586results about "Blade accessories" patented technology

A wind energy plant with a nacelle, a rotor, which features at least one rotor blade adjustable in its blade pitch angle, a central control device for controlling the wind energy plant and a control unit disposed in the rotor for controlling the blade pitch angle of the at least one rotor blade, wherein the central control device and the control unit in the rotor can exchange data with each other via a data link, which comprises at least one first sending and receiving device at the nacelle side and at least one second sending and receiving device at the rotor side, wherein a wireless network connection with a safety-oriented communication protocol is provided between the at least one first sending and receiving device and the at least one second sending and receiving device, and that one monitoring- and communication device at a time is associated to the first and / or the second sending and receiving device, which can monitor the function of the sending and receiving device and initiate a predetermined action in the case of an error.

The application belongs to the field of turbine blade thermal barrier coating of an aero-engine. Specifically, it is a structure design, a preparation method and a prepared coating of a double-layer radiation-resistant thermal barrier coating with a high-reflection layer and a high-absorption layer. The coating comprises a high-absorption layer and a high-reflection layer. The high-absorption layer can absorb part of the energy of photon radiation and re-emit it into the environment, while relying on a cooling gas film to quickly remove heat. Part of the energy of the photon radiation is reflected by the high-reflection layer again to the high-absorption layer after passing through the high-absorption layer, and the process of absorption-emission-cooling gas film heat dissipation is repeated. Within a set temperature range, the absorption rate of the high-absorption layer is determined based on the temperature difference requirement between the inner and outer surfaces of the thermal barrier coating, and the high-absorption layer material is determined based on the absorption rate. The application greatly reduces the radiation heat flow penetrating the coating, avoids the direct heating of the high-temperature alloy by infrared radiation, and significantly improves the heat insulation capacity of the thermal barrier coating and the service life of the turbine blade.

A method for manufacturing a tubular fiber preform for a composite material part. A method for manufacturing a tubular preform comprises: - the creation by three-dimensional weaving of two fiber skins, - the shaping of the first fiber skin so as to form a first part (310) of a tubular fiber preform comprising a first portion (314) of the outer shell of a cylinder (C) of predetermined diameter (D), - the shaping of the second fiber skin so as to form a second part (320) of a tubular fiber preform comprising a second portion (324) of the outer shell of the cylinder of predetermined diameter, - the fitting of the second part (320) into the first part (310) of the tubular fiber preform so as to form a tubular fiber preform (300) having an external surface corresponding to the joining of the first and second portions (314,324) of the outer shell of the cylinder (C) of predetermined diameter (D), the remaining portions (3211, 3212) of the second skin (200) being folded inside the tubular fibrous preform. Figure for the abstract: Fig. 9.

Disclosed is a turbine airfoil including an airfoil body including an outer wall configured to define an internal space, in which a cooling hole is formed in the outer wall and allows the internal space S and an external space of the airfoil body to communicate with each other. The inner surface of the outer wall, which defines the cooling hole, includes a first inner surface, and a second inner surface configured to define an outer hole region of the cooling hole H that communicates with the external space, in which the outer hole region is provided as a plurality of outer hole regions spaced apart from one another, and in which the outer wall includes a partition section provided between the two adjacent outer hole regions.

The present disclosure provides a kind of blade for inhibiting the blade of leakage vortex break, blade assembly, impeller machinery and blade shaping method.The blade profile is defined according to the cross section profile of the height direction accumulation, the cross section profile has the ratio of the height of the cross section profile and the total height of the blade, the dimensionless blade height H=1 is the blade tip, the dimensionless blade height H=0 is the blade root;The cross section profile also has the back bending angle δ;At the dimensionless blade height H=1, the back bending angle δ=13 °, and has the tolerance range of the numerical ratio ±20%.

The invention relates to a stator assembly (10) for a turbine engine compressor, comprising two successive rows of upstream stator vanes (121) and downstream stator vanes (124) defining a tandem configuration, the vanes (121, 124) extending radially between inner and outer support members, wherein, within the tandem configuration, the chord (26) of the upstream vane (121), extending between the leading edge and the trailing edge of the vane, increases with the height of the upstream vane: at 0% of the height of the upstream vane (121), the chord (26) of the upstream vane is at least 1.9 times smaller than the chord of the upstream vane at 100% of the height of the upstream vane and, at 70% of the height of the upstream vane (121), the chord (26) of the upstream vane is at least 1.2 times smaller than the chord of the upstream vane at 100% of the height of the upstream vane.

