Abstract: A turbine vane assembly for use in a gas turbine engine includes turbine vanes, a segmented inner vane support, and an outer vane support. The turbine vanes are arranged around a central axis of the engine. The inner vane support is arranged radially inwardly of the turbine vanes. The outer vane support includes a full-hoop outer support ring located radially outward of the plurality of turbine vanes and extends entirely circumferentially about the central axis. The outer vane support further includes a plurality of discrete support spars coupled to the full-hoop outer support ring and extending radially inward from the full-hoop outer support ring through an interior cavity of a respective turbine vane.

Abstract: An airfoil, such as a vane or a blade for a gas turbine engine, may be provided. The airfoil may include a platform; a spar; a fillet; and a coversheet. The spar may include a passageway inside of the spar for a cooling fluid, a pedestal on an outer surface of the spar, and a spar hole configured to direct the cooling fluid from the passageway to the outer surface of the spar. The fillet may be located at the intersection of the platform and the spar, and may couple the spar to the platform. The fillet and/or the coversheet may include a protrusion extending along the edge of the coversheet. The protrusion, along with the pedestal and the outer surface of the spar, may define a purge groove. The purge groove and a purge groove outlet may form a cooling path for cooling fluid to flow onto the platform.

Abstract: A component is provided and comprises at least one wall comprising a first and a second surface. At least one film cooling; hole extends through the wall between the first and second surfaces and has an outlet region at the second surface. The film cooling hole includes a first expansion section being a side diffusion portion and a second expansion section being a layback diffusion portion, wherein the side diffusion portion is upstream and spaced from the layback diffusion portion.

Abstract: An airfoil includes a leading edge, a trailing edge, a base, and a tip. The airfoil further includes a pressure side wall and a suction side wall that extend between the leading edge, the trailing edge, the base, and the tip. The airfoil further includes a plurality of passages that are defined within the airfoil and extend from an inlet at one of the base or the tip. Each passage of the plurality of passages is defined at least partially by a primary impingement wall and a solid side wall. The primary impingement wall is spaced apart from one of the pressure side wall or the suction side wall such that a primary impingement gap is defined therebetween. The primary impingement wall defines a plurality of impingement apertures that direct air in discrete jets across the impingement gap to impinge upon an interior surface of the airfoil.

Abstract: A molding tool (10) for manufacturing cooling features in a ceramic core for a casting process includes a first mold portion (12) comprising a crossover hole forming feature (18); a second mold portion (24) comprising an impingement jet receiving forming feature (30) for forming an impingement jet receiving feature having a desired aiming point (32); and a sacrificial alignment member (34) for extending at least partially into the crossover hole forming feature (18) at least partially into the aiming point (32) of the impingement jet receiving forming feature (30) for substantially aligning a central axis (38) of the crossover hole forming feature (18) with the aiming point (32) to maintain a crossover hole and aiming point alignment during casting to ensure that the jet is directed at the aiming point (32) in a resultant cast part.

Abstract: A turbomachine turbine blade, includes a platform, a vane, a cooling cavity supplying a plurality of cooling outlets provided along the trailing edge, two radially adjacent cooling outlets defining therebetween a rib. At least one cooling hole is formed in the thickness of at least one rib and/or in the thickness of a portion of the trailing edge fillet located in the axial extension of at least one rib, so as to ensure fluid communication for a cooling flow between the inside and the outside of the blade for cooling the at least one rib.

Abstract: A compressor includes a rotor including a plurality of disks, a shaft portion connected on a downstream side of the disks; and rotor blade rows fixed to the plurality of disks; a stator including a compressor casing; and a plurality of stator vane rows each provided between corresponding adjacent ones of the rotor blade rows; an outlet guide vane including blade main bodies disposed at an interval in a circumferential direction on the downstream side of the disk located most downstream, and inner shrouds connecting the blade main bodies in the circumferential direction, on an inner side in a radial direction; and an inner casing disposed on the downstream side of the disk located most downstream with a gap between the disk and the inner casing.

Abstract: A vane ring assembly adapted for use in a gas turbine engine includes an airfoil set, an outer band, and an inner band. The airfoil set extends radially across a primary flow path annulus arranged around an axis. The outer band extends around the airfoil set. The inner band extends around the axis at the airfoil set.

Abstract: A fluid supercharging device (3, 5), comprising: a rotating shaft (7); a vane disc (308) coaxially fixed to the rotating shaft (7); a plurality of fan blades (301) fixed around a perimeter of the vane disc (308); the back side of the fan blades 301 being provided with at least one fluid guiding inlet (305), an end of the back side distal from the vane disc (308) is provided with a fluid guiding outlet (306, 307), a fluid channel (304) communicating the fluid guiding inlet (305) with the fluid guiding outlet (306, 307) is provided along a lengthwise direction inside the fan blades; the fan blades (301) rotate to generate a centrifugal force such that a fluid flows into the fluid channel via the fluid guiding inlet on the back side, and flows out of the fluid guiding outlet along the lengthwise direction of the fan blades.

Abstract: A steam turbine hollow stationary blade is able to reduce the amount of water droplets captured on a blade surface. The steam turbine hollow stationary blade, which has a cavity therein, includes a partition wall dividing the cavity into a pressure chamber on a leading edge side and an exhaust chamber on a trailing edge side, at least one steam inlet hole connecting the pressure chamber and an outside of the stationary blade to each other, and at least one pressure conditioning hole connecting the pressure chamber and the exhaust chamber. Total opening area of the pressure conditioning hole is smaller than total opening area of the steam inlet hole.

Abstract: An apparatus is provided for a turbine engine. This turbine engine apparatus includes a turbine engine component that includes a sidewall and a cooling aperture. The cooling aperture includes an inlet, an outlet, a meter section, a diffuser section and a transition section between and fluidly coupled with the meter section and the diffuser section. The cooling aperture extends through the sidewall from the inlet to the outlet. The meter section is at the inlet. The diffuser section is at the outlet. The transition section is configured to accommodate lateral misalignment between the meter section and the diffuser section.

Abstract: A diffuser assembly includes a casing at a compressor aft end; an inner barrel member radially inward of the casing; and an array of radial flow splitters extending between the inner barrel member and the casing. Each radial flow splitter includes a leading edge facing into a flow of air, a trailing end wall opposite the leading edge, a pair of side walls extending between the leading edge and the trailing end wall, and an axis extending through the leading edge and the trailing end wall. A width of each radial flow splitter increases from the leading edge to the trailing end wall. The side walls diverge away from the axis in a downstream direction corresponding to the flow of air. Optionally, the side walls also diverge away from the axis in a radial direction between the inner barrel member and the casing.

Abstract: The film cooling structure includes a wall part extending forward and rearward, and a cooling hole including a tubular inner peripheral surface and inclined such that an outlet is positioned rearward of an inlet. The cooling hole includes a throat having a minimum cross section, and a diffuser part extending from the throat to the outlet. The diffuser part includes a channel cross section expanding rearward and along the wall part as it approaches the outlet. The inner peripheral surface of the cooling hole includes a flat portion extending in a direction perpendicular to the cooling hole and along the wall part at a front part of the inner peripheral surface, and a convex portion projecting from a rear part of the inner peripheral surface toward the flat portion, extending in parallel with the flat portion, and forming the throat between the flat portion and the convex portion.

Abstract: A cooling assembly includes a coolant chamber disposed inside an airfoil of a turbine assembly that directs coolant inside the airfoil. The airfoil extends between a leading edge and a trailing edge along an axial length of the airfoil. Inlet cooling channels are fluidly coupled with the coolant chamber and direct the coolant in a direction toward a trailing edge chamber of the airfoil. The trailing edge chamber is fluidly coupled with at least one inlet cooling channel. The trailing edge chamber is disposed at the trailing edge of the airfoil and includes an inner surface. The inlet cooling channels direct at least a portion of the coolant in a direction toward the inner surface of the trailing edge chamber. One or more outlet cooling channels direct at least a portion of the coolant in one or more directions away from the trailing edge chamber.

Abstract: A gas turbine engine defining an axial direction and a radial direction, and including a primary cooling circuit configured to receive a first flow of air; and a turbine rotor comprising a rotor blade, the rotor blade defining at least in part a forward wheelspace that is located forward of the rotor blade, the forward wheelspace configured to receive a second flow of air, the rotor blade further defining: a first cooling circuit internal to the rotor blade and in fluid communication with the primary cooling circuit for receiving the first flow of air from the primary cooling circuit; a second cooling circuit internal to the rotor blade and in fluid communication with the forward wheelspace for receiving a portion of the second flow of air from the forward wheelspace; and a means for drawing a portion of the second flow of air into the second cooling circuit.

