Abstract: A low pressure turbine rotor disk for a small twin spool gas turbine engine in which the rotor disk includes a forward side cavity large enough to allow for the bearing assembly that rotatably supports the rotor disk to fit within the cavity in order to shorten the axial distance between the bearings that support the inner rotor shaft on which the turbine rotor disk is secured. Minimizing the bearings spacing allows for a high critical speed for the inner rotor shaft and therefore allows for the small twin spool gas turbine engine to operate at this small scale. The turbine rotor disk also includes a plurality of axial aligned cooling air holes to allow for cooling air from the bearings to flow out from the aft end of the rotor disk. The inner surface of the cavity is an annular surface that forms a seal with knife edges extending outward from the bearing support plate also located within the cavity.

Abstract: Cooling of turbine blades within a gas turbine engine is important. Coolant flows are taken from the engine to provide cooling effects but diminish the efficiency of the engine. Blades rotate and therefore centrifugal effects stimulate flow and pressure to maintain coolant flow presentation upon the blade. More cooling effectiveness is required towards the root of a blade in comparison with the tip. By providing cavities which incorporate return apertures coolant flow can be recycled. The cavities incorporate return portions on one side of a feed passage and a constriction is provided in passage. Thus, a proportion of coolant within the cavities is returned to the passage with pressure maintained by the rotational and centrifugal effects upon the coolant flow through the feed passage. Coolant flow is presented through outlet apertures as a film upon a surface of a blade.

Abstract: A rotor (10) for a power plant, in particular an aero engine or a gas turbine, whereby the rotor (10) includes at least two bonded rotor discs (12, 14). The rotor (10) has an inlet opening (30) for inlet of fluid from around the rotor (10) into a chamber (26, 44) in one or between two adjacent rotor discs (12, 14) and at least one interrupter element (38) with at least one fluid passage (40) is arranged in the chamber (26).

Abstract: Embodiments of the invention can provide systems, methods, and an apparatus for thermally isolating a turbine rotor wheel. According to one embodiment, a method for thermally isolating a turbine rotor wheel can be provided. The turbine rotor wheel can include a pair of blade posts, or shanks, supporting respective turbine rotor blades such that a shank cavity is formed between respective blade posts. To thermally isolate the turbine rotor wheel, at least one seal operable to partition the shank cavity into an upper region and a lower region can be provided. In addition, this seal can be inserted within the shank cavity to partition the shank cavity into an upper region and a lower region. In doing so, cooling air flow can be divided between the upper region and the lower region so that the lower region can be maintained at a lower temperature than the upper region.

Abstract: A method is provided for coating a hollow, internally cooled blade (1) of a gas turbine, in which method an outer coating (5) comprising an MCrAlY-based bonding layer (6) and a ceramic thermal barrier layer (9) of zirconium oxide is applied to the base material (3) of the blade (1) on the outer side of the blade (1) and an inner coating (4) comprising a Cr diffusion layer (7) is applied to the base material (3) of the blade (1) on the inner side of the blade (1). The MCrAlY-based bonding layer (6) is thereby applied to the finished blade (1). At the same time, along with the inner coating (4), the Cr diffusion layer (7) is also applied to the MCrAlY-based bonding layer (6) of the outer coating (5) by chemical vapour deposition. Subsequently, an Al diffusion layer (8) and an outer brittle Al build-up layer are applied by chemical vapour deposition to the bonding layer (6) coated with the Cr diffusion layer (7).

Abstract: A blade tip of a turbine rotor blade for a gas turbine engine, the turbine rotor blade including an airfoil and a root portion for mounting the airfoil along a radial axis to a rotor disk inboard of a turbine shroud, a pressure sidewall and a suction sidewall that join together at a leading edge and a trailing edge, the pressure sidewall and suction sidewall extending from the root portion to the blade tip, and a squealer tip cavity formed at the blade tip, the blade tip comprising: a trailing edge trench originating at the squealer tip cavity, wherein the trailing edge trench generally extends toward the trailing edge of the blade tip.

Abstract: A heat transfer system is provided for a turbine engine of the type including an annular casing with an array of generally radially-extending strut members disposed therein. The heat transfer system includes at least one primary heat pipe disposed at least partially inside a selected one of the strut members; at least one secondary heat pipe disposed outside the fan casing and thermally coupled to the at least one primary heat pipe and to a heat source. Heat from the heat source can be transferred through the secondary heat pipe to the primary heat pipe and to the selected strut member.

Abstract: A heater assembly for deicing and/or anti-icing a component includes a metallic heating element adjacent to a densely woven fabric layer impregnated with a resin that is capable of withstanding temperatures of up to 550° F. (288° C.).

Abstract: An apparatus for ice removal from a surface has an electrically resistive layer on the surface. An actuation device is provided for mechanically disturbing the surface, as for example deflecting, deforming, or vibrating the surface. When ice has accumulated, an interface layer of ice is melted by heating the electrically resistive layer with an electric current, and an electric current is applied to the actuation device to disturb the surface and release the ice. Alternative embodiments having various forms of actuation device are disclosed. An icemaker using the ice removal apparatus for ice release is described.

Abstract: A cover member defining a rotor chamber adjacent to a rotor wheel that supports a rotating blade in a turbine includes a first aperture for introducing a cooling gas stream into the rotor chamber, and a second aperture positioned in a radially outward portion of the cover member for allowing a portion of the cooling gas stream to exit the rotor chamber. The portion of the cooling gas stream exiting the rotor chamber carries dirt particles to purge the rotor chamber.

Abstract: A turbine blade for use in high temperature applications includes an as-cast airfoil portion and an as-cast outer tip shroud portion and the outer tip shroud portion has at least one as-cast internal cooling circuit for cooling the outer tip shroud, and the at least one as-cast internal cooling circuit having a plurality of exits for discharging cooling air over exterior surfaces of the shroud. A process for forming a turbine blade comprises the steps of forming an as-cast turbine blade having an airfoil portion and a tip shroud and the forming step comprising forming at least one as-cast cooling circuit within the tip shroud.

Abstract: There is described a vane wheel of a turbine comprising at least one vane, the footing thereof being held on a wheel disk. At least one cooling channel is arranged between the wheel disk and the vane footing. The vane wheel having a plurality of turbulators is embodied on at least one of the walls of the cooling channel, the turbulators being configured in such a way that the turbulence and thus the heat transfer of a cooling fluid flowing through the cooling channel are increased.

Abstract: A gas turbine rotor blade and rotor disk arrangement is provided. The rotor blade and the rotor disk includes a coating in order to seal the gap which normally exists between the bottom part of the blade root and the bottom part of the disk slot. The coating may be applied to the blade root or to the disk itself on the bottom part of the disk slot. The coating is preferably an abradable nickel-alloy Bentonite coating. The coating may also be applied to only a portion of the areas being sealed, omitting a portion at the start of the slot or at the start of one of the lobes of the blade root, so that a lead-in is formed, whereby the blade can more easily be inserted into the slot.

Abstract: A cooling circuit serves to provide cooling air in a rotor having a compressor rotor and a turbine rotor. The cooling circuit includes a first cooling path through the compressor rotor in series with a second cooling path through the turbine rotor such that one amount of cooling air is used for cooling both the compressor rotor and the turbine rotor. The construction enables a reduction in cooling air requirements, thereby increasing power plant efficiency.

Abstract: An engine cooling system includes a combustion chamber assembly configured to generate detonation waves and a first vapor cooling assembly. The combustion chamber assembly defines a flowpath between an inner liner and an outer liner. The first vapor cooling assembly includes a vaporization section located adjacent to the flowpath and a condenser section spaced from the flowpath, and is configured to transport thermal energy from the vaporization section to the condenser section through cyclical evaporation and condensation of a working medium sealed within the first vapor cooling assembly.

Abstract: A cooling system and method is provided for a wind turbine. The cooling system is used to cool at least one component in the wind turbine, and includes a filter system having at least one primary filter and at least one secondary filter. A switch selects between the primary filter and the secondary filter. A cooling circuit includes at least one conduit and a coolant medium. The filter system can switch from the primary filter to the secondary filter.

Abstract: A wind turbine generator capable of enhancing starting performance of the wind turbine generator when the outside air temperature is low and capable of securing temperature adjusting ability of a generator set in a wide outside air temperature range is provided. A wind turbine generator includes a nacelle (3) for accommodating a generator set (11, 12, 14) that generates electricity by rotation of a rotor having a plurality of blades, a ventilation opening (51) provided in a wall surface on a wind upstream side when the nacelle is normally operated, and a heat exchanging unit (13) disposed in the ventilation opening (51) to exchange heat between an outside air and a lubricant of the generator set (11, 12) such as an lubricant oil or cooling water.

Abstract: A cooling system and method is provided for a wind turbine. The cooling system is used to cool at least one component in the wind turbine, and includes a filter system having at least one primary filter and at least one secondary filter. A switch selects between the primary filter and the secondary filter. A cooling circuit includes at least one conduit and a coolant medium. The filter system can switch from the primary filter to the secondary filter.

Abstract: A gas turbine engine includes turbine blades having film cooling holes at an outer face of an airfoil wherein the film cooling holes are designed to be better filled with air. In a disclosed embodiment, the film cooling holes include a meter section extending along a direction having a main component extending from a blade tip to a blade root. In addition, a diffused section communicates with the meter section at a face of the airfoil. The diffused section is spaced toward the blade tip from the meter section. In this manner, centrifugal force ensures the diffused section is also filled with air.

Abstract: A lubrication system for a wind turbine having a control system, the lubrication system including a sump configured to collect a lubrication fluid, and at least one heating unit configured to heat the lubrication fluid based on at least one forecasted condition.

Abstract: A gas turbine engine component having shaped cooling holes that further enhances the cooling of a desired region while reducing stress levels in and around the cooling holes is disclosed. The cooling holes are generally elliptically-shaped and diffuse from a cooling fluid supply side to a discharge side and are oriented on the turbine component to reduce stress concentrations while directing the cooling fluid to a desired surface or location. The elliptical cooling holes have openings in the surface that have high points that are concentric and planar.

Abstract: An airfoil assembly for a turbine engine, that may comprise: an airfoil that includes at least one panel opening formed therein; and an aperture panel that includes at least one cooling aperture; wherein the panel opening and the aperture panel are configured such that, upon installation of the aperture panel within the panel opening, a substantially smooth outer surface of the airfoil is formed and at least one of the cooling apertures of the aperture panel forms an exit passageway from at least one hollow internal cavity within the airfoil.

Abstract: A stator vane for use in a gas turbine engine, the vane having a plurality of near wall mini serpentine flow cooling modules arranged in an array on the airfoil walls. Each module includes a series of multiple pass serpentine flow cooling channels extending in the airfoil chordwise direction. Each module is connected by a cooling air feed hole to a cooling air supply cavity and a cooling air discharge hole connected to a cooling air discharge cavity, where both cavities are formed between the airfoil walls. Each series of multiple pass serpentine cooling channels is connected together by a spanwise channel. The spanwise channels can include a row of film cooling holes to discharge film cooling air onto the airfoil external surface. The discharge cavity can also include film cooling holes to discharge cooling air from the cavity onto the airfoil external surface.

Abstract: A blade attachment apparatus for a turbine is provided and includes a member, which rotates about an axis of the turbine, including fore and aft disk posts that retain corresponding fore and aft turbine blades, respectively, fore and aft of a turbine nozzle and a central portion proximate the turbine nozzle and an attachment retained between the fore and aft disk posts and the fore and aft turbine blades and coupled to the central portion, such that a portion of the attachment interrupts a purge flow through the turbine.

Abstract: The invention relates to a component having a substrate and a protective layer, which consists of an intermediate NiCoCrAlY layer zone on or near the substrate and an outer layer zone which is arranged on the intermediate NiCoCrAlY layer zone, which is characterized in that the intermediate NiCoCrAlY layer zone comprises of (in wt %): 11-13% Co, 20-23% Cr, 10.2-11.5% Al, 0.3-0.5 Y, 1.0-2.5% Re and Ni balance.

Abstract: A rotor of a thermal fluid flow machine, especially a gas turbine, is provided. The rotor includes a plurality of rotor components that are held together by a common tie-bolt that extends through the center of the rotor components. The tie-bolt is fixed in at least one of the rotor components using at least one star spring that surrounds the tie-bolt in a circumferential direction.

Abstract: Systems and methods of deicing aircraft propellers include a deicing heater timing cycle, i.e. heater on/heater off schedule that is calculated as a function of ambient temperature and the airspeed of the aircraft.

Abstract: A heat dissipation device includes a heat sink and a cooling fan. The heat sink includes a substrate and a plurality of fins. The cooling fan includes a fan housing defining an air inlet and an air outlet at two opposite sides thereof, and an impeller received in the fan housing and mounted to a supporting base provided at the air inlet. The air outlet is positioned adjacent to the heat sink. The impeller includes a hub, a plurality of main blades and auxiliary blades. The hub includes a top wall located at the air outlet and a sidewall extending from the top wall. The main blades extend radially and outwardly from the sidewall. The auxiliary blades located on the top wall extend from a center of the top wall towards an outer periphery of the top wall.

Abstract: A heating and cooling system for a wind turbine is provided and includes a gearbox, gearbox heat exchanger, generator, generator heat exchanger, and a cooling duct. The cooling duct is connected to the gearbox and generator heat exchangers, and is used to transport air across both heat exchangers to cool the gearbox and generator.

Abstract: A process for producing a component of a gas turbine having a substrate with a metallic layer is provided. The metallic layer is a MCrAlX layer which is treated at temperatures elevated above the operating temperature, by at least 50° C., so that the oxidation and corrosion behavior are improved. In particular a MCrAlX of the type, NiCoCrAlX is used.

Abstract: A turbine blade damper arrangement in which a damper is positioned against the undersides of the platforms of adjacent turbine blades. In operation, the damper is centrifugally urged into engagement with the blade platforms to provide damping of relative movement between the blades. The damper and platform surfaces that it engages are of part-cylindrical configuration in order to minimise gas leakage paths between the damper and blade platforms.

Abstract: An air flow heating system for wind machine is provided, which can be mounted to a conventional, propeller driven wind machine's mast or tower, to provide a stream of heated updrafting air, for use by the wind machine, to supplement the overall convective air current in the vicinity of the wind machine. The system includes a heater assembly with a plurality of heater arms, each heater arm having at least a single burner, and each burner provided with a combustible fuel.

Abstract: A wind turbine, in an exemplary embodiment, includes a hub and at least one blade coupled to the hub, at least one electric pitch motor operatively coupled to the at least one blade, and a pitch motor control system operatively coupled to the at least one pitch motor. The pitch motor control system includes a plurality of components that produce excess heat during operation.

Abstract: A turbine vane of a gas turbine, especially a gas turbine aircraft engine, is disclosed. The turbine vane includes a vane base body with an outer surface forming a suction side and a pressure side, the outer surface of the vane base body being at least partially coated with a thermal barrier coating. The thermal barrier coating extends continuously or uninterruptedly at least largely over the suction side and largely over the pressure side of the surface of the vane base body, with the layer thickness of the thermal barrier coating being variable or adjustable.

