620results about "Turbines" patented technology

Disclosed herein is an integrated in-line emergency lube pump assembly for a planetary gearbox, including a main lube pump, a coaxial lube pump coaxially aligned with and coupled to the planetary gearbox, and an oil tank having a main section and an auxiliary section. There is no physical boundary between the main section and the auxiliary section of the oil tank. When the detected pressure in the oil supply line is equal to or exceeds a predetermined pressure threshold, the main lube pump operates to causes the main pump to pump oil from the main section of the oil tank to the planetary gearbox. Alternatively, when the detected pressure in the oil supply line is less than the predetermined pressure threshold the main lube pump ceases operation to cause the coaxial lube pump to draw oil from the auxiliary section of the oil tank to the planetary gearbox.

The invention relates to the technical field of high-temperature alloy additive manufacturing, and discloses a method for preparing a low-crack-sensitivity nickel-based high-temperature alloy through selective laser melting and application of the low-crack-sensitivity nickel-based high-temperature alloy. And selective laser melting forming treatment is conducted on the core-shell structure powder, specifically, regional laser scanning, interlayer pulse pressurization and in-situ heat treatment are conducted, and the low-crack-sensitivity nickel-based high-temperature alloy is obtained. According to the invention, synchronous regulation and control of microcrack closing and nanophase precipitation are realized through design of core-shell structure powder, laser regional dynamic energy control scanning of hot lines and interlayer in-situ densification. By means of the method, the SLM forming crack density can be reduced to be smaller than or equal to 0.1 mm / mm, the anisotropy index is smaller than or equal to 1.05, the post-treatment hot isostatic pressing link is omitted, and the method can be applied to thin-wall high-cycle fatigue parts such as aero-engine flame tubes and aerospace turbine pump rotors.

The invention discloses a laser additive repairing method for 17-4PH steam turbine blades, and relates to a laser repairing method for steam turbine blades. The method aims at solving the technical problem that in existing laser additive 17-4PH turbine blade repairing, the difference between the thermal expansion coefficient of Co-based corrosion-resistant alloy and the thermal expansion coefficient of 17-4PH alloy is large, and consequently the interface performance is reduced after repairing is conducted. According to the laser additive repairing method for the 17-4PH steam turbine blade, due to the fact that 17-4PH is poor in water erosion resistance, Co-based alloy powder is used for preparing a water erosion resistance structure, on the basis, in consideration of the problem that the performance difference of the Co-based alloy powder and the 17-4PH material is large, a powder component gradient transition mode is adopted, and the water erosion resistance of the 17-4PH steam turbine blade is greatly improved. The problem of low interfacial strength during direct transition is solved, the binding capacity of the Co-based water erosion resistant layer and the 17-4PH matrix is improved, and finally the service time of the 17-4PH blade is prolonged.

The invention relates to the field of industrial manufacturing, and provides a defect correction method and system for aerospace engine turbine blade additive manufacturing. The method includes; a quantum dot marking agent is adsorbed to the surface of metal powder, signal data in the additive manufacturing process are collected, and a multi-dimensional monitoring characteristic parameter set is obtained; constructing a graph neural network model, taking the multi-dimensional monitoring characteristic parameter set as node characteristic input, and performing spatio-temporal evolution prediction on a molten pool area to obtain a defect probability distribution field; performing multi-target optimization on the process parameter vector through a Bayesian optimization algorithm to obtain a self-adaptive regulation and control parameter combination, and performing partition regulation and control on the manufacturing process; and for the residual defects of the partition regulation and control result, a reinforcement learning strategy network is adopted to select a repair action sequence, the repair action sequence is input into an execution unit to perform cross-level cooperative processing on the residual defects, and a repaired blade structure is obtained. According to the method, intelligent prediction and active correction of the turbine blade additive manufacturing defects are achieved.

The invention belongs to the technical field of aero-engine materials, and particularly relates to a high-temperature-resistant aero-engine blade and a preparation method thereof.According to the blade, titanium-aluminum alloy is adopted as a base body, and two novel materials including a core-shell structure self-healing reinforcement phase and a rare earth niobate ferroelastic protection compound are introduced, so that the high-temperature-resistant aero-engine blade is obtained; and the comprehensive performance of the blade in a high-temperature environment is remarkably improved. The core-shell reinforcement phase takes titanium diboride as a core and boron nitride as a shell, and can realize a self-healing function at a high temperature; the rare earth niobate compound has ferroelastic behavior and ultralow thermal conductivity through doping modification of tantalum and rhenium. The preparation process comprises the following steps: mixing the titanium-aluminum alloy powder, the reinforcing phase, the silicon carbide short fibers and the binder, carrying out spark plasma sintering molding, carrying out machining, applying the protective coating through plasma spraying, and finally carrying out heat treatment. According to the method, the process is simplified, and the prepared blade has excellent high-temperature resistance, oxidation resistance and mechanical property and is suitable for aero-engine hot end components.

A vacuum induction melting (VIM) furnace and method of using. A VIM is configured for use with metal-containing carbon powders. In operation, the metal-containing carbon powders are formed into a pellet so as to minimize or eliminate ejection of material during introduction of the pellet into the VIM processor. The VIM processor may be substituted or used in combination with a vacuum arc melt processing apparatus, an electron beam melt furnace, an ion plating furnace, a plasma flame source, a smelter, a traditional metal-metal melt furnace, or any equivalent. Pelletizing can be accomplished through use of a press or through application of any pelletizing technique and / or use of any apparatus that is able to generate pellets that have sufficient mass to avoid ejection from a VIM processor.

An oil supply system, has: an oil tank; a closed-loop oil circuit including a component, a pump having a pump inlet and a pump outlet, and a de-aerator; a make-up flow conduit connecting the oil tank to the closed-loop oil circuit; a make-up pump connected to the closed-loop oil circuit; a recirculation conduit in parallel to the make-up pump and bypassing the make-up pump; a fixed orifice fluidly connected to the make-up flow conduit upstream of the closed-loop oil circuit and downstream of the recirculation conduit, the fixed orifice having an area selected to increase a pressure at an outlet of the make-up pump above a pressure at the de-aerator inlet; and a metering orifice connected on the recirculation conduit and having a flow circulating area selected to reduce a flow rate inputted into the closed-loop oil circuit below a flow rate generated by the make-up pump.

Inventive techniques for forming unique compositions of matter are disclosed, as well as associated physical characteristics and properties of the materials. In particular, particles comprising a metal lattice are characterized by having carbon (preferably graphene) disposed within the crystalline lattice structure thereof. The carbon is at least partially disposed in interstitial sites of the metal lattice, and may be present in amounts ranging from about 15 wt % to about 90 wt % of the total particle mass, with about 15 wt % to about 60 wt % being disposed in the interstitial sites, e.g., between basal planes, of the metal lattice. The carbon, moreover, is substantially homogeneously dispersed throughout the resultant material, conveying unique and advantageous properties such as strength-to-weight ratio, density, mechanical toughness, sheer strength, flex strength, hardness, anti-corrosiveness, electrical and / or thermal conductivity, etc. as described herein. In some approaches, the graphene is pristine, and has corresponding physical characteristics as described herein.

An oil distributor is provided for a bearing chamber. The oil distributor includes an outer annular manifold, which is integrally formed with an interior surface of the bearing chamber and which defines a first interior cavity receptive of pressurized oil which defines a first interior cavity receptive of pressurized oil, an inner annular manifold, which defines a second interior cavity and comprises an interior facing surface defining holes respectively communicative with the second interior cavity and one or more support struts by which the inner annular manifold is supported within the outer annular manifold. Each of the one or more support struts defines interior lines by which the pressurized oil flows from the first interior cavity to the second interior cavity and to the holes.

A braze tape includes a first portion, a second portion, and a central portion located between and adjacent to the first portion and the second portion. The first portion includes a low melting temperature alloy material, the second portion includes a high melting temperature alloy material, and the central portion includes a mixture of the low melting temperature alloy material and the high melting temperature alloy material. The central portion may gradually change between the low melting temperature alloy material and the high melting temperature alloy material. A related method is also disclosed.

