edge, and after passing through pin fins, then flows over the dimples along the cutback surface to generate a spiral vortex which is guided to the lands on both sides thereof.

Abstract: Components for gas turbine engines are described. The components include an airfoil having a leading edge cavity with a baffle portion and a leading edge portion. A baffle is installed within the baffle portion and includes a first metering flow aperture. A first support element retention feature is located within the leading edge cavity. A first axial extending rib extends between an aft end of the cavity and a forward end proximate the first support element retention feature and is formed on an interior surface of the airfoil. A first axial extending flow channel extends along the first axial extending rib between an exterior surface of the baffle and an interior surface of the airfoil and the first metering flow aperture is located proximate the aft end of the first axial extending flow channel to generate a forward flowing cooling flow.

Abstract: A turbine engine includes a compressor section, a combustion section, a turbine section, and an exhaust section in serial flow order and together defining a core air flowpath; a fuel delivery system for providing a flow of fuel to the combustion section; and a waste heat recovery system. The waste heat recovery system includes a heat source exchanger in thermal communication with the turbine section, the exhaust section, or both; a heat sink exchanger in thermal communication with the fuel delivery system, the core air flowpath, or both; a thermal transfer bus including a thermal transfer fluid and extending from the heat source exchanger to the heat sink exchanger; and a pump in fluid communication with the thermal transfer bus downstream of the heat source exchanger and upstream of the heat sink exchanger for increasing a temperature and a pressure of the thermal transfer fluid in the thermal transfer bus.

Abstract: A turbine of a turbo engine including a casing and a turbine nozzle stage, the turbine nozzle stage including an outer metal shroud secured to the casing, an inner metal shroud, and a plurality of ring sectors made of ceramic-matrix composite material forming a crown extending between the outer metal shroud and the inner metal shroud and having an internal shroud and an external shroud, each ring sector having an internal platform forming a portion of the internal shroud, an external platform forming a portion of the external shroud, and at least one blade extending between the external platform and the internal platform and secured thereto. Each sector is fixed to the outer metal shroud using at least one assembly including a screw and a nut, the screw passing through the outer platform of the sector and the outer metal shroud.

Abstract: A fan blade anti-icing system comprises a fan hub and a fan blade extending radially outwardly from the fan hub. The fan blade has a base and an airfoil extending radially outwardly from the fan base. The airfoil having a leading edge, a trailing edge, a convex side surface between the leading and trailing edge and a concave side surface between the leading and trailing edge. The fan blade further has a radial passage extending from a blade air inlet in the blade base in communication with a source of heated air, and a rearwardly directed passage in communication with the radial passage and having a blade air outlet forward of the trailing edge and oriented tangentially to the convex side surface or concave side surface of the airfoil.

Abstract: A structure for cooling a turbine airfoil includes: a cooling passage formed between a first airfoil wall curved as to be concave relative to a high-temperature gas passage and a second airfoil wall curved so as to be convex relative to the high-temperature gas passage; lattice structure bodies each formed by stacking a plurality of ribs in a lattice pattern; a partition body provided between the adjacent lattice structure bodies; a cooling medium discharge port for discharging a cooling medium within the cooling passage to the outside; and an exposed wall portion formed as a portion of the second airfoil wall extending beyond the cooling medium discharge port to the outside. At outlet portions of the lattice structure bodies adjacent to each other across the partition body, adjacent first rib sets and second rib sets are inclined in opposite directions relative to the partition body, respectively.

Abstract: A sealing arrangement is provided between inner and outer coaxial ring-shaped walls of an aircraft engine. The outer wall has an axial end with a U-shaped cross-section parallel to axis, the opening therein being oriented axially in a first direction. The inner wall has an axial end with a U-shaped cross-section parallel to axis, the opening therein being oriented axially in the opposite direction. The U-shaped end of the outer wall includes a ring-shaped free edge axially engaged in the opening in the axial end of the inner wall. The U-shaped end of the inner wall includes a ring-shaped free edge axially engaged in the opening in the axial end of the outer wall. The walls define therebetween a ring-shaped gas channel having a substantially S-shaped axial cross-section.

Abstract: A flow-straightener vane of a bypass turbomachine includes a plurality of vane sections stacked radially with respect to a longitudinal axis (X) along a stacking line (L) between a root end and a tip end. Each vane section has a pressure-face surface and a suction-face surface extending axially between an upstream leading edge and a downstream trailing edge. Between the leading and trailing edges of each vane section there is formed a profile chord (CA) the length of which is substantially constant between the tip end and the root end, and the stacking line (L) exhibits a curvature in a plane passing more or less through the axis (X) and through the stacking line (L), situated in the vicinity of the tip end and oriented from downstream towards upstream.

Abstract: A turbine blade tip includes a tip cap disposed over a blade airfoil and having a pressure side edge and a suction side edge. A notch is formed by a radially inward step adjacent to the suction side edge of the tip cap. The notch is defined by a radially extending step wall and a radially outward facing land. The step wall extends radially inward from the suction side edge of the tip cap to the land, whereby the land is positioned radially inward in relation to a radially outer surface of the tip cap. The notch extends along at least a portion of the suction sidewall in a direction from the leading edge to the trailing edge. In a further aspect, a method is provided for servicing a blade that includes machining a suction side notch as described above.

Abstract: A method of manufacturing a turbine blade includes forming a blade body divided body constituting a blade body of the turbine blade by a three-dimensional lamination method; individually forming a plurality of shroud divided bodies constituting a shroud of the turbine blade for each of the shroud divided bodies by a three-dimensional lamination method; joining the shroud divided bodies together; and joining the blade body divided body and the shroud divided bodies.

Abstract: A turbine blade for a gas turbine engine has: an airfoil extending along a span from a base to a tip and along a chord from a leading edge to a trailing edge, the airfoil having a pressure side and a suction side, a tip pocket at the tip of the airfoil, the tip pocket at least partially surrounded by a peripheral tip wall defining a portion of the pressure and suction sides; at least one internal cooling passage in the airfoil and having at least one outlet communicating with the tip pocket; and a reinforcing bump located on the pressure side of the airfoil and protruding from a baseline surface of the peripheral tip wall to a bump end located into the tip pocket, the reinforcing bump overlapping a location where a curvature of a concave portion of the pressure side of the airfoil is maximal.

Abstract: Component for gas turbine engines are described. The components include an airfoil body having leading and trailing edges and pressure and suction sides. The airfoil has a leading edge cavity located proximate the leading edge defined between the leading edge and a separator rib and between the pressure side and the suction side. An insert member is installed within the leading edge cavity. The insert member has one or more metering flow apertures at an aft end and one or more impingement apertures at a forward end and at least one axially extending rib along an exterior surface thereof. At least one axial extending flow channel passage is defined along the axial extending rib between an exterior of the insert member and an interior of the airfoil body. Air flow through the metering flow apertures flows into the axial extending flow channel passage and flows forward toward the leading edge.

Abstract: An apparatus and method of cooling a hot portion of a gas turbine engine, such as a multi-stage compressor of a gas turbine engine, by reducing an operating air temperature in a space between a seal and a blade post of adjacent stages by routing cooling air through an inlet in the vane, passing the routed cooling air through the vane, and emitting the routed cooling air into the space.

Abstract: A nozzle for a turbine system includes an airfoil, an inner sidewall, and an outer sidewall. Each of the inner sidewall and outer sidewall each includes a peripheral edge defining a pressure side slash face, a suction side slash face, a leading edge face, and a trailing edge face. At least one of the inner sidewall pressure side slash face, the inner sidewall suction side slash face, the outer sidewall pressure side slash face, or the outer sidewall suction side slash face includes a first swept surface extending at a first angle relative to a nominal slash face angle and a second swept surface extending at a second angle relative to the nominal slash face angle. The first and second swept surfaces meet at a joining line that is circumferentially aligned with a stiffening member extending circumferentially on a respective sidewall.

Abstract: A vane arc segment includes an airfoil section, a spar, and a baffle. The spar supports the airfoil section and has a leg that extends in an internal cavity of the airfoil section. The leg is spaced from the airfoil wall such that there is a gap there between. The baffle divides the gap into a plenum space between the leg and the baffle and an impingement space between the baffle and the airfoil wall. The baffle has impingement holes directed toward the airfoil wall that connect the plenum space and the impingement space. The impingement space is of substantially uniform thickness and the plenum space is of non-uniform thickness along a portion of the span length of the airfoil section.

Abstract: A turbine blade for a gas turbine engine. The turbine blade having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.

