Abstract: A gas turbine blade comprises a base metal, a platform, an airfoil extending upwardly from the platform, a shank extending downwardly from the platform. The shank has an exterior wall and an internal passage, and the airfoil has a cooling flow channel inside the airfoil for flowing a cooling flow therethrough. The blade has a first chromide coating contacting the base metal of at least a portion of an interior surface of the shank and interdiffused therewith, wherein the first chromide coating does not have an aluminide coating deposited over it. The blade has a second chromide coating contacting the base metal of at least a portion of an interior surface of the airfoil and interdiffused therewith. A method for preparing a gas turbine blade comprises the steps of applying chromide coatings, sealing the interior passages of the shank and airfoil and applying an aluminide or platinum aluminide coating and an optional ceramic layer on the airfoil.

Abstract: An internal heating arrangement for a centrifugal impeller for a gas turbine engine is provided having at least one heating passage extending through into the rotor for directing air bled from the rotor exit along the backface and forwardly through the impeller.

Abstract: In a gas turbine, annular overhang portions are formed on adjacent surfaces of a plurality of rotor discs so as to face each other, surrounding the rotor axis, groove portions are provided circumferentially to the surfaces of the overhang portions facing each other, and sealing structures are annularly installed to the inside of the groove portions. The gas turbine comprises sealing plate assemblies which include the overhang portions, the groove portions and a plurality of plates being formed annularly by being mutually piled up; and detachable retaining members which are provided so as to fix the overhang portions and the sealing plate assemblies together by way of disc engagement portions provided to the overhang portions and sealing plate engagement portions provided to the sealing plate assemblies.

Abstract: A turbine blade having a cooling supply passage formed within the blade, and a blade tip with a squealer tip formed thereon. The blade tip includes a blade top or cap, and includes a plurality of mini-serpentine cooling channels formed within the tip. The mini-serpentine channels can be 2-pass, 3-pass, 4-pass, or 5-pass serpentine channels, and each includes an inlet hole connected to the internal cooling supply passage to pass cooling air through the channels. Each channel includes an exit hole with a diffuser that opens onto the pressure side of the blade to provide film cooling. The mini-serpentine cooling channels can be arranged to flow substantially from blade side to side or from blade edge to edge.

Abstract: A turbine blade with a film cooling holes arrangement to supply film cooling to hot sections on the pressure side and suction side of the blade. The cooling flow circuit for the blade uses a serpentine flow cooling path to supply cooling air to a number of impingement cavities, the impingement cavities being connected to film cooling holes to discharge film cooling air to key spots on the pressure side wall and suction side wall of the blade that require film cooling air. The serpentine flow cooling channels are located on the opposite side of the blade from the impingement cavity in which the channel communicates with through metering and impingement holes. The first and second legs of the serpentine flow cooling channels are located on the pressure side, and the third leg or channel is located on the suction side of the blade downstream from the gage point and where no further film cooling is preferred.

Abstract: An apparatus for distributing heated air to an environment comprises a ceiling fan assembly secured to a heating assembly. The ceiling fan assembly comprises a plurality of rotatable blades arranged substantially in a horizontal plane to drive airflow in a desired direction when the blades are rotated. The heating assembly comprises a heating element and a power supply for energizing the heating element to produce heat, whereby air driven by the ceiling fan contacts the heating element and passes into the environment. A lighting assembly and a guard may be included.

Abstract: A passageway is provided through which a cooling fluid flows in a first direction. The passageway has a plurality of trip strips positioned within the passageway. Adjacent one of the trip strips are oriented to converge towards each other at a first end to form an apex portion and to form a region in which turbulence is created. The apex portion is oriented at an angle with respect to the first direction.

Abstract: In accordance with the present invention, there is provided a turbine engine component having an airfoil portion with a pressure side and a suction side, a first microcircuit embedded in a wall forming the pressure side, an internal cavity containing a supply of cooling fluid, the first microcircuit having an inlet leg, an intermediate leg, and an outlet leg, and a plurality of communication holes between the internal cavity and the outlet leg. In a preferred embodiment, the outlet leg is provided with at least one set of features for locally accelerating cooling flow in the outlet leg and for increasing heat pick-up ability.

Abstract: A turbine airfoil includes a leading edge having a concave cooling flow passage. An apex of the concave cooling flow passage divides the flow passage into adjacent regions. The turbine airfoil includes a first plurality of turbulators disposed in one of the adjacent regions, and a second plurality of turbulators disposed in the other of the adjacent regions. The first and second pluralities of turbulators are positioned relative to one another to divert cooling flow in opposing swirl streams that recombine along the apex and to effect a desired heat transfer and pressure loss.

Abstract: A cooled airfoil includes an impingement rib having a multiple of openings which supply a cooling airflow from a cooling circuit flow path toward an airfoil leading edge. The multiple of openings are offset in the impingement rib opposite an outer airfoil wall which includes gill holes. Offsetting the multiple of openings opposite an outer airfoil wall which includes the gill holes focuses the cooling airflow across turbulators to increase the cooling airflow dwell time to increase the thermal transfer therefrom in higher temperature airfoil areas.

Abstract: A turbine airfoil with a plurality of breakout passages located just beneath a thermal barrier coating or just beneath the metal surface of the airfoil. The breakout passages are connected to an internal cooling air passage and allow for film cooling air to flow when a surface over the breakout passage has been chipped or eroded away to provide for additional cooling to the damaged airfoil surface. The breakout passages are formed during the casting process of the airfoil using a plurality of molding pieces. The breakout passages are formed along a direction of the pulling direction of the respective mold piece. This allows for the solidified airfoil to be easily removed from the mold assembly.

Abstract: Described is a cooled gas turbine guide blade for a gas turbine, with a hollow blade profile which comprises an onflow edge onto which a working medium is capable of flowing, and with, for guiding a cooling medium, an onflow edge duct running inside the blade profile along the onflow edge. Furthermore, to the use of such a gas turbine guide blade and to a method for operating a gas turbine with an abovementioned gas turbine guide blade. In order to provide a gas turbine guide blade with an increased lifetime by means of the invention, it is proposed that, in the onflow edge duct, an electrical heating element be provided, which extends approximately completely over the entire length of the onflow edge duct and through which a heating current flows, before the operation of the gas turbine, for the preheating of the gas turbine guide blade.

Abstract: A blade is provided for a turbine engine that includes an exterior surface. The exterior surface includes a portion having a thermal barrier coating and an uncoated shelf adjacent to the thermal barrier coating without the thermal barrier coating. A cooling hole extends from an internal passageway through the exterior surface to an exit. A scarfed channel is recessed in the exterior surface and interconnected to the cooling hole at the exit. The scarfed channel extends to a blade tip end surface. The scarfed channel protects the cooling fluid exiting the cooling hole from secondary flows surrounding the blade that would otherwise mix with and disperse the cooling fluid. The scarfed channels also increase the surface area exposed to the cooling fluid to increase the heat transfer rate.

Abstract: A method of repairing an airfoil for a gas turbine engine includes removing a damaged portion of the airfoil at a radially outward tip of the airfoil, securing temporarily a portion of a ceramic core stub within an internal cavity of the airfoil, applying new metallic material to the airfoil covering an exposed portion of the ceramic core stub, machining the airfoil to remove an excess portion of the new metallic material, and removing the ceramic core stub from the internal cavity of the airfoil.

Abstract: The invention relates to a turbine blade comprising a vane that runs along a blade axis and a platform region, which is located at the root of the vane and has a platform that extends transversally to the blade axis. The aim of the invention is to guarantee the improved cooling of a platform region and the transition region from a turbine blade to a platform of a turbine blade, thus ensuring the cooling of the delimitation of a flow channel in a gas turbine. To achieve this, the platform comprises a first platform wall that does not support the vane and a second platform wall that supports said vane. According to the inventive concept, at the root of the vane and over the course of the transition region from the turbine blade to the platform, said first platform wall is aerodynamically curved and the course of the second platform wall has a receding shoulder in relation to the first platform wall, as a continuation of the vane.

Abstract: A rotor of a thermally loaded turbomachine, in particular of a compressor or a gas turbine, is mounted such that it can rotate about a rotor axis and is concentrically surrounded by a hot-gas duct or cooling-air duct. A significant improvement to the ability of the rotor to withstand thermal loads is achieved, with little increased outlay in terms of materials, by virtue of the fact that the rotor comprises a rotor core made of a first material, that the rotor core is concentrically surrounded by shielding rings made of a second material, which shield the rotor core from the temperature in the hot-gas duct or cooling-air duct, the second material having a higher heat resistance than the first material, and that the shielding rings are cohesively joined to the rotor core.

Abstract: The invention relates to a device for cooling the slots of a turbomachine rotor disk, the device comprising a rotor disk having a plurality of slots and a flange that extends upstream from a radially upstream face of the disk. The device also comprises a retaining annulus having a flange that is placed around the flange of the disk, co-operating therewith to provide a space forming a cooling air diffusion cavity, said cavity opening out into the bottom of each of the slots of the disk, and a plurality of air admission orifices opening out into the diffusion cavity at the upstream end thereof, the cavity having means at its downstream end for reducing the flow rate of air penetrating into the slots of the disk that are situated radially in register with air admission orifices.

Abstract: The present invention provides near-surface cooled airfoils that can be made with near-surface cooling passages that are completely free of any leachable or otherwise sacrificial material in the recessed portion of the outer surface of the core. The turbine airfoil comprises a metallic core or substrate having an outer surface and one or a plurality of recessed portions of the outer surface; an intermediate metallic skin or foil having a back surface and a top surface, the back surface of the intermediate skin being bonded to the outer surface of the core such that the recessed portion(s) is sufficiently enclosed so as to form at least one or more near-surface cooling passages or pathways; and at least one or more metallic coatings of a high temperature-resistant metallic material deposited on a top surface of the intermediate skin.

Abstract: A turbine blade having a pressure sidewall and a suction sidewall connected at chordally spaced leading and trailing edges to define a cooling cavity. Pressure and suction side inner walls extend radially within the cooling cavity and define pressure and suction side near wall chambers. A plurality of mid-chord channels extend radially from a radially intermediate location on the blade to a tip passage at the blade tip for connecting the pressure side and suction side near wall chambers in fluid communication with the tip passage. In addition, radially extending leading edge and trailing edge flow channels are located adjacent to the leading and trailing edges, respectively, and cooling fluid flows in a triple-pass serpentine path as it flows through the leading edge flow channel, the near wall chambers and the trailing edge flow channel.

Abstract: A nose cone of a turbomachine, such as an aircraft engine, is heated to prevent ice formation. The nose cone is configured as a conical shell positioned concentrically over an inner shell with a space therebetween. Spiral spacers in the space provide distributed passages through which heated lubricating oil flows. A system for separating entrained air from the lubricating oil as part of this mechanism is disclosed.

Abstract: A gas turbine combustor liner, including a shell having a first end adjacent to an upstream end of the combustor and a second end adjacent to a downstream end of the combustor, where the shell also has a hot side, a cold side, and a centerline axis therethrough. A plurality of small, closely-spaced film cooling holes are formed in the shell through which air flows for providing a cooling film along the hot side of the shell. Each cooling hole has a non-uniform diameter as it extends through the shell. In particular, each cooling hole includes a first opening located at the cold side of the shell having a first diameter and a second opening located at the hot side of the shell having a second diameter, wherein the second diameter of the second opening is larger than the first diameter of the first opening. It is preferred that the shape of each cooling hole be substantially frusto-conical.

Abstract: A device for de-icing an aircraft wall that comprises a composite structure (47) and layer (41) of electrically conductive material, which device comprises a plurality of thin inductor elements (33 to 35) designed to extend substantially parallel to the layer of conductive material at a distance therefrom that is small enough to ensure that when the inductor is powered by an AC source of appropriate frequency, the layer or skin heats rapidly and uniformly, without significantly heating the composite structure.

Abstract: A method of forming a rotor for a turbine engine such that the rotor is formed of two materials including: an inner disk formed from a first material, such as steel, and an outer ring formed from a second material, such as a nickel alloy, having a larger thermal expansion coefficient than the first material forming the inner disk. The ring may include an inner aperture having a conical shape, and the disk may have an outer surface with a conical shape and a diameter with a portion that is larger than a portion of the ring. The ring may be heated such that the aperture expands to a size greater than the largest diameter of the inner disk. The ring may be positioned over the disk and allowed to cool to allow the ring to be attached to the disk. The ring and disk may then be co-forged.

Abstract: The purpose of the invention is to protect, in a horizontal axis wind turbine with an exhaust aperture for exhausting air from within the nacelle, the exhaust aperture and the internal structures from storm winds through facing the exhaust aperture away from the storm wind, even when storm winds are received from all directions relative to the nacelle. A rotary duct 4 is attached to the nacelle 1 so as to be able to rotate freely in the yaw direction, and is rotated in the yaw direction in accordance with changes in the wind direction. The exhaust aperture 4a is provided at a position that changes direction depending on the rotation of the rotary duct 4 in the yaw direction. By rotating the rotary duct 4 in the yaw direction so that the exhaust aperture 4a will face the downwind side in accordance with changes in the wind direction, the exhaust aperture 4a can be faced away from storm winds, even when storm winds are received from all directions in relation to the nacelle 1.

Abstract: A cooling system includes a fan assembly that is rotated to move air in a primary path through a housing. The fan assembly includes a hub and a plurality of spaced reinforcing ribs that establish channels between a motor casing and the interior of the hub. A heat sink assembly is axially spaced from a hub bottom to define an air cavity that is in communication with a gap. The hub includes a tube that defines an air passage to establish communication with the cavity. A plurality of air vanes extend in a spiral path from the hub bottom and into the air cavity to move air through the air passage and create turbulence over the heat sink assembly. The air vanes move the air over the heat sink assembly and through the air passage to establish communication between the channels and the air cavity.

Abstract: According to some embodiments, an apparatus may comprise a cold plate comprising a plurality of fins and an impeller comprising a plurality of blades. An output fluid velocity angle defined by the plurality of fins may be aligned to an impeller inlet velocity angle. The impeller inlet velocity angle may be based on an operational speed of the plurality of blades and an angle of the plurality of blades with respect to a center of the impeller.

Abstract: A method for manufacturing a turbine engine component includes forming a component to include a first side and an opposite second side, positioning at least one capillary adjacent to an external surface of at least one of the first and second sides, and securing the at least one capillary to the component with at least one composite layer.

Abstract: A fan includes an impeller, a motor, a plurality of first engaging members and a plurality of second engaging members. The impeller includes a hub and a plurality of blades disposed around the hub. The motor includes a rotor housing coupled with the hub, and the motor is for driving the impeller to rotate. The first engaging members are disposed on an inner side of the top surface of the hub, and the second engaging members are disposed on the top surface of the housing. When the impeller and the rotor housing are assembled, the second engaging members are disposed corresponding to the first engaging members, so that parts of the first engaging members are engaged into and assembled with the second engaging members.

Abstract: A turbine engine component has an airfoil portion having a pressure side and a suction side, a trailing edge discharge slot, and a suction side lip downstream of an exit of the trailing edge slot. The suction side lip is provided with negative features for increasing local heat transfer coefficient in the region of the suction side lip.

Abstract: A wind turbine blade that includes integrated erosion, lightning and icing protection for a leading edge. The leading edge includes a frontal surface with an erosion shield positioned external to the frontal surface. An internal lightning conveyance path is provided with an electrical connection from the erosion shield to the internal lightning conveyance path. A heat-generating element is positioned between the erosion shield and the frontal surface and a heat conduction path is provided from the heat-generating element to the erosion shield. Further, a structure is provided for connecting the heat-generating element to an energy source within the blade.

Abstract: A cooling system for an airfoil portion of a turbine engine component is provided. The cooling system includes a first cavity dedicated to cooling a trailing edge portion of an airfoil portion and a second cavity dedicated to cooling an aft portion of a pressure side wall.

Abstract: A compressor with a centripetal air takeoff system that includes pipes installed between an outer collar and an inner collar. Each pipe is fitted in a hole in the outer collar and in a hole in the inner collar. The system further includes a locking mechanism between the pipe and the inner collar.

Abstract: An ice management system is configured for an aircraft having at least one rotor (11), each rotor having a plurality of blades. The system has an alternating current (AC) electrical power supply for supplying AC power (13), the AC power supply being adapted to be carried by the aircraft remote from the rotor (11). A converter converts the AC power to direct-current (DC) power. At least one ice protection device is carried by the rotor and operated using the DC power.

Abstract: A compressor with a centripetal air takeoff system that includes pipes installed between an outer collar and an inner collar. Each pipe is fitted in a hole in the outer collar and in a hole in the inner collar. The system further includes a locking mechanism between the pipe and the inner collar.

Abstract: A turbine blade for turbo-engines as well as a method for manufacturing such a turbine blade is disclosed. According to the task set, the turbine blades should be capable of withstanding high thermal stress and able to maintain an adequate mechanical strength even at raised operating temperatures. The turbine blades are so designed that on the surface of a core element a heat-insulating layer of a metallic open-cell foam is integrally connected to said core element by sintering. The outer contour of the turbine blade is formed with at least one shell element. The shell element comprises a nickel-base alloy, which is also integrally connected by sintering to the open-cell foam which forms the heat-insulating layer.

Abstract: A turbine blade arrangement comprises turbine blades which are secured in adjacent positions to a rotor disc with a cavity defined between root segments and platform segments. A flow deflector within the cavity is provided as an insert such that through a recessed portion coolant flow .from a coolant path is constrained to remain adjacent to a rim surface. By constraining the coolant flow to remain adjacent to the surface greater cooling efficiency is achieved. Inner surfaces of the deflector may also be coated with low emissivity materials to reduce radiant heat flux transfer. The flow deflector supports a damper member in association with the platform segments.

Abstract: A component such as a turbine blade of a gas turbine engine has a cooling arrangement comprising a cascade impingement array in which cooling air flows from a supply chamber through impingement passages in webs to an internal passageway comprising first and second limbs. The cooling air flow through the limbs, provides improved heat transfer compared with continued impingement cooling in the chordwise direction of the blade.

Abstract: A component, such as a blade, vane or combustor wall of a gas turbine engine comprises two walls defining a coolant passage and has an array of pedestals extending between the two walls for heat removal. Each pedestal changes in cross-section along its length. Alternate rows of pedestals are arranged such that their larger cross-sectional area is adjacent one wall then the other. When a coolant flows through the passage it is forced to flow between one wall and the other wall so as to increase turbulence and hence mixing for a more even coolant temperature. The array of pedestals can also be used to tailor the individual heat loads on each wall independently and has the ability to use differing levels of blockage to counter adverse pressure gradients along successive rows of pedestals.

Abstract: The invention relates to a turbine 1, in particular a gas turbine, which along a swivel-mounted rotationally symmetrical rotor 2 has a compressor 3, a combustion chamber 5 and a turbine section 7 formed of a plurality of turbine stages 15, in which each turbine stage 15 contains mutually interacting blades 17 and vanes 16 which a hot working fluid 13 can flow around, with a coolant provided by the compressor 3 for cooling the blades 16, 17, which can flow in a channel along the rotor 2 from the compressor 3 to the turbine section 7 and into which a liquid can be introduced for cooling. In order to specify a turbine in which less wear occurs and in which the lifetime of the components is increased, it is proposed that the channel extends outside the rotor 2 and that the liquid can be introduced at the end of the channel which faces the compressor 2.

Abstract: A cooling device for a wind turbine is provided, the cooling device including at lease one liquid-tight tube which is partially filled with a liquid coolant, wherein one end of the tube is located in a first area and another end of the tube is located in a second area, the first and second areas having different temperatures when operating the wind turbine.