The invention provides a nickel-based superalloy with high cobalt and gamma'content and a preparation method and application thereof, and belongs to the technical field of superalloy additive manufacturing. The invention provides spherical nickel-based high-temperature alloy powder which comprises the following components in percentage by mass: 20 to 30 percent of Co, 8 to 10 percent of Cr, 4 to 7 percent of Al + Ti, 14 to 16 percent of W + Mo + Ta, 0 to 3.5 percent of Nb, 0.001 to 0.015 percent of B, 0.05 to 0.1 percent of C and the balance of Ni. And the spherical nickel-based superalloy powder serves as a printing raw material, laser additive manufacturing is conducted, then heat treatment is conducted, and the nickel-based superalloy with the high cobalt content and the high gamma'content is obtained. The nickel-based superalloy with high cobalt and gamma'content prepared by the method has excellent high-temperature mechanical properties and excellent printability, and is suitable for direct forming of turbine blades or additive manufacturing repair of damaged turbine blades.

The invention relates to a method for the additive manufacturing of a combustion chamber (5). A cross-sectional shape (2) is defined as a closed, parameterized curve, and an extrusion curve (3) is defined as an open, three-dimensional curve. The material is applied along the extrusion curve (3) to form a strand (1), the cross-section of which is defined by the cross-sectional shape. A first end of the strand (1) forms one of the inlet openings (1a), and a second end of the strand (1) forms one of the outlet openings (1b). The application step is repeated at least once to produce further strands (1). The strands (1) are arranged such that they form a wall of the combustion chamber (5), with the outlet opening (1b) of each strand facing an interior region of the combustion chamber (5).

A fan exit guide vane with acoustic treatment includes a leading edge and a trailing edge opposite chordwise from the leading edge; a radially inner attachment region opposite spanwise from a radially outer attachment region; a span dimension extending between the radially inner attachment region and the radially outer attachment region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the fan exit guide vane; and an acoustic zone integrally formed within the fan exit guide vane. A gas turbine engine includes the fan exit guide vane with acoustic treatment. A process forms the fan exit guide vane with acoustic treatment.

Variable pitch fan blade (8a) for an unfaired turbomachine fan, the blade (8a) comprising an aerodynamic blade (9) and a base (10), the blade (9) comprising a recess opening simultaneously onto the upper and lower surfaces (14) and a leading edge (15) of the blade (9), the blade (8a) further comprising a measuring device (26) housed in the recess and clipped onto the blade (9), the measuring device (26) comprising a streamlined body (28) whose outer surface (29) is complementary to the outer surface (30) of the blade (9) to ensure aerodynamic continuity, the body (28) carrying at least one measuring instrument (31) configured to measure at least one parameter of an airflow upstream of the blade (9). Figure for the abstract: 2

The present invention relates to a valve arrangement. The present invention further relates to a fluid flow control element. Furthermore, the present invention relates to a turbomachine system comprising such a valve arrangement and / or fluid flow control element. Moreover, the present invention relates to a method and a use relating to the valve arrangement and / or the fluid flow control element.

There is provided a fluid flow machine (10) comprising: a casing structure (24), a turbomachine blade (311) disposed within the casing structure (24), and a sealing arrangement (330) coupled to the casing structure (24). The sealing arrangement comprises (330) an abrasion portion (334) including a lattice structure (400). The abrasion portion (334) is configured to provide a seal with a tip (321) of the turbomachine blade (311). There is also provided a method (600) of manufacturing such a fluid flow machine (10).

A method of making a fluid injector for a gas turbine engine includes depositing material (104) onto a piece of tube stock (102). The method includes machining the deposited material into a fluid injector component (106). Depositing can include laser cladding the material onto the piece of tube stock. The method can include placing or flowing braze into a braze joint location (116) between the deposited material and another fluid injector component and forming the braze into a braze joint in the braze joint location.