Abstract: A wind energy converter includes a wind turbine, a wind turbine foundation and a temperature control mechanism for controlling the temperature of one or more areas of the wind turbine. The temperature control mechanism including a mechanism for exchanging heat. The wind energy converter is characterized in that the mechanism for exchanging heat is positioned in the ground outside the foundation. Also contemplated is a method for controlling the temperature of one or more areas of a wind energy converter and use thereof.

Abstract: According to an embodiment of the invention, an article of manufacture for use in a gas turbine engine is disclosed. The article comprises a part having a surface covered with a ceramic thermal barrier coating. The thermal barrier coating has an outer surface covered with a sacrificial phosphate coating, wherein the sacrificial phosphate coating reacts with contaminant compositions to prevent contaminant infiltration into the thermal barrier coating.

Abstract: A motion generating system for propelling and/or lifting a craft exploiting explosive and implosive processes, whereby a propulsive or lifting force on the craft arises from two sources: generation of a stream of fluids which imparts thrust to the craft; and generation of zones of reduced pressure and density in front of and above the craft allowing the craft to be thrust forward and lifted by ambient pressure on the rear and underside of the craft. Furthermore, reduction of fluid density in front of the craft results in a reduction of frontal drag allowing the attainment of higher speeds. The motion generating system may have other applications; for example, the motion generating system may be used to propel a stream of liquid in the manner of a pump.

Abstract: A cooled airfoil includes a concave pressure wall extending radially from a base to a tip of the airfoil, a convex suction wall connected to the concave pressure wall at a leading edge and a trailing edge spaced axially from the leading edge, and cooling channels extending radially between the base and the tip of the airfoil between the concave pressure wall and the convex suction wall and configured to receive a cooling fluid supply through the base of the airfoil. The cooling channels include a leading edge channel, a trailing edge channel, a serpentine cooling circuit, and a dedicated up-pass channel. The serpentine cooling circuit includes a first up-pass channel forward of the trailing edge channel and configured to be in flow communication with a supply channel through the base of the airfoil, a down-pass channel forward of and in flow communication with the first up-pass channel, and a second up-pass channel forward of and in flow communication with the down-pass channel.

Abstract: A cooled airfoil includes a concave pressure wall extending radially from a base to a tip of the airfoil, a convex suction wall connected to the concave pressure wall at a leading edge and a trailing edge spaced axially from the leading edge, and a plurality of cooling channels formed between the concave pressure wall and the convex suction wall and configured to receive a cooling fluid supply from the base of the airfoil. The cooling channels include a leading edge channel extending radially from the base toward the tip, a trailing edge channel extending radially from the base toward the tip and in flow communication with a plurality of trailing edge apertures adapted to exhaust cooling fluid to the exterior of the airfoil, a serpentine cooling circuit including a plurality of channels, and a dedicated up-pass channel extending radially from the base toward the tip between the leading edge channel and the forward most channel of the plurality of channels in the serpentine cooling circuit.

Abstract: With regard to gas turbine engines it will be appreciated that blades are typically cooled in order to ensure that the materials from which the blades are formed remain within acceptable operational parameters. Coolant is judiciously used in order to maintain engine operational efficiency. Unfortunately with regard to rotor blades horseshoe vortices tend to increase heating towards a pressure side of a blade resulting in localised overheating. Such localised overheating may result in premature failure of the blade component. Traditionally coolant flows have been presented over a forward projection of a blade platform. In such circumstances coolant flow will not be used as efficiently as possible with regard to protecting a pressure side of a platform in a blade assembly and arrangement. By provision of a deflector element on the forward blade platform coolant flow can be proportioned either side of a leading edge of the blade.

Abstract: A shroud suitable for use in a gas turbine engine exhibits substantially uniform thermal growth.