Abstract: An example gas turbine engine includes a compressor including a casing defining a passageway and a shaft extending through the passageway. The shaft drivingly coupling the compressor and a turbine of the gas turbine engine. The shaft has an opening to receive airflow from the passageway. The gas turbine engine also includes an inner shroud, stator vanes coupled to and extending radially between the casing and the inner shroud, and an offtake scoop disposed on a downstream side of the inner shroud. The offtake scoop has a channel to direct the airflow radially inward toward the opening in the shaft.

Abstract: A method for machining a blade and a blade for a turbomachine comprising a shroud which is positioned on a tip side of the blade. The shroud has an outer surface with at least one circumferential fin arranged thereon, whereby at least one section of the outer surface beside the at least one fin is processed in at least two manufacturing steps. At least one first section of the outer surface is processed to have a first shape and at least one second section of the outer surface is processed to have a second shape.

Abstract: An integral cooling system for a turbine casing and guide vanes in an aeroengine is provided, belonging to the field of research on flow and heat exchange of a turbine casing in an aeroengine. An inner guide ring and multiple of guide vanes are mounted on the turbine casing; the cooling system includes an electromagnetic pump, a heat exchanger, an expansion joint and a cooling pipeline; an annular cavity is provided in the turbine casing, the cooling pipeline is mounted on the inner wall of the annular cavity and periodically and uniformly distributed along the circumferential direction of the turbine casing, and the cooling pipeline is filled with cooling liquid; a mounting cavity is further provided in the turbine casing, and the mounting cavity communicates with the annular cavity; the electromagnetic pump, the expansion joint and the heat exchanger are all mounted in the mounting cavity.

Abstract: There are provided a throttle mechanism and the like that are capable of easily changing a cross-sectional area of a flow path according to an operating state. The throttle mechanism in an embodiment is a throttle mechanism that controls a flow rate of a fluid flowing through a flow path, and is configured to make a cross-sectional area of the flow path change autonomously according to temperature.

Abstract: Flow diverters for installation in mid-turbine frame systems at a conduit outlet of gas turbine engines are described. The flow diverters include a diverter body having a connector portion defining a diverter inlet, a diverter extension at least partially defining a diverter outlet, and a curved portion arranged between the connector portion and the diverter extension, the curved portion configured to change a direction of flow from a first direction to a second direction that is about 90° from the first direction as the flow passes from the diverter inlet to the diverter outlet.

Abstract: An inlet guide vane assembly for an aircraft engine includes an array of inlet guide vanes having radially inner ends at a radially inner shroud, radially outer ends at a radially outer shroud, and airfoils extending therebetween. An internal passage extends radially through the airfoil from a vane air inlet at the radially inner end to a vane air outlet at the radially outer end. The vane air inlet is in fluid communication with an inner plenum, disposed radially inwardly of the inlet guide vane and in fluid communication with an anti-icing air source. The vane air outlet in fluid communication with an outer plenum, disposed radially outwardly of the inlet guide vane and having an exhaust port. A vane anti-icing pathway is defined in a radially-outward direction from the inner plenum, through the internal passage of the vane, and into the outer plenum.

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 turbine rotor in an embodiment is configured by joining a rotor component member and a rotor component member together by bolt fastening with an abutting end surface of the rotor component member and an abutting end surface of the rotor component member abutting on each other. The turbine rotor includes: a cylindrical recessed portion that is formed at the abutting end surface and is recessed in an axial direction; an axial passage bored from a bottom surface of the cylindrical recessed portion in the axial direction; an introduction passage introducing the cooling medium into the axial passage; a discharge passage discharging the cooling medium from the axial passage; and a sealing member that is arranged in the cylindrical recessed portion and seals one end of the axial passage.

Abstract: Various methods and systems are provided for a fluidic variable turbine. In one example, a system for an engine comprises a turbocharger turbine including a nozzle ring, the nozzle ring including a plurality of stationary vanes, each vane of the plurality of stationary vanes including a plurality of injection ports arranged at an outer surface of the vane, and a gas supply system to supply variable gas flow to and out of the plurality of injection ports.

Abstract: A method of making an impeller includes building the impeller in a layer by layer process in a build direction along the rotational axis starting from a base of the hub. The plurality of blades includes a plurality of perforated blades that support the shroud during additively manufacturing the impeller. The method can include installing the impeller in a fuel pump, air compressor, or the like, without removing the perforated blades from the impeller.