Abstract: A method for manufacturing a turbine engine part in which a first rough casting element includes a first face and a second face opposite to each other is assembled by the second face on an orifice which has a second element of the part. The method includes machining a through-cavity in the first element which opens at the first face and from the second face of the first element and machining the first face of the first element so as to form an area suitable for ensuring the attachment of a conduit on the first element. The machining of the cavity and of the first face is achieved by using a machining reference frame based on the second element.

Abstract: A gas turbine engine article includes an article wall that defines leading and trailing ends and first and second sides that join the leading and trailing ends. The article wall defines a cavity. A cooling passage network is embedded in the article wall between inner and outer portions of the article wall. The cooling passage network has an inlet orifice through the inner portion of the article wall to receive cooling air from the cavity, a plurality of sub-passages that extend axially from the at least one inlet orifice, at least one outlet orifice through the outer portion of the airfoil wall, and trip strips for mixing cooling air in the cooling passage network.

Abstract: A turbine blade includes an airfoil portion, a cooling passage inside the airfoil portion, and a plurality of cooling holes formed in a trailing edge part of the airfoil portion. The cooling holes communicating with the cooling passage and opening in a surface of the trailing edge part. A relation of d_up<d_mid<d_down is satisfied, where d_mid is an index indicating opening densities of the cooling holes in a center region including an intermediate position between a first end and a second end of the airfoil portion in the blade height direction, d_up is an index in a region positioned upstream of a flow of a cooling medium in the cooling passage from the center region in the blade height direction, and d_down is an index in a region positioned downstream of the flow of the cooling medium from the center region in the blade height direction.

Abstract: A method for manufacturing a stage of a steam turbine comprising the steps of milling a block of material to define a sector having a plurality of blades, each blade having an external surface; machining an opening in the external surface of at least one of the blades; machining a cavity in fluid communication with the opening; the step of machining the cavity being performed by wire electric discharge machining.

Abstract: A gas turbine engine article includes a cooling passage network embedded in an article wall between inner and outer portions of the article wall. The network has an inlet orifice through the inner portion to receive cooling air from a cavity, a sub-passage region that includes an array of pedestals, and at least one outlet orifice through the outer portion. The array of pedestals includes first pedestals arranged in a first row and second pedestals arranged in a second, adjacent row. The first pedestals and the second pedestals define inter-row sub-passages there between. Each of the inter-row sub-passages has an inlet mouth, an outlet mouth, and a compound channel connecting the inlet mouth and the outlet mouth. The compound channel includes a first channel length over which the inter-row sub-passage has a constant cross-section and a second channel length over which the inter-row sub-passage has a non-constant cross-section.

Abstract: An aerofoil for a turbine blade, having a leading edge onto which a hot gas can flow and from which a suction-side wall and a pressure-side wall extend to a trailing edge of the aerofoil, the aerofoil extending in a direction transverse thereto from a root-side end having an aerofoil height of 0% to a tip-side end having an aerofoil height of 100%, and having at least two rows of cooling holes arranged along the leading edge, which cooling holes are at a mutual spacing which can be measured perpendicularly to the leading edge. In order to provide a turbine blade which can be used, with reduced cooling complexity, for additional reliable cooling of the leading edge in various operating conditions, the at least two rows of cooling holes are arranged at least in part in a wavy line along the leading edge.

Abstract: A stator vane includes an airfoil having an airfoil shape. The airfoil shape has a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in one of Table I, Table II, Table III, Table IV, Table V, Table VI, Table VII, Table VIII, Table IX, Table X, Table XI, or Table XII. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances expressed in a unit of distance by multiplying the Cartesian coordinate values of X, Y and Z by a scaling factor of the airfoil in the unit of distance. The X and Y values, when connected by smooth continuing arcs, define airfoil profile sections at each Z value. The airfoil profile sections at Z values are joined smoothly with one another to form a complete airfoil shape.

Abstract: A blade includes an airfoil portion including a pressure surface and a suction surface each extending between a leading edge and a trailing edge along a spanwise direction, and at least one communication hole extending m the airfoil portion and having a first and a second opening end opening to the pressure and suction surface, respectively. The first opening end is located on a first cross-section perpendicular to the spanwise direction at a first position in the spanwise direction, the second opening end is located on a second cross-section perpendicular to the spanwise direction at a second position in the spanwise direction, and a dimensionless blade chord length position of the first opening end with respect to the leading edge on the first cross-section is larger than a dimensionless blade chord length position of the second opening end with respect to the leading edge on the second cross-section.

Abstract: An engine component for a gas turbine engine includes a film-cooled substrate having a hot surface facing hot combustion gas and a cooling surface facing a cooling fluid flow. The substrate includes one or more film holes that have a metering section defining a metering diameter and a diffusing section that defines a hooded length.

Abstract: A vane arc segment includes an airfoil fairing that has first and second fairing platforms and a hollow airfoil section. A spar has a spar platform adjacent the first fairing platform and a hollow spar leg that extends from the spar platform and through the hollow airfoil section. The hollow spar leg has an internal passage for receiving cool air there through, a clevis mount, and a pin fairing. The clevis mount is distal from the spar platform and protrudes from the second fairing platform. The clevis mount includes first and second prongs with aligned holes. A pin extends through the aligned holes. The pin fairing extends over the pin between the first and second prongs for guiding the cooling air around the pin. There is a support platform adjacent the second fairing platform. The pin locks the support platform to the spar leg.

Abstract: The present invention relates to a panel for tip clearance control formed by: a first perforated sheet adapted for being seated on the turbine casing; a second sheet arranged on said first sheet and configured for being attached to the first sheet leaving a gap between both sheets; and a third sheet arranged between both first sheet and second sheet such that respective spaces in fluid connection by at least one hole are configured: a distribution chamber and an impingement chamber, both chambers extending from one end of the panel to the other. The panel further comprises a closure element together with a sealing element at one of its lateral ends for allowing the passage of fluid with another adjacent panel, whereas the sealing element is configured for allowing relative movement between both adjoining panels.

Abstract: A gas turbine engine is provided. The gas turbine engine includes an engine casing structure and a part retained relative to the engine casing structure by a channel-cooled hook. The channel-cooled hook includes at least a portion of a hook cooling channel. A vane assembly for the gas turbine engine is also provided.

Abstract: In order to limit the air flow required to cool its blade member, a blade for an aircraft turbomachine turbine includes, in its blade member, a main passage configured to collect a main air flow from an air inlet and to circulate it in the blade member, and a side passage located between the main passage and the pressure-side wall, while being separated from the main passage, wherein the side passage is configured to collect a side air flow from the air inlet and to circulate it in the blade member, the main passage having at least three portions connected together end-to-end to form a coil, the side passage being located between each of these portions and the pressure-side wall.

Abstract: An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges to provide an exterior airfoil surface. The pressure and suction side walls are spaced apart from one another in a thickness direction. A stagnation line is located near the leading edge. A cooling passage is provided between the pressure and suction side walls. The showerhead cooling holes are arranged at least one of adjacent to or on the stagnation line. At least one of the showerhead cooling holes has a metering hole fluidly connecting the cooling passage to a diffuser arranged at the exterior airfoil surface. At least one showerhead cooling hole is arranged on each of opposing sides of the stagnation line. Each showerhead cooling hole has the diffuser with a first diffuser angle that expands downstream in the thickness direction in opposing directions from one another when separated by the stagnation line.

Abstract: A heat exchanger comprises: a plurality of first fluid channels defining a flow path for a first fluid; a first fractal channel for conveying the first fluid to the plurality of first fluid channels; and a second fractal channel for conveying the first fluid from the plurality of first fluid channels; wherein the first fractal channel and the second fractal channel each comprises at least one divergence point along its length where a parent channel splits into a plurality of sub-channels which diverge away from each other, and wherein one or each of the first fractal channel and the second fractal channel defines one or more through-channels for a second fluid to pass through the respective fractal channel, each through-channel extending from a parent channel side of a divergence point and emerging between the corresponding sub-channels on a sub-channel side of the divergence point.

Abstract: A gas turbine includes a plurality of first-stage stator vanes arranged in a circumferential direction, the first-stage stator vanes each including a vane surface having a pressure surface and a suction surface, a first combustor disposed on a suction surface side of one of the first-stage stator vanes, the first combustor having a first combustor outlet which includes a first side wall portion extending along a radial direction, and a second combustor disposed on a pressure surface side of the one of the first-stage stator vanes and adjacent to the first combustor in the circumferential direction. The one of the first-stage stator vanes satisfies 0.05??y/P?0.25, where ?y is a protruding amount of the suction surface from the inner wall surface of the first side wall portion to the circumferential direction, and P is an arrangement pitch of the first-stage stator vanes in the circumferential direction.

Abstract: A turbine vane assembly adapted for use with a gas turbine engine includes an airfoil and a spar. The airfoil is formed to define a cavity that extends into the airfoil. The spar is located in the cavity to define a cooling passage that extends around the spar between the spar and the airfoil. The turbine vane assembly includes cooling features to aid heat transfer of the turbine vane assembly during operation in the gas turbine engine.

Abstract: An airfoil includes an airfoil body that has a trailing edge region. The trailing edge region includes first and second monolithic ceramic exterior walls, a flow discharge passage between the first and second monolithic ceramic exterior walls, a ceramic matrix composite (CMC) liner at least a portion of which is disposed in the flow discharge passage between the first and second monolithic ceramic exterior walls, and an array of pedestals disposed in the flow discharge passage. Each of the flow guides bridges the CMC liner and at least one of the first and second monolithic ceramic exterior walls.

Abstract: A turbine vane assembly adapted for use in a gas turbine engine includes an aerofoil configured to interact with gases flowing through the gas turbine engine along a gas path, an outer platform that defines an outer boundary of a gas path, and an inner platform that defines an inner boundary of the gas path.

Abstract: Airfoil assemblies for gas turbine engines are described. The airfoil assemblies include an airfoil body having a leading edge, a trailing edge, a pressure side, and a suction side, the airfoil body extending in a radial direction between a first end and a second end, wherein the airfoil defines an internal cavity bounded by interior surfaces of the airfoil body, the airfoil body formed from a high-temperature-material material and a metallic insert member installed within the internal cavity. One or more radially extending ribs are arranged on an exterior surface of the metallic insert member and defining one or more radially extending passages between the exterior surface of the metallic insert member and the interior surface of the airfoil body.

Abstract: A variable vane actuation system for a gas turbine engine includes a stem section that forms a base and a contoured section that extends from the base along an axis. A vane arm comprising a claw section received onto the contoured section and a fastener fastened to the contoured section to load the claw section to the base.

Abstract: A vane arc segment includes an airfoil piece that has a first platform, a second platform, and an airfoil section between the first platform and the second platform. At least the first platform is formed of a fiber-reinforced composite (FRC) and defines a radial flange. A cap is fitted on the radial flange.

Abstract: A tip shroud for a turbine blade of a gas turbine system includes a body coupled to a radial outer end of an airfoil of the turbine blade. The tip shroud may include at least one circumferentially extending tip rail. A first edge wall of the tip shroud extends axially and radially outwardly from the body along at least one of a leading circumferential-facing edge or a trailing circumferential-facing edge of the body and includes a circumferentially facing surface. Cooling passages are defined in the body and extend circumferentially therein to cool an area near the first edge wall. The tip shroud includes an exit surface adjacent the first edge wall, where the exit surface includes an exit opening through which at least one of cooling passages exits the body. The exit surface is angled relative to the circumferentially facing surface of the first edge wall in a range of 15° to 80°.

Abstract: An airfoil includes an airfoil section that has an airfoil wall that define leading and trailing ends and first and second sides that join the leading and trailing ends. The first and second sides span in a longitudinal direction between first and second ends. The airfoil wall circumscribes an internal core cavity. The airfoil section extends form a platform. The platform defines a shelf that extends forward from the first end at the leading end of the airfoil section to a platform leading edge, aft from the first end at the trailing end of the airfoil section to a platform trailing edge, laterally from the first end at the first side of the airfoil section to a first platform side edge, and laterally from the first end at the second side of the airfoil section to a second platform side edge. The internal core cavity extends from the airfoil section into the shelf.

Abstract: Example apparatus, systems, and articles of manufacture to control deflection mismatch are disclosed herein. Further examples and combinations thereof include: A deflection limiter comprising an inner shroud segment to support a stator structure, the inner shroud segment including a first end face and a first outer upper portion, the first end face positioned radially inward and aft relative to the first outer upper portion, and an outer shroud segment to support the inner shroud segment, the outer shroud segment including a second end face and a second outer upper portion, the second end face positioned aft relative to the first end face and the second outer upper portion positioned aft relative to the first outer upper portion of the inner shroud segment, the second end face coupled to the first end face of the inner shroud segment and the second outer upper portion coupled to the first outer upper portion.