Abstract: A turbine blade rotor assembly is disclosed for a gas turbine engine. The assembly includes a rotor having nickel alloy turbine blades secured thereto. Each of the blades includes a root and an airfoil. The roots are supported by the rotor. A ceramic matrix composite platform separate from the turbine blades is supported between each pair of the turbine blades adjacent to the airfoils. In another example, the airfoil includes a perimeter. A shroud having an aperture receives the airfoil with a single shroud substantially surrounding the airfoil at the perimeter. In one example, the turbine blade includes high and low pressure sides opposite one another that extend from a tip to a root. The airfoil is free from any protrusions extending from the high and low pressure sides on a portion of the blade axially outward from the root.

Abstract: A method and apparatus for a rotor assembly for gas turbine engine are provided. A first rotor blade including an airfoil, a platform, a shank, an internal cavity, and a dovetail is provided, wherein the airfoil extends radially outward from the platform, which includes a radially outer surface and a radially inner surface, the shank extends radially inward from the platform, and the dovetail extends from the shank, such that the internal cavity is defined by the airfoil, the platform, the shank, and the dovetail. The first rotor blade is coupled to a rotor shaft such that during engine operation, cooling air is channeled from the cavity through an impingement cooling circuit for impingement cooling the first rotor blade platform radially inner surface, and a second rotor blade is coupled to the rotor shaft such that a platform gap is defined between the first and second rotor blade platforms.

Abstract: An airfoil assembly includes an airfoil with at least one cavity that is in communication with a source of cooling air. A baffle is disposed within that cavity and includes a plurality of openings for directing cooling air against the hot wall. A plurality of dividers extends between the baffle walls and the internal cavity to direct cooling air toward one of a leading edge chamber and a trailing edge chamber.

Abstract: A CMC wall (22) with a front surface (21) heated (24) by a working fluid in a gas turbine. A back CMC surface (23) is coated with a layer (42) of a thermally conductive material to accelerate heat transfer in the plane of the CMC wall (22), reducing thermal gradients (32-40) on the back CMC surface (23) caused by cold spots (32) resulting from impingement cooling flows (26). The conductive material (42) may have a coefficient of thermal conductivity at least 10 times greater than that of the CMC material (22), to provide a minimal thickness conductive layer (42). This reduces thermal gradient stresses within the CMC material (22), and minimizes differential thermal expansion stresses between the CMC material (22) and the thin conductive layer (42).

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: Fan assemblies, mechanical draft systems and methods are provided. In this regard, a representative mechanical draft system for use with multiple appliances includes: a chimney operative to direct combustion products from multiple appliances; a chimney fan operative to draw combustion products from the chimney; and a controller operative to adjust an operating speed of the chimney fan such that, responsive to a change in pressure in the chimney, the controller adjusts the operating speed of the chimney fan to maintain a desired pressure in the chimney.

Abstract: A turbine stage includes a stator nozzle having a row of vanes mounted between inner and outer bands. The inner band terminates in an ovate ledge converging aft from the vanes with radially outer and inner convex surfaces joined at a convex apex. The ovate ledge reduces aerodynamic losses at the rotary seal with a row of following turbine rotor blades.

Abstract: A composite turbine blade and fan blade assembly for use in a turbo machine such as a gas turbine engine. The turbine blade assembly includes a plurality of metallic turbine blades secured to a rotor disk and forming an outer shroud assembly. A fiber reinforced composite fan blade assembly includes an annular ring with a plurality of fan blades extending there from. The annular ring has a conical shape to allow for easy installation over the outer shroud assembly, and is secured to the outer shroud assembly to force the fan blade assembly to rotate with the turbine blades, yet allow for a slight radial movement between the fan blades and the turbine blades. The turbine blades are formed of a high temperature super alloy to allow for use in a gas turbine engine, and the fan blade assembly is made from a fiber reinforced composite to allow for higher rotational speeds and smaller diameters for smaller gas turbine engines without exceeding allowable stress limits.

Abstract: An impeller (10) of a cooling fan includes a hub (11) and a plurality of blades (12) radially and outwardly extending from the hub. Each of the blades includes a windward surface (127) and a leeward surface (128). A groove (124) is defined in each of the blades for guiding airflow from the windward surface to the leeward surface.