Abstract: An aeration apparatus for aerating water discharged from a hydraulic turbine includes: a manifold disposed within a crown of a runner of the hydraulic turbine; a plurality of radial pipes extending radially from an outer perimeter of the manifold and in fluid communication with the manifold; and one or more air injectors having a first end disposed within an aeration pipe, each of the one or more air injectors having a second end extending into a nozzle at a first end of one of the radial pipes. Rotation of the aeration apparatus resulting from rotation of the runner causes pumping of water from the manifold through the radial pipes past the one or more air injectors, and water flowing past the one or more air injectors causes air to become entrained in the water. The radial pipes discharge the water and entrained air from the aeration apparatus.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter. The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module including a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the fan assembly. Each heat transfer element may be individually and independently fluidly isolated from the remaining heat transfer elements.

Abstract: A turbine blade includes an airfoil body, and a plurality of cooling passages extending along a blade height direction inside the airfoil body and being in communication with each other to define a serpentine flow passage. The plurality of cooling passages include first turbulators on an inner wall surface of an upstream cooling passage of the plurality of cooling passages, and second turbulators on an inner wall surface of a downstream cooling passage of the plurality of cooling passages. A second angle formed by the second turbulators with respect to a flow direction of a cooling fluid in the downstream cooling passage is smaller than a first angle formed by the first turbulators with respect to the flow direction of the cooling fluid in the upstream cooling passage.

Abstract: An endwall may be disposed at one end of a vane assembly. The endwall may comprise an endwall spar and a coversheet on the hot surface of the endwall spar. The endwall may further comprise a cooling fluid channel between the hot surface of the endwall spar and the cold surface of the coversheet. The cooling fluid channel may include a cooling fluid inlet disposed in the endwall spar, and a cooling fluid outlet. The cooling fluid outlet may be formed at an angle with the axis of the endwall spar. A plurality of pedestals may be disposed on the hot surface of the endwall spar extending into the cooling channel. The pedestals may be formed at an angle with the axis of the endwall spar to direct a cooling fluid.

Abstract: Airfoil for gas turbine engines are described. The airfoils have internal walls and internal cavities. A leading edge cavity is defined within the airfoil body along a leading edge and a leading edge feed cavity is arranged aft of the leading edge cavity. A bent leading edge rib is arranged between the leading edge cavity and the leading edge feed cavity. A main body cavity is arranged aft of the leading edge feed cavity and defined at least in part by two interior ribs that define a part of the leading edge feed cavity. The main body cavity is fluidly connected to the leading edge feed cavity by an interior fluid connection through the intersection of the two interior ribs. A shield cavity is arranged to thermally shield the leading edge feed cavity from heat pickup along the suction side of the airfoil body.

Abstract: The invention concerns a gas turbine having a picture frame, a first vane, and a sealing arrangement to seal a gap between the picture frame and the first vane, the sealing arrangement including two seals arranged in series between the picture frame and the first vane. In exemplary embodiments, one of the seals is a honeycomb seal, a dogbone seal, a hula seal or a piston seal and the other seal is a honeycomb seal, a dogbone seal, a hula seal or a piston seal. A method of supplying cooling fluid to the gap between the picture frame and the first vane is also disclosed.

Abstract: A turbomachine defines an axial direction, a radial direction perpendicular to the axial direction, and a circumferential direction extending concentrically around the axial direction. The turbomachine includes a nozzle having an inner platform, an outer platform, and an airfoil. The airfoil includes a leading edge, a trailing edge downstream of the leading edge, a pressure side surface, and a suction side surface opposite the pressure side surface. The trailing edge is orthogonal with the outer platform in an axial-radial plane and the trailing edge is oblique to the inner platform in the axial-radial plane.

Abstract: An airfoil of either of a turbine blade or a turbine vane includes a cooling passage; at least one disk body disposed on an inner wall of the cooling passage and configured to reduce a flow cross-sectional area of the cooling passage to increase a fluid pressure of cooling fluid flowing through the cooling passage; and at least one through-hole formed in each of the at least one disk body such that the cooling fluid flows through the at least one through-hole and forms a vortex on a downstream side of the at least one through-hole. The cooling passage includes an inlet supplied with the cooling fluid and an end opposite to the inlet, and the at least one disk body is disposed at the inlet of the cooling passage and is configured to increase the fluid pressure of the cooling fluid flowing into the cooling passage.