The invention relates to a component (50) obtained by injection molding, in particular from a mixture comprising an adhesive and a powder, the component having a plurality of protrusions (40) protruding from a main surface (241) of the component (50), the plurality of protrusions (40) comprising at least one first rib (41) and at least one second rib (42) intersecting at an intersection point (400) comprising a midpoint (44) of the component (50).

A method (100) for manufacturing turbomachine blades, the manufacturing method (100) comprising the steps of: providing a tubular part (20) extending along a first axis (A1) between a first (22) end and a second end (24); machining the tubular part (20) to produce a plurality of blade blanks (10), each blade blank (10) comprising a core (16) extending along a longitudinal axis (X) between a first end portion (12; 14) and a second end portion (12; 14); machining the first end portion (12; 14) and / or the second end portion (12; 14) of each blade blank (10) to produce a reference element; separating the blade blanks (10) from each other. Abstract figure: Figure 4

The invention relates to a metal powder for an additive manufacturing method, the metal powder comprising an alloy comprising, by weight, between 23% and 24.5% of chromium, between 9% and 11% of nickel, between 6.5% and 7.5% of tungsten, between 3% and 4% of tantalum, between 0.55% and 0.65% of carbon, between 0.3% and 0.5% of zirconium, between 0.15% and 0.25% of titanium, at most 2% of iron, at most 0.3% of silicon, at most 0.1% of manganese, at most 0.1% of copper, at most 0.015% of sulfur, at most 0.015% of phosphorus, at most 0.01% of boron, at most 0.025% of oxygen, at most 0.020% of nitrogen and at most 0.010% of hydrogen and less than 0.050% of other elements in total, the remainder being cobalt.

A damper element and a vibration dampening system (120, 320) including a plurality of the stacked damper elements (170, 370), are provided. The damper element(s) (170, 370) may be used in a body opening (160) in a rotating blade (114) in a turbine (108). The damper elements (170, 370) include a head member (172, 202, 224, 342) having an at least partially ramped surface (174, 204, 226, 344), and a plurality of flexible legs (180) extending from the head member. Each flexible leg (180) includes a radially extending body section (182) having an end section (184), and the end section has an outer end surface (186) and an inner end surface (188). Centrifugal force caused by rotation of the rotating blade (114) cause the head member (172, 202, 224, 342) of an adjacent damper element (170, 370) to force the outer end surfaces (186) of the plurality of flexible legs (180) into frictional engagement with an inner surface (162) of the body opening (160) to dampen vibration. In certain embodiments, an elongated body (392) may extend through an opening (390) in the head member of the damper element(s).

A niobium silicide-based alloy and a turbine having a turbine component formed from at least the niobium silicide-based alloy are provided, the niobium silicide-based alloy comprising about 14 atomic percent to about 24 atomic percent titanium (Ti), about 11 atomic percent to about 19 atomic percent silicon (Si), about 4 atomic percent to about 8 atomic percent chromium (Cr), about 2 atomic percent to about 6 atomic percent hafnium (Hf), up to about 4 atomic percent aluminum (Al), about 0.5 atomic percent to about 1 atomic percent tin (Sn), about 5 atomic percent to about 15 atomic percent tantalum (Ta), about 1 atomic percent to about 5 atomic percent tungsten (W), up to about 5 atomic percent rhenium (Re), up to about 5 atomic percent zirconium (Zr), up to about 6 atomic percent yttrium (Y), and the balance niobium (Nb).

The invention relates to a method for securing a spring washer (05) to a combustion chamber housing (02). The combustion chamber housing (02) has a cylindrical receiving portion (03) at least at the downstream end. The spring washer (05) comprises a cylindrical fastening portion (06) and a plurality of spring tabs (07) distributed over the circumference. First, a solder material is applied to the inner face of the fastening portion (06). Then, the spring washer (05) is joined onto the combustion chamber housing (02) such that the fastening portion (06) is located on the receiving portion (03). A connection of the fastening portion (06) to the receiving portion (03) is achieved by heating the fastening portion (06) and / or the receiving portion (03). According to the invention, the spring tabs (07) are cooled at the same time.

