2759results about "Leakage prevention" patented technology

A component (202) is disclosed that includes a body (206) and an additive manufactured (AM) metal mass (200) that includes a first section (292) and a stress relief section (294). The stress relief section (294) includes a porous region (300) having a porosity in which an open spatial volume accounts for 2% to 50% of the total volume of the stress relief section (294). The stress relief section (294) has a smaller at least one of thermal conductivity, tensile strength, ductility, and fatigue strength than the first section (292). A braze material (310) couples the AM metal specimen block (200) to a specimen block opening (204) in the body (206). The stress relief section (294) creates a low stress zone in the metal specimen block (200), which can be located near or adjacent to a high stress zone. The stress relief section (294) reduