Abstract: A cooling system for a gas turbine engine turbine section includes a rotor supporting a blade having a cooling passage. A disc is secured relative to the rotor and it forms a cavity between the rotor and the disc. A bleed air source is in fluid communication with the cavity. An impeller is arranged in the cavity. The impeller is configured to increase a fluid pressure within the cavity to drive bleed air from the bleed air source and thereby provide a pressurized cooling fluid to the cooling passage.

Abstract: A platform cooling device directs a cooling fluid onto a root side of a platform of a turbomachine part. The platform cooling device includes a first segment to be positioned at a root of the turbomachine part and a second segment, at an angle to the first segment, to be positioned at the root side of the platform of the turbomachine part. The second segment may include at least one impingement channel having an inlet for receiving at least a part of the cooling fluid and an outlet for releasing the received cooling fluid onto the root side of the platform. The first segment and the second segment may define a path for the cooling fluid via the impingement channel. A turbomachine component includes the platform cooling device.

Abstract: A turbine airfoil includes: a concave pressure sidewall and a convex suction sidewall joined together at a leading edge and a trailing edge, and extending between a root and a tip; an internal rib extending between the pressure sidewall and the suction sidewall; and a crossover hole formed in the rib, the crossover hole having a noncircular cross-sectional shape with a major axis defining a maximum dimension of the cross-sectional shape; wherein the major axis of the crossover hole lies in plane with the rib and is non-parallel to an imaginary curvilinear lateral centerline which defines a locus of points lying halfway between the pressure and suction sidewalls. The orientation of the crossover holes minimizes stress concentration caused by the presence of the crossover holes.

Abstract: In order to maximize cooling while minimizing drag in aerial vehicles of airborne wind turbines, it may be preferable to dissipate the cooling energy of the motors via a radiator in a region with advantageous airflow parameters. Aerial vehicle rotors operating in thrust mode may produce relatively more airflow velocity in certain regions further away from the center of the rotor blades, both radially and longitudinally. Placing a radiator in a rotor-supporting pylon and offset from the center of the rotor blades and aft of the rotor blades may allow for greater cooling while the aerial vehicle while in thrust mode.

Abstract: A rotor blade includes a mounting portion comprising a mounting body that is formed to receive a coolant therein. An airfoil portion extends substantially radially outwardly from the mounting body and includes an airfoil body. The mounting body and the airfoil body define a plurality of primary cooling passages that extend substantially radially therein for routing the coolant through the airfoil body. Each of the primary cooling passages includes a cooling flow outlet that is formed along a tip portion of the airfoil body. The airfoil body further defines a plurality of trailing edge cooling passages, each having a coolant outlet that is formed along a trailing edge portion of the airfoil body. At least a portion of the trailing edge cooling passages are formed along a radially outer portion of the trailing edge proximate to the tip portion of the airfoil body.

Abstract: A fan impeller for the conveyance of a gaseous fluid includes several blades arranged with constant angular spacing around a rotation axis. At least one blade has a boundary zone and an inner zone, together with a bead extending between them along a curved path and/or the boundary zone is determined by a first blade surface and the inner zone is determined by a second blade surface, which are arranged with varying geometry.

Abstract: A system for forming a cooling hole in an airfoil includes a liquid-jet guided laser. The liquid-jet guided laser generates a laser beam confined within a fluid column. The fluid column/laser beam is directed at an outer surface of the airfoil. The system also includes a purge medium supply that is fluidly coupled to an aperture of the airfoil. The purge medium supply provides a purge medium into an inner cavity of the airfoil. The purge medium flow is oriented to flow in a flow direction that is substantially parallel to an inner surface of the cavity. A cooling hole is formed in the airfoil and extends through the outer surface and penetrates the inner surface of the cavity. A centerline of the cooling hole forms an acute angle with respect to the purge medium flow direction. The system described herein provides a method for manufacturing an airfoil.

Abstract: Various embodiments of the invention include turbine buckets and systems employing such buckets. Various particular embodiments include a turbine bucket having: a base including: a casing having at least one exhaust aperture on an outer surface of the casing; and a core within the casing, the core having: a serpentine cooling passage; and at least one outlet passage fluidly connected with the serpentine cooling passage and the exhaust aperture, wherein the at least one outlet passage permits flow of a coolant from the serpentine cooling passage to the at least one exhaust aperture on the outer surface of the casing; and an airfoil connected with the base at a first end of the airfoil, the airfoil including: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side.

Abstract: A compressor rotor is provided having a rotor blade groove thereon and also includes a device for cooling the rotor in the region of a compressor rotor exit. Efficient cooling is achieved by the compressor rotor, in a compressor rotor exit region, having a ring which is pushed concentrically, and at a distance, forming a gap, over a rotor disk of the rotor, and is fastened on the disk, by the rotor blades, in the compressor rotor exit region, being inserted into corresponding grooves on the ring and being retained there, by first means for directing an axial flow of cooling medium from the compressor rotor exit through the ring, and by second means for deflecting the cooling medium which issues from the ring such that the cooling medium flows back in the axial direction through the gap between the ring and the rotor disk, encompassed by the ring.

Abstract: A damper pin for coupling platforms of adjacent turbine blades includes a first flat longitudinal end region, a second flat longitudinal end region and a reduced cross sectional area. The reduced cross sectional area is separated from the first flat longitudinal end region by a first main body region and the reduced cross sectional area is separated from the second flat longitudinal end region by a second main body region. The cross sectional area of the reduced cross sectional area is less than the cross sectional area of each of the first and second main body regions.

Abstract: The present application provides a cover plate assembly for use with a rotor disk. The cover plate assembly may include a radial flange extending from the rotor disk, a flange aperture extending through the radial flange, a cover plate segment with a fastening aperture and a hook for receiving the radial flange, and a fastener extending through the flange aperture and the fastening aperture.

Abstract: In a turbine for aeronautic engines, a first and at least a second turbine disk rotating around a common axis respectively carrying a first and a second moving-blade crown, with which first and, respectively, second axial passages are defined through which a cooling air mass for the turbine disks can pass, a device for conveying cooling air being interposed between the first and the second axial passages to receive the mass of air passing through the first axial passages and send it through the second axial passages.

Abstract: A bladed rotor of a gas turbine for aeronautic engines is cooled by making an air mass flow through a plurality of passages, each one defined by a turbine disk on one side and by a root of a corresponding blade coupled to this turbine disk on the other, the incoming air of each passage being subdivided into two or more airflows and each flow being directed towards the turbine disk in a predefined direction.

Abstract: A system includes a cover segment set configured to mount in first and second grooves circumferentially along a blade mounting region of a turbine rotor. The cover segment set includes a seal segment and a cover segment. The seal segment has a seal body with a lock feature, wherein the seal segment is configured to mount circumferentially along the second groove. The cover segment includes first and second lips extending along first and second circumferential portions of a cover body. The first and second lips are at an offset from one another. The first lip is configured to mount circumferentially along the first groove, the second lip is configured to mount circumferentially along the second groove adjacent the seal segment, and the second lip includes a mating lock feature configured to couple with the lock feature of the seal body.

Abstract: A gas turbine engine preswirler (470) includes an outer ring (471) and an inner ring (472). The inner ring (472) includes a plurality of angled holes (490). Each angled hole (490) follows a vector which is angled in at least one plane. A component of the vector is located on a plane perpendicular a radial from an axis of the preswirler (470). The component of the angle is angled relative to an axial direction of the preswirler (470).

Abstract: A wind turbine has a nacelle wherein a component is equipped with a first temperature sensor for detecting a temperature of the component, a second temperature sensor for detecting an outside temperature inside or outside the nacelle, a heater associated with a component to heat the latter, and a heater controller. The detected temperature of the component and the detected outside temperature are provided as input signals to the heater controller which activates the heater for heating when the detected temperature of the component is below a predetermined switch-on value for the heater. The switch-on value for the heater is, in at least one predetermined temperature interval, dependent on the detected outside temperature. The switch-on value for a first outside temperature is greater than the switch-on value for a second outside temperature in the temperature interval when the second outside temperature is greater than the first outside temperature.

Abstract: A rotor disk for a gas turbine engine includes a hub; a rim with a forward portion and an aft portion; and a plurality of circumferentially spaced radial scallops in the rim to form cooling slots between two disks when the rotor disk is connected to another rotor disk.

Abstract: According to one embodiment, a propeller blade includes a carbon foam core and a structural layer that surrounds at least a portion of the carbon foam core.

Abstract: Provided is a turbine blade including: a platform; a blade body including a cooling flow channel including a meandering serpentine cooling flow channel; a fillet portion provided in a joint surface between the blade body and the platform; and a base portion including a cooling flow channel communicated with the serpentine cooling flow channel. The cooling flow channel includes a bypass flow channel that is branched off from a high-pressure part of the cooling flow channel, passes through along the inside of the fillet portion, and is connected to a low-pressure part of the cooling flow channel.

Abstract: A turbine blade rotor assembly provides sealing for the rotor components and bi-direction axially locking of the blades on the rotor using a seal plate on the aft exhaust side without axial locking on the forward intake side of the rotor and blade array. As a result, the blades can be removed from the aft direction without moving a turbine casing.

Abstract: A gas turbine engine configured to rotate in a circumferential direction about an axis extending through a center of the gas turbine engine comprises a turbine stage. The turbine stage comprises a disk, a plurality of blades and a mini-disk. The disk comprises an outer diameter edge having slots, an inner diameter bore surrounding the axis, a forward face, and an aft face. The plurality of blades is coupled to the slots. The mini-disk is coupled to the aft face of the rotor to define a cooling plenum therebetween in order to direct cooling air to the slots. In one embodiment of the invention, the cooling plenum is connected to a radially inner compressor bleed air inlet through all rotating components so that cooling air passes against the inner diameter bore.

Abstract: The present application provides a near flow path seal for a gas turbine. The near flow path seal includes a base, a pair of arms extending from the base, and a curved indentation positioned between the pair of arms.

Abstract: The present invention proposes a rotor device for a jet engine with a disk wheel and several rotor blades connected to said disk wheel, with the rotor blades being arranged in each case via a blade root substantially in the axial direction inside recesses of the disk wheel. Several locking segments are provided in the recesses of the disk wheel for axially locking the rotor blades, said locking segments interacting on the one hand with grooves of the rotor blades and on the other hand with at least one groove of the disk wheel. In the area of the groove of at least one of the rotor blades between the groove of the rotor blade and a locking segment there is a positive fit preventing a relative movement between the rotor blades and the locking segments.

Abstract: A rotorcraft is disclosed. The rotorcraft may include an airframe, at least one engine connected to the airframe, and a rotor connected to the airframe. The rotor may include a hub, a rotor blade, and a feathering spindle connecting the rotor blade to the hub. The rotorcraft may further include a flow of oil passing proximate the feathering spindle. The flow of oil may cool the feathering spindle during take off and landing of the rotorcraft. Additionally, the flow of oil may heat the feathering spindle during travel of the rotorcraft at altitude.

Abstract: The present invention relates to a wind turbine nacelle having a top face with a longitudinal extension in a wind direction. The nacelle comprises a cooling device having a cooling area and extending from the first face of the nacelle, and a cover having at least one inner face and at least a front edge facing the wind direction. The cooling device is enclosed by the first face of the nacelle and the inner face of the cover and is arranged in a front distance of at least 440 mm from the front edge of the cover.

Abstract: A turbine rotor body having at least one inlet orifice in fluid communication with a pre-swirl system such that the inlet orifice receives cooling fluids from the pre-swirl system is disclosed. The inlet orifice may be configured to reduce the relative velocity loss associated with cooling fluids entering the inlet orifice in the rotor, thereby availing the cooling system to the efficiencies inherent in pre-swirling the cooling fluids to a velocity that is greater than a rotational velocity of the turbine rotor body. As such, the system is capable of taking advantage of the additional temperature and work benefits associated with using the pre-swirled cooling fluids having a rotational speed greater than the turbine rotor body.