The present invention relates to a tool (4) for fixing two hollow shafts (10, 20) together, the tool being configured to be engaged in the two shafts which respectively comprise first and second annular flanges (100, 200) adapted to be fixed together by screws (70), the tool comprising: - a guide (6) adapted to be engaged in the two shafts (10, 20), - a sleeve (5) mounted on the guide and comprising a body (50) having an end (51) adapted to be applied axially against one of the shafts, - at least one arm (53) which comprises a series of holes (532), and - several rods (7) each comprising a portion (72) which can include one of the screws, the rods being adapted to slide axially in the holes (532) to a position in which the portions (72) are at an axial distance from the end (51). Figure for the summary: Figure 10

This invention belongs to the field of laser surface engineering and metal matrix composite technology, and provides a Ti / Nb microalloyed Inconel 625 toughening coating, its composite powder, and preparation method. The composite powder is composed of gas-atomized Inconel 625 alloy powder, titanium powder, and niobium powder, with a Ti to Nb mass ratio of (0.8–1.2):1 and a total addition amount of 0.8–1.5 wt%. A dual-laser cladding process is employed. An initial coating is first formed by laser cladding in an inert atmosphere, followed by in-situ pulsed laser shock blasting when cooled to 900–1000℃. After slow cooling, a toughened coating is obtained. The resulting coating contains 10–20 nm dispersed Ni3(Nb,Ti) composite γ-rays. , The coating contains no Ni2(Nb,Si) Laves phase. It exhibits both high strength and high toughness, along with excellent high-temperature stability.

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An airfoil (123) of an engine nozzle segment (140, 800) includes: a convex panel (920, 1120); a concave panel (1020); an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140). The convex panel (920, 1120) includes a plurality of metering holes (910, 1010, 1110), the airfoil (123) characterized by the following equation, in which a convex panel (920, 1120) metering hole area (1230) (CVXA (860, 980)), inner band (940, 1040, 1140) radius (IBR), and loading area (840, 1090) (LA) are related as follows: EQ1=CVXAm20.000005m2+0.000005IBRmπmLAm20.00001m2, wherein 0.0000001 meters squared (m2) ≤ CVXA ≤ 0.000004 m2, 0.213 m ≤ IBR ≤ 0.221 m, 0.0000147 m2 ≤ LA ≤ 0.0000153 m2, and 4.835 ≤ Equation 1 ≤ 281.150.

An inlet cone (20) for an aircraft turbine engine (1) is configured to be rotated about an axis (X). The inlet cone comprises a frustoconical body (22) and a tip (21) made of elastically deformable material fixed to a smaller-diameter end (221) of said body. The body comprises through-housings (222) distributed around the axis, and parts (23) made of elastically deformable material located in the through-housings. Each of the parts (23) made of elastically deformable material comprises at least one peripheral flange (230) for retaining the part in the corresponding through-housing (222). The peripheral flange bears against an internal surface (22a) of the body.

The present invention discloses a system and method for performing congestion control at an apparatus. The apparatus may establish at least one first connection with a first node and establish a plurality of second connections with a second node. After the at least one first connection and the plurality of second connections are established, the apparatus may transmit data packets to the second node through the plurality of second connections according to an outbound traffic policy when receiving the data packets from the first node through the at least one first connection. If a first criterion is satisfied, the apparatus may establish a plurality of third connections with the second node and transmit data packets to the second node through the plurality of third connections when receiving data packets from the first node through the at least one first connection.

A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The fan comprises a plurality of fan blades that define a fan diameter and a BEAL. The fan case comprises an inlet and an inlet length between the inlet and the fan. The acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the BEAL determines an acoustic spacing ratio of the gas turbine engine.

Provided herein are compositions and methods for removing a coating, such as an environmental barrier coating (EBC), from a substrate. By some approaches, a method for removing a coating from a substrate comprises contacting the coating with a salt, the salt being a solid. The method also involves heating the coating and the salt to a temperature at or above the melting point of the salt to convert the coating to a dissolvable reaction product by fusion of the coating with the salt. After forming the dissolvable reaction product, the product is then dissolved using an acid to remove at least a portion of the coating from the substrate.

