73results about "Turbines" patented technology

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.

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).

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.

A pump assembly (10) of an aircraft engine includes a pump housing (36), a first pump stage (12) positioned in the housing (36) and including two first stage rotating elements (20, 22) meshed at a first mesh location (24), and a second pump stage (14) positioned in the housing (36) and including two second stage rotating elements (26, 28) meshed at a second mesh location (30). The first stage rotating elements (20, 22) are retained at respective shafts (16, 18) by one or more woodruff keys (44), and at least one second stage rotating element (26, 28) of the two second stage rotating elements (26, 28) is secured at its respective shaft (16, 18) via a spline connection (70) of the at least one second stage rotating element (26, 28) to the respective shaft (16, 18).

A method of manufacturing a turbine vane (116) within an engine case includes additively manufacturing a combustor liner (118) within an engine case, additively manufacturing a support structure (130) attached to the combustor liner (118) at a radially distal position, and additively manufacturing the turbine vane (116) attached to the support structure (130) at an inwardly adjacent position to the radially distal position.

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,

A de-aerator is provided that includes a body, a cover panel, a fluid inlet, a helical fluid passage, at least one partition, and at least one fluid outlet. The body extends between opposing first and second axial ends. The body has a sidewall, a base panel, and an internal cavity. The fluid inlet is in communication with the body and is configured to direct fluid tangentially into the internal cavity. The helical fluid passage is disposed within the internal cavity and has entry and exit ends. The helical fluid passage has circumferential turns that each include an air passage. The at least one partition is disposed within the internal cavity at or below the passage exit end, spaced above the base panel. The at least one fluid outlet is configured to permit liquid passage from the internal cavity of the body to outside the body.

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.

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.

This invention relates to the field of aero-engine maintenance technology and discloses a laser cladding repair method for damage to the turbine blade rim and rounded corner transition area. The method involves: comprehensively inspecting the damaged areas of the turbine blade rim and rounded corner transition area using inspection equipment; treating the damaged areas of the turbine blade rim and rounded corner transition area using mechanical grinding; using the rounded corner vertex as the origin, and based on the length L of the right-angle side to be clad in the rounded corner transition area and the length L1 to be clad on the rim end face, the weld width W of a single cladding layer, and the height h of a single cladding layer, using an interpolation method to clad the areas of the rounded corner transition area and the rim end face; and finally, using a grinding equipment to finely repair the rim end face and rounded corner transition area. This invention solves the problems of large deformation of the rim end face and rounded corner transition area caused by traditional repair methods, and low bonding strength between the repair layer and the substrate, which cannot meet the assembly requirements of the rim and the usage requirements under high temperature and high vibration conditions.

The invention relates to a cobalt superalloy having (in % by weight): carbon (C): 0.04%-0.05%; chromium (Cr): 21.0%-23.0%; tungsten (W): 7.2%-8.2%; titanium (Ti): 0.05%-0.14%; aluminum (Al): 1.7%-2.7%; tantalum (Ta): 2.9%-3.6%; yttrium (Y): 0.01%-0.03%; nickel (Ni): 11.5% bis 13.5%; hafnium (Hf): 0.45%-0.65%; cobalt (Co): 49.0%-53.0%; zirconium (Zr): up to 0.02%; boron (B): up to 0.0014%; silicon (Si): up to 0.018%.

The invention comprises a superalloy target wherein the superalloy target has a polycrystalline structure of random grain orientation, the average grain size in the structure is smaller than 20 μm, and the porosity in the structure is smaller than 10%. Furthermore, the invention comprises a method of producing a superalloy target by powder metallurgical production, wherein the powder-metallurgical production starts from alloyed powder (s) of a superalloy and comprises the step of spark plasma sintering (SPS) of the alloyed powder (s).

The invention relates to a method and a device for producing micro parts and micro components by means of additive manufacturing using micro laser sintering. A pulverulent agglomerating material (1) is applied in layers from a first powder reservoir (2), and the material is melted using laser light (4) after being applied. The invention is characterized in that the material (1) located in the first powder reservoir (2), which is open in the direction of a processing plane, is an agglomerating powder, the particle size (1) of which equals maximally 20 μm, and a force (F1) is applied to the agglomerating material (1) in the direction of a substrate (3) in a process region (100) at least during the application process, the agglomerating material (1) behaving as a fluid while being applied as a result of said force. For this purpose, the device has a horizontally aligned base (5) which is interrupted at least in some regions in a process region (100) above a substrate (3) and has a first powder reservoir (2) which is arranged above the base (5) in a horizontally movable manner over the base (5) and which comprises an interior for receiving a pulverulent agglomerating material (1), the particle size (1) of which equals maximally 20 μm, wherein the first powder reservoir (2) has a first means (6) in the interior thereof for applying a first axial force (F1), which acts in the direction of the base (5) and the substrate (3), to the pulverulent agglomerating material (1) located in the interior, the agglomerating material (1) behaving as a fluid in the process region (100) at least while being applied as a result of said force.

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 (330) 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). The abrasion portion (334) includes the lattice structure (400) having a lattice density that increases along a radial direction (42) of the machine (10). There is also provided a method (600) of manufacturing such a fluid flow machine (10).

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).

The present invention provides an alloy that is excellent in oxidation resistance and structure stability, having high strength and a significantly improved temperature capability, and that is suitable for turbine disk applications. The nickel-cobalt-based alloy of the present invention has a composition comprising: 15% by mass or more and 43% by mass or less of cobalt; 6% by mass or more and less than 12% by mass of chromium; 3% by mass or more and 9% by mass or less of tungsten; 1% by mass or more and 6% by mass or less of aluminum; 1% by mass or more and 8% by mass or less of titanium; 7% by mass or less of tantalum; 0.01% by mass or more and 0.15% by mass or less of carbon; 0.01% by mass or more and 0.15% by mass or less of boron; 0.01% by mass or more and 0.15% by mass or less of zirconium; and the balance nickel and inevitable impurities.

A pump assembly of an aircraft engine includes a pump housing, a first pump stage positioned in the housing and including two first stage rotating elements meshed at a first mesh location, and a second pump stage positioned in the housing and including two second stage rotating elements meshed at a second mesh location. The first stage rotating elements are retained at respective shafts by one or more woodruff keys, and at least one second stage rotating element of the two second stage rotating elements is secured at its respective shaft via a spline connection of the at least one second stage rotating element to the respective shaft.

The present invention relates to a method for extending a fatigue lifetime of a turbine blade (1) affected by pitting. A corrosion pit (2) is identified and treated by laser shock peening. In some embodiment, the blade (1) is treated by single shot LSP. In another embodiment, pattern strategy used for LSP is provided. The pattern of the single shock peening can be overlapping or nonoverlapping. In a preferred embodiment, the application of LSP is angle dependent.

A component includes a body and an additively manufactured (AM) metal coupon that includes a primary member having a member porosity between 2% to 50% open space volume to total volume of the primary member. The primary member includes a first portion distanced from a second portion by a gap. At least one thermal transfer structure extends between the first portion and the second portion across the gap and a portion has a thermal transfer structure porosity. A braze material couples the AM metal coupon in a coupon opening in the body and infiltrates the primary member based at least on a characteristic of the member porosity.

A titanium alloy having 1.50 to 7.00 wt. % aluminium, 3.00 to 5.00 wt. % vanadium, 1.00 to 3.00 wt. % molybdenum; 0.50 to 2.50 wt. % zirconium, 0.05 to 0.40 wt. % oxygen, 0.05 to 2.00 wt. % tin, 0.00 to 1.00 wt. % iron, 0.00 to 0.3 wt % silicon, 0.01 to 0.15 wt. % carbon, and 0.001 to 0.05 wt. % nitrogen; the balance being titanium and incidental elements and unavoidable impurities.

A method for manufacturing a component, in particular a turbomachine component, in the form of a plurality of superposed layers of a material, includes a step of supplying the material into a feed nozzle and a step of projecting energy towards the feed nozzle to cause the material passing through the latter to melt. The method further includes a step of cooling the feed nozzle with a cooling element during the step of projecting energy.

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 (330) 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). The abrasion portion (334) includes a filler (500) within the lattice structure (400), the filler comprising one or more selected from: a polyester, a polyimide, a cyanate ester, a siloxane, a polysiloxane or a polyepoxide. There is also provided a vehicle (200) comprising such a fluid flow machine (10). There is also provided a method (600) of manufacturing such a fluid flow machine (10).

The invention relates to a component for a turbomachine, in particular for an aircraft propulsion system, wherein the component (1, 2, 3) is made of alloy powder and has a first region (11) and at least one second region (12, 13, 14), a material for a component (1, 2, 3) for a turbomachine, in particular for an aircraft propulsion system, and a manufacturing method for such a component.

The invention relates to a method for controlling the permeability of a lubrication circuit for an aircraft turbomachine, the turbomachine comprising at least one guide bearing, the lubrication circuit being configured to circulate an oil flow from upstream to downstream between an oil inlet and an oil outlet for lubricating the guide bearing, the method comprising the steps of: measuring a first temperature (T1) of the oil upstream of the guide bearing; measuring a second temperature (T2) of the oil downstream of the guide bearing; calculating a temperature difference between the second temperature (T2) and the first temperature (T1); comparing the temperature difference (ΔT) with a predetermined expected temperature difference (ΔT0); and, when the temperature difference (ΔT) is greater than the expected temperature difference (ΔT0), signalling a permeability fault in the lubrication circuit.

An arrangement (1) for the rear impact cooling of an impact plate (2) thermally stressed on its front side (3) by a hot medium (4) comprises the impact plate (2), a perforated plate (6) arranged at a distance from the rear side (5) of the impact plate (2), cooling fluid passage holes in the perforated plate (6) directed towards the rear side (5) of the impact plate (2), a cooling fluid outflow channel (9) formed between the rear side (5) of the impact plate (2) and the perforated plate (6), open in a transverse flow direction (10) at a cooling fluid outlet opening, and a cooling fluid supply channel (13) to which the cooling fluid passage holes are fluidically connected. The cooling fluid supply channel (13), the cooling fluid passage holes, and the cooling fluid outflow channel (9) form a cooling fluid passage through the arrangement (1) that is closed up to the cooling fluid outlet opening.To increase heat transfer from the baffle plate (2) to a cooling fluid (15) flowing through the cooling fluid passage, fluidic oscillators (19) are arranged in the cooling fluid passage.