Abstract: The present invention relates to a wind turbine nacelle having a first face with a longitudinal extension in a wind direction, comprising a cooling device having a cooling area and extending from the first face of the nacelle, and a cover having at least one inner face, the cooling device being enclosed by the first face of the nacelle and the inner face of the cover. A first distance between at least one of the faces and the cooling area is at least 30 mm.

Abstract: The present invention relates to a wind turbine nacelle having a top face with a longitudinal extension in a wind direction, comprising a cooling device extending from the top face of the nacelle and a cover having at least one inner face. The cooling device is enclosed by the top face of the nacelle and the inner face of the cover.

Abstract: The invention relates to a cooled wall segment in the hot gas path of a gas turbine, particularly to a cooled stator heat shield. Such components have to be properly cooled in order to avoid thermal damages of these components and to ensure a sufficient lifetime. The wall segment according to the invention includes a first surface, exposed to a medium of relatively high temperature, a second surface, exposed to a medium of relatively low temperature, and side surfaces connecting said first and said second surface and defining a height of the wall segment. At least one cooling channel for a flow-through of a fluid cooling medium extends through the wall segment. Each cooling channel is provided with an inlet for the cooling medium and an outlet for the cooling medium.

Abstract: A brush configured to engage a rotating component to transfer an electrical signal and/or power is provided including a body formed from a core material. A coating material is disposed over a portion of the body configured to contact the rotating component. A contact area between the brush and the rotating component increases as the coating material wears.

Abstract: A turbine rotor blade is provided with a blade root, platform adjoining it, and turbine blade on that side of the platform which faces away from the blade root, with at least one opening for feeding coolant into the turbine rotor blade interior on an underside of the blade root, which opening merges into a coolant duct. An axial rotor section for a rotor is provided, having an outer circumferential surface adjoining two end-side first side surfaces with rotor blade holding grooves distributed over the circumference and extending along an axial direction, wherein a turbine rotor blade is arranged in every holding groove, wherein a multiplicity of sealing elements are at the side of a side surface of the rotor section, and lie opposite the end sides of blade roots to form a gap. Multiple outlet holes for impingement cooling of the sealing elements are provided in the end surface.

Abstract: A rotating inlet cowl for a turbine engine, having a rotation axis and for which the forward end is arranged to be eccentric relative to this rotation axis. Furthermore, a forward cone of the cowl is truncated by a truncation surface defining the forward end of the inlet cowl.

Abstract: Multi-scale turbulation features, including first turbulators (46, 48) on a cooling surface (44), and smaller turbulators (52, 54, 58, 62) on the first turbulators. The first turbulators may be formed between larger turbulators (50). The first turbulators may be alternating ridges (46) and valleys (48). The smaller turbulators may be concave surface features such as dimples (62) and grooves (54), and/or convex surface features such as bumps (58) and smaller ridges (52). An embodiment with convex turbulators (52, 58) in the valleys (48) and concave turbulators (54, 62) on the ridges (46) increases the cooling surface area, reduces boundary layer separation, avoids coolant shadowing and stagnation, and reduces component mass.

Abstract: A turbine bucket is disclosed herein. The turbine bucket may include a platform and a shank portion extending radially inward from the platform. The shank portion may include a slash face, a radial seal pin groove formed in the slash face, and at least one cooling hole disposed in the slash face about the radial seal pin groove.

Abstract: Through the use of different ceramic heat insulation coatings (7, 13) on turbine blades, it is possible to produce or restore different configurations of gas turbines, which are then optimized for a respective baseload operation application or peak load operation application. Differences in the sprayed-on heat insulation coatings (7, 13) can be generated by means of a different material composition of the powder, for example zirconium oxide, by means of a different number of layers (10), by means of different coating thicknesses and/or by means of different porosities.