Abstract: The present invention is an integral composite structural (ICS) material comprising an open metal structure having at least one external side and internal surfaces defining a plurality of open shapes with a ceramic matrix composite bonded to at least one external side and the surfaces of at least a substantial portion of the plurality of open shapes and occupying at least a substantial portion of the plurality of open shapes. The open metal structure, independent of the ceramic matrix composite, has a total metal volume percent in the range of about 10% to about 90%, with no dimension of any open shape being greater than about ¾ inch. The ceramic matrix layer covers a substantial portion of at least one external side of the open metal structure.

Abstract: The device and method are used for heating a central section of a rotor mounted for rotation in a gas turbine engine.

Abstract: A method for deicing a rotor blade having a blade root, a blade tip, and a leading edge in which the rotor blade is operably coupled to a hub of a turbine includes circulating heated air through an outflow channel from the blade root towards the blade tip, recirculating the heated air via a return channel from the blade tip to the blade root, whereupon the recirculated heated air becomes returned air, and reheating the returned air for further circulation.

Abstract: A cooling arrangement comprises a support (20) for a plurality of blades (22). The support (20) comprises a plurality of blade mounting members (28) provided between adjacent blades (22). The blades (22) are mounted upon the blade mounting members (28). The cooling arrangement defines a pathway (36) for a cooling fluid. The cooling arrangement further includes a fluid directing formation (40, 44, 44A) to direct the cooling fluid across the blade mounting member (28).

Abstract: A ceiling fan that incorporates into its blades both a system for air purification by removing dust and other particulate pollutants from the air in a room through ionization of the air stream passing over an ionizer grid and also a system for heating and cooling the air stream passing over the blades. The device provides a room conditioner for efficiently heating or cooling and maintaining a room at a temperature comfortable to the user. The device also provides negative ion generation through the use of the ionizer.

Abstract: A turbine blade having improved cooling characteristics is provided. The turbine blade has an airfoil portion having a span, and at least one cooling passageway in the airfoil portion extending from a root portion of the airfoil portion to a tip portion of the airfoil portion. A plurality of turbulation promotion devices are placed in the at least one cooling passageway. The turbulation promotion devices have a P/e ratio which varies along the span of the airfoil portion, where P is the pitch between adjacent turbulation promotion devices and e is the height of the turbulation promotion devices.

Abstract: A turbine (1) with a rotationally symmetrical rotor (4) on rotary bearings which has on its outer perimeter moving blade retaining grooves running transverse to the outer perimeter and set at the same angular distance from one another, having a moving blade (13) that has a vane (24) and a moving blade foot (19) which corresponds to the moving blade retaining groove (18) and which is inserted in one of the moving blade retaining grooves (18), wherein a platform (23) is located between the moving blade foot (19) and the vane (24) and extends transverse to the moving blade foot (19), having a baffle cooling element (22) which is located between two directly adjacent moving blades (13) between its platforms (23) and the rotor (4) and through which a cooling fluid can flow to cool the platform (23).

Abstract: A rotor system for a gas turbine engine is provided having an outer rotor which includes an annular outer spool that comprises at least one ring portion of increased thickness with respect to the remainder of the outer spool, attached to a relatively thinner, axially extending arm portion. A plurality of outer turbine blades are attached to the ring portion of the outer spool and extend radially inwardly into a working gas flowpath of the engine. A casing surrounds the outer spool and defines an annular cavity between the casing and the outer spool, which is sealed by forward and aft brush seals at the forward and aft ends thereof. At least one inlet port is formed through the casing for admitting a flow of cooling air to the annular cavity. Means are provided for directing this cooling air flow to the ring portion of the outer spool.

Abstract: An airfoil having extended life due to reduction in stresses. The stresses are reduced by extending the thermal barrier coating below the radius between the outer band perimeter and the inner flow path surfaces and tapering the coating thickness. This additional tapered thermal barrier coating reduces the temperature gradient across a region already having high mechanical stresses resulting from geometric considerations thereby lowering thermally-induced stresses so that low cycle fatigue life is improved.

Abstract: A turbine nozzle includes a plurality of vanes joined at opposite ends to outer and inner bands. The inner band has a forward hook which is segmented to reduce thermal mismatch. And, in additional embodiments the vane includes an impingement baffle having preferential cooling.