The present disclosure may be embodied as a blade for a wind turbine. The blade includes a spar and a blade body arranged around the spar. The blade may include a root, a tip, and one or more body sections, each body section having a length, a stiffness ratio. The blade may further include two or more boundary actuators, each boundary actuator positioned at a boundary end of a body section, wherein each boundary actuator is configured to engage the corresponding boundary end to twist the body section. The length and stiffness ratio of each section may be optimized for maximum efficiency during Region 2 operation.

An engine oil system includes an oil tank and an access port cover. The oil tank includes a wall forming an internal oil cavity. The wall forms an access port extending along an access port axis through the wall. The access port cover is mounted on the wall and seals the access port. The access port cover includes a cover body and a packing ring. The cover body extends between and to an inner axial end and an outer axial end. The cover body forms a cover lip at the outer axial end. The cover lip extends circumferentially about the access port axis. The cover lip extends radially outward and axially outward, relative to the access port axis, along a lip axis to the outer axial end. The packing ring forms a fluid seal between the cover body and the wall within the access port.

The invention relates to an additive manufacturing method for a wall (10) of a turbine engine, the wall (10) comprising a first cooling hole (40). The manufacturing method comprises additively manufacturing the wall (10) by selective melting or selective sintering on a powder bed. The wall (10) is manufactured with at least one lower region (19a) at least partially surrounding the first cooling hole (40) having a minimum length (La) between 0.01 mm and 0.4 mm and / or a lower region (19a) having a total thickness (ea) between 0.06 mm and 0.22 mm.

Disclosed is a counter-form (1) for producing metal aeronautical parts, in particular a turbine part, by solid-phase densification, comprising a composite material containing, on the one hand, a first phase having formula Mn+1AlCn, where n = 1 to 3 and M is a transition metal selected from the group consisting of titanium and / or molybdenum and / or niobium and / or chromium, and, on the other hand, a second phase having formula Al4C3.

An engine oil system for an aircraft engine includes an oil tank and an oil tank return. The oil tank forms an internal oil cavity. The oil tank return forms a portion of an oil flow path through the engine oil system. The oil tank return includes a return body forming a return passage extending between and to a passage inlet and a passage outlet. The return body includes an inner diameter wall and an outer diameter wall. The inner diameter wall and the outer diameter wall form a turn of the return passage between the passage inlet and the passage outlet. The return body forms a top side, a corner, and an oblique side at the passage outlet. The corner and the oblique side are disposed at the outer diameter wall. The corner extends between the top side and the oblique side.

A manufacturing method of a composite part, which utilizes a hot isostatic pressing process to prepare the part, the composite part comprising a plurality of sub-parts, an independent jacket is provided for each sub-part, each sub-part is filled with raw material powder and subjected to warm isostatic pressing preforming, then the jacket connection part is cut open and welded together, and the whole jacket is subjected to hot isostatic pressing processing to obtain a composite part blank. The method can realize full and effective filling of raw material powder in a complex structure, and improve the production efficiency and product quality of the composite part. The application also provides a manufacturing method of a double-alloy integral blisk, a guide vane and a blade ring.

A method of manufacturing a component (100) for a turbo machine. The component (100) comprises a main body (104) having a fluid inlet (103) and fluid outlet (200). A cooling passage (204) extends between a fluid inlet (103) and a fluid outlet (200). The cooling passage (204) is divided into a first (204A) and a second section (204B) which extend between the fluid inlet (103) and fluid outlet (200). The first section (204A) has a first predetermined roughness; and the second section (204B) has a second predetermined surface roughness; the predetermined surface roughness in at least one of the first section (204A) and second section (204B) is defined by a plurality of spaced apart micro ribs (212) which extend at least part of the way across the cooling passage (204), at least one of the cooling passage sections is formed to further comprise macro ribs (206) which extend across the cooling passage (204), at least one micro rib is being formed between adjacent macro ribs (206).