Abstract: One embodiment of the present application includes a hot section component of a gas turbine engine having a covering. The covering includes a protrusion and is attached to the hot section component though a flexible retainer. In one form the covering is made from ceramic matrix composite. The flexible retainer has a closed position and an open position. The retainer secures the protrusion to the hot section component when it engages part of the protrusion when in the closed position.

Abstract: A turbine wheel for a turbine engine, comprising a disk carrying blades, each having a platform carrying an airfoil and connected by a tang to a root, and sealing and damping sheets housed in the inter-tang spaces, the platforms including projections on their radially internal faces against which the sheets bear radially in operation, in order to define radial clearance and create at least one space between the sheets and the platforms, and the sheets including holes for feeding air to the or each space.

Abstract: A film cooling hole for an air cooled turbine airfoil, where the film cooling hole includes a first expansion section with expansion only on the downstream wall and a second expansion section with expansion on the downstream wall and the two side walls. No expansion is formed on the upstream walls on the first and second expansion sections.

Abstract: A system and method to cool the air inside the nacelle and the heat generating components, particularly of an offshore wind turbine is presented. The ambient air first enters an air handling unit near the tower bottom where the airborne water droplets and salt particles are removed. The clean air is then compressed adiabatically, thus increasing the dew point temperature of the water vapor in the air. The high pressure, high temperature air from the compressor is then cooled and dehumidified in a sea water-to-air heater exchanger or an air-to-air heat exchanger. The high pressure air from the heat exchanger then enters the turbine at the tower bottom and flows up to the nacelle where it is allowed to expand adiabatically in a duct. The duct helps direct the resulting cold air over the heat generating components. The cold air can also used to cool these components internally. The warm air ultimately exits the nacelle at the rear top.

Abstract: A turbine blade includes a core element having a base portion, a tip portion, and an intermediate portion extending between the base portion and the tip portion. The intermediate portion includes a non-uniform cross-section and is a high-strength fiber material. The turbine blade further includes a shell disposed around the core element, and the volume between the core element and the shell forms a void.

Abstract: A connector for connecting an impeller to a shaft is provided. The impeller has a shaft-side hub extension with a central recess. The impeller is formed of a material having a greater coefficient of thermal expansion than the material of the shaft. The connector is inserted into the recess to frictionally connect an outwardly facing surface of the connector with a radially inner surface of the hub extension. The connector has a threaded portion carrying a thread which screws onto a corresponding threaded portion of the shaft, such that the connector provides a rotationally fixed connection between the impeller and the shaft. The connector is formed of a material having a coefficient of thermal expansion which is greater than the coefficient of thermal expansion of the material of the shaft.

Abstract: A method for casting a collar (44, 46) for a heat shield (36) of a strut (32) in a gas turbine exhaust section (20). A casting geometry (60, 70) is defined with extra wall thickness (56, 68) in an area of wall curvature (53, 54), which provides a flow path beyond a final geometry of the collar to facilitate a flow of molten metal in the mold (63, 64). The extra thickness is removed after casting, leaving the collar in its final geometry, which may have uniform wall thickness (T, T2). The extra thickness in the casting geometry may be provided by increased radius (R3) in the wall curvature and/or by casting feed portals (66, 68) that span the wall curvature between a tubular portion (50) and a flange (52) of the collar.

Abstract: A separator device is provided for separating dirt particles from a flow of cooling air fed to airfoils of the turbine section of a gas turbine engine. In use the separator device extends across a conduit which bypasses the combustor of the engine to convey pressurized cooling air carrying dirt particles from the compressor section of the engine to openings which direct the air into the airfoils. The separator device is configured to direct a first portion of the impinging cooling air flow away from the openings and to allow a second portion of the impinging cooling air to continue to the openings. The first portion of cooling air has a higher concentration of the coarsest dirt particles carried by the cooling air than the second portion of cooling air.

Abstract: A spacer for a gas turbine engine includes a rotor ring defined along an axis of rotation and a plurality of core gas path seals which extend from the rotor ring, each of the plurality of core gas path seals extend from the rotor ring at an interface, the interface defined along a spoke.