Abstract: A gas turbine engine includes at least one trap that absorbs or adheres to calcium-magnesium-alumino-silicate (CMAS) entrained in intake air entering the engine.

Abstract: Systems and methods for air injection passageway integration and optimization in turbomachinery using surface vortex generation. An airfoil including a leading edge, a trailing edge, a pressure side, and a suction side, and is configured to influence an airflow as it passes from the leading edge to the trailing edge. The airfoil defines an aerodynamic passageway having an inlet on the pressure side and an outlet on the suction side to deliver air from the airflow through the airfoil to the suction side. The outlets are configured to inject the air at areas on either airfoil side targeted due to their propensity to generate undesirable boundary layer growth and associated flow losses. Outlet may also be included in the hub and the shroud of the turbomachine.

Abstract: An airfoil assembly includes first and second fiber-reinforced composite airfoil rings that each have inner and outer platform sections, a suction side wall extending between the inner and outer platforms, a pressure side wall extending between the inner and outer platforms, and suction and pressure side mate faces along, respectively, edges of the suction and pressure side walls. The suction side mate face of the first fiber-reinforced composite airfoil ring and the pressure side mate face of the second fiber-reinforced composite airfoil ring mate at an interface to form an airfoil that circumscribes an internal cavity. A least one of the suction or pressure side mate faces includes protrusions along a trailing edge of the airfoil. The protrusions define a toothed exit slot for emitting cooling air from the internal cavity.

Abstract: A stator of a turbine section, has: vanes distributed around a central axis, a vane of the vanes extending along a spanwise axis and defining an internal passage; an insert received within the internal passage, the insert defining a cavity for receiving cooling air and defining impingement cooling apertures facing an inner face of the vane; a splitter plate secured within the cavity and being transverse to the spanwise axis, the splitter plate having a base secured to the insert and a tip protruding from the base; and a flow passage defined between the tip and the insert, the flow passage fluidly connecting a first section of the cavity to a second section of the cavity, the tip of the splitter plate secured to the insert at at least one location along a perimeter of the tip.

Abstract: The gas turbine engine includes a casing assembly located proximate a turbine section of the gas turbine engine, and a tangential on-board injector (TOBI) having a body defining a plurality of air passages extending in a radial direction, the plurality of air passages circumferentially distributed and directing cooling air toward a turbine rotor of the turbine section of the gas turbine engine. An interference fit is provided between a face of the body and a face of a member of the casing assembly, the interference fit defining a fastener-free engagement between the bearing housing and the TOBI to prevent relative movement between the member of the casing and the TOBI.

Abstract: An airfoil including: an airfoil portion having an airfoil surface; and a communication hole extending at least in the airfoil portion and a first opening open in the airfoil surface, through which the first opening is communicated with a second opening provided in a portion of the airfoil. On a cross-section perpendicular to the spanwise direction through a position of the first opening of the spanwise direction, an angle A1 satisfying a condition (a) exists within an angle range ?10 degrees to 10 degrees with respect to an extension line obtained by extending a camber line of the airfoil portion from a leading edge. The condition (a) is a static pressure at a position of the first opening is equal to a static pressure at a position of the second opening when the airfoil portion receives a fluid flow from a direction of the angle A1 toward the leading edge.

Abstract: An internal core profile for a turbine nozzle airfoil of a gas turbine is provided. The turbine nozzle may include an airfoil core having an uncoated nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, wherein the X, Y, and Z coordinates are distances in inches measured in a Cartesian coordinate system, the corresponding X and Y coordinates, when connected by a smooth continuous arc, define one of a plurality of airfoil core profile sections at each Z distance, and the plurality of airfoil core profile sections, when joined together by smooth continuous arcs, define an airfoil core shape.

Abstract: A turbine assembly for use with a gas turbine engine includes a bladed wheel assembly and a vane assembly. The bladed wheel assembly is adapted to interact with gases flowing through a gas path of the gas turbine engine. The vane assembly is located upstream of the bladed wheel assembly and adapted to direct the gases at the bladed wheel assembly.