TITLE: Rough of a part produced by additive manufacturing and method of obtaining it The invention relates to a rough (10) of a part (1) comprising a first lateral wall (11) from which extends a supported portion (13), the rough (10) being produced by additive manufacturing comprising the melting and solidification of successive layers of powder, the rough (10) comprising a support (3) intended to support the supported portion (13) during its manufacture, the support (3) extending under the supported portion (13) and comprising a plurality of lamellae (31) extending side by side, each lamella (31) extending along the manufacturing direction V from an internal end (311) connected to the supported portion (13) and forming a breakable zone, to an external end (312), the support (3) being configured to be removed from the part (1),Each of the lamellae (31) is configured to be broken at least at the level of said breakable zone by application of a force at its outer end (312) and in a direction transverse to a direction of elongation of the lamella (31). Figure 2,

The utility model relates to turbine manufacturing technical field especially a kind of high-strength light-weight turbine blade, including the blade root, blade body and blade cap connected in turn, the sequential quantity ratio of three is:1:2:1, arc line is provided with straining connecting bridge between every two adjacent blade body;The inside of blade root and blade body is provided with inner weight-reducing cavity, the outside of blade root is provided with outer weight-reducing cavity;Honeycomb filling frame is provided in the inner weight-reducing cavity, the outer weight-reducing cavity is connected with "Z" continuous support frame;The outside of one end of blade root and blade body is provided with arc-shaped clamping slot, the arc-shaped curve of arc-shaped clamping slot and the arc-shaped curve of straining connecting bridge are both common center;The utility model can effectively enhance the overall strength and rigidity of blade, improve the stability of blade under complex stress condition, reduce the risk of blade deformation and fracture;Realize to turbine blade two blade body is a group of synchronous installation, improve installation efficiency.

A coating system includes a bond layer including CoNiCrAlY; a thermal barrier layer at least partially disposed on the bond layer and including yttria stabilised zirconia; a mesh layer connecting the thermal barrier layer to the bond layer and including CoNiCrAlY; and an abradable layer disposed on the thermal barrier layer and including magnesium spinel. The mesh layer includes a plurality of cells connected to each other and is at least partially embedded in the thermal barrier layer. Each cell defines a cell opening therethrough and the cell opening of each of the plurality of cells at least partially receives the thermal barrier layer therein. The abradable layer and the thermal barrier layer are at least partially removable from the mesh layer and the bond layer without damaging them.

TITLE: Ventilated Turbine Rotor The invention relates to a turbine rotor centered on an axis X and comprising: a disc (20) having a web (21), a shell (60) having an annular radial mounting flange (61), the radial flange (61) of the shell (60) being fixed to the web (21) by a fastening member (70), a flange (80) comprising a first wall (82) extending axially and a second wall (81) extending radially outwards from the first wall (82), a ventilation circuit (90) configured to connect a radially internal cavity (15) formed by a radially internal side of the shell (60), with a radially external cavity (16) delimited by the flange (80) and by the disc (20), the ventilation circuit (90) comprising a first opening (91) in the shell (60) in communication Fluidic with a second opening (92) in the flange. Figure 3

The invention relates to a process for manufacturing a turbomachine blade, wherein: —a part (4) is manufactured comprising a foot (4), a heel (6) and an air stream zone (10) extending between the foot and the heel, the air stream zone comprising at least one protuberance (20, 24, 26) projecting from a main face (12) of the zone, the manufacturing being performed by injecting a mixture comprising a binder and a powder, the powder comprising at least a metal or a ceramic: —debinding is performed on the part so as to eliminate a greater quantity of the binder from the part: —heat treatment is performed on the part; and—the or each protuberance is eliminated from the air stream zone.

Methods, apparatus, systems, and articles for providing damping of airfoils are disclosed. An example airfoil is disposed in a flow path and includes a shell defining an outer surface of the airfoil and forming a cavity in an inner surface of the airfoil, and a grid damper disposed in the cavity for reducing vibrational loads applied to the airfoil.

A turbine blade for use in a turbine engine or a gas turbine comprises a locking portion and an airfoil with a uniform structure, in which there is at least one pocket filled with powder. In the at least one pocket there is at least one pin connected to the airfoil and at least one bar connected to the airfoil and / or at least one bar connected to the pin.

A process for manufacturing a part in SiC / SiC composite material. A process for manufacturing a part in SiC / SiC composite material comprising: - the production of a fibrous preform (20) from refractory material fibers; - a first densification step by chemical infiltration in the gas phase; - a second densification step by chemical deposition in the gas phase so as to obtain an intermediate part comprising the fibrous preform (20) densified throughout by the first matrix phase with a surface coating constituted by the second matrix phase; - the formation in the coating formed by chemical vapor deposition of one or more access openings to the porosity of the fibrous preform; - a third densification step comprising the impregnation of the fibrous preform with an impregnation composition containing at least silicon so as to obtain a part in ceramic matrix composite material.Figure for the summary: Fig. 2.