Abstract: An exhaust-gas turbocharger (1) having a shaft (2), a turbine wheel (5), which is fastened to the shaft (2), and a heat throttle (8) between the shaft (2) and the turbine wheel. An end face (3) of the shaft (2) is provided with a protrusion (4), with an outside diameter (A4) which is smaller than the outside diameter (A2) of the shaft (2). The turbine wheel (5) has a hollow receiving portion (7), which is formed integrally on the wheel rear side (6) and the inside diameter (I7) corresponds to the outside diameter (A4) of the protrusion (4) and the outside diameter (A7) corresponds to the outside diameter (A2) of the shaft (2). The protrusion (4) engages into the receiving portion (7). The heat throttle (8) is formed by a cavity (8A, 8B), which has an outside diameter (A8) which is smaller than the outside diameter (A4) of the protrusion (4) and extends from the protrusion (4) into the receiving portion (7).

Abstract: The invention relates to a control device (1) of the pitch of blades of a rotor of a propeller, comprising: a radial shaft (6), a pivot (8) connected to the blade, the rotation of the radial shaft (6) driving the rotation of the pivot (8) for the modification of the pitch of the blade, characterised in that it comprises: an insert (12) fixed rigidly to the pivot (8) so as to block their relative displacement, the radial shaft (6) being configured to drive in rotation the insert (12), at least one peg (24) passing through the insert (12) and the pivot (8) for transmission of the rotation of the insert (12) to the pivot (8), a heat insulation element (15) arranged between the insert (12) and the pivot (8), whereof the thermal conductivity is less than the thermal conductivity of the pivot (8) and of the radial shaft (6). The invention also relates to a propeller comprising this device.

Abstract: A turbine arrangement is provided, particularly a gas turbine arrangement, having at least one rotor blade and a turbine disc, the rotor blade having a root portion, the turbine disc having at least one slot in which the root portion of the rotor blade is secured. The slot has a plurality of opposite pairs of slot lobes and a plurality of opposite pairs of slot fillets, and a slot bottom of the slot. The slot bottom is arranged to have a first convex surface section. Furthermore the root portion of the rotor blade has a root bottom with a first concave surface section corresponding to the first convex surface section of the slot bottom. Additionally, the first convex surface section is pierced by an outlet of a cooling duct through the turbine disc.

Abstract: A turbine component is provided. The turbine component includes an airfoil having a first surface and a second surface. A thermal barrier coating is coupled to the second surface, wherein the thermal barrier coating includes a first portion, a second portion and a trench defined between the first and second portions. A channel is coupled in flow communication to the first surface and the trench, wherein the channel includes a first sidewall and a second sidewall opposite of the first sidewall. The first and second sidewalls extend from the first surface and toward the trench at an angle. The turbine component includes a cover coupled to the second surface, wherein the cover includes a first end coupled to the first portion and a second end extending into the trench and spaced from the second portion.

Abstract: A cooling system of a wind turbine is provided which includes a generator (10), a housing portion (1), and a hub (2) which is rotatable with respect to the housing portion (1) and drivingly connected to the generator (10). The hub (2) supports at least one blade (4). The housing portion (1) and the hub (2) enclose a cavity for accommodating at least said generator. Heat exchange means (6) are provided on the outer surface of said hub (2) for transferring at least a part of heat produced by said generator (10) in operation to the outside.

Abstract: A wind turbine blade having an elongated blade body extending along a longitudinal axis and having an upper skin and a lower skin, the lower skin spaced from the upper skin in a thickness direction of the blade body, the upper skin and/or lower skin having a laminated layer, the laminated layer having an outer layer wherein the outer layer forms part of the upper and/or lower skin respectively, an inner layer spaced from the outer layer in the thickness direction; and an intermediate layer sandwiched between the outer layer and inner layer, the intermediate layer having a plurality of openings extending through the intermediate layer from the inner layer to the outer layer; and a plurality of corresponding heat conductor elements extending through the plurality of openings from the inner layer to the outer layer for transferring heat from the inner layer to the outer layer.