Abstract: Gas turbine engines and rotor arms thereof are described. The gas turbine engines include a first disk, a second disk, and a rotor arm arranged between and connecting the first disk to the second disk, wherein a cavity is defined at least between the rotor arm and the first disk. The rotor arm includes a radial portion having an inner diameter end and an outer diameter end, an axial portion having a first end and a second end, wherein the first end of the axial portion is connected to the outer diameter end of the radial portion, at least one entrance flow path defined within the radial portion extending from the inner diameter end to the outer diameter end, and at least one exit aperture arranged proximate the second end of the axial portion.

Abstract: A flowpath component for a gas turbine engine includes a body having a leading edge and a trailing edge. A first exterior wall connects the leading edge to the trailing edge and a second exterior wall connects the leading edge to the trailing edge. At least one first internal cooling passage has a first inlet at a first end of the body. At least one second internal cooling passage has a second inlet at a second end of the body. The at least one first internal cooling passage is isolated from the at least one second internal cooling passage.

Abstract: Disclosed herein are a ring segment having an air pouch and a first cooling hole formed therein, and a turbine including the same. The air pouch and the first cooling hole are formed in a shield wall, thereby achieving an improvement in cooling performance as well as simplification of production process.

Abstract: A turbofan engine includes an outer bypass duct and a gas turbine engine having an outer casing. The gas turbine engine is disposed in the outer bypass duct such that a bypass airflow passage is formed between the outer casing of the gas turbine engine and the outer bypass duct. The turbofan engine includes a turbine rear frame link assembly including a set of links coupled between the outer bypass duct and the outer casing to support the gas turbine engine relative to the outer bypass duct. The links are arranged around the gas turbine engine on a plane that is perpendicular to a centerline axis of the turbofan engine. None of the links extends through the bypass airflow passage at a position that intersects a radius extending in a vertically downward direction from the centerline axis.

Abstract: A cooling structure for a turbine airfoil includes: a lattice structure body formed such that a first rib set and a second rib set arranged in a cooling passage are stacked on each other in a lattice pattern; and lattice communication portions that allow passages formed between ribs of the first rib set to communicate with passages formed between ribs of the second rib set. Each of the first and second rib sets has rib walls each including a pair of ribs that are inclined in directions opposite to each other relative to an imaginary boundary line extending in a movement direction of a cooling medium and that are in contact with each other on the imaginary boundary line. A plurality of lattice communication portions are formed between two lattice communication portions at opposite end portions of each rib that forms the rib wall.

Abstract: A gas turbine engine includes a compressor section, a combustor, and a turbine section. The turbine section includes a high pressure turbine comprising a plurality of turbine blades. The gas turbine engine includes a tap for tapping air that is compressed by the compressor, to be passed through a heat exchanger to cool the air, the cooled air to be passed to the plurality of turbine blades. A sensor is located downstream of a leading edge of the combustor, and is configured to measure a characteristic of the cooled air. A controller is configured to compare the measured characteristic to a threshold and control an operating condition of the gas turbine engine based on the comparison.

Abstract: A turbine rotor in an embodiment includes: a rotor body portion; and a plurality of turbine disks provided on the rotor body portion in a center axis direction of the rotor body portion. The turbine rotor includes: a high-pressure cooling passage formed in the rotor body portion, the high-pressure cooling passage to which a high-pressure cooling medium is supplied, and the high-pressure cooling passage that discharges the high-pressure cooling medium to the high-pressure side turbine stage; and a low-pressure cooling passage formed in the rotor body portion, the low-pressure cooling passage to which a low-pressure cooling medium whose pressure is lower than the pressure of the high-pressure cooling medium is supplied, and the low-pressure cooling passage that discharges the low-pressure cooling medium to the low-pressure side turbine stage.

Abstract: A turbine airfoil includes a body including a wall defining pressure and suction sides, and a leading edge extending between the pressure and suction sides. A cooling circuit inside the wall of the body includes at least one of: a) a suction side to pressure side cooling sub-circuit including a first cooling passage(s) extending from the suction side to the pressure side around the leading edge to a first plenum, and a plurality of first film cooling holes communicating with the first plenum and extending through the wall on the pressure side; and b) a pressure side to suction side cooling sub-circuit including second cooling passage(s) extending from the pressure side to the suction side around the leading edge to a second plenum, and a plurality of second film cooling holes communicating with the second plenum and extending through the wall on the suction side.

Abstract: A component, such as for a turbine engine, can include an airfoil with an outer wall defining an exterior surface bounding an interior and defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip to define a span-wise direction. The component can also include at least one cooling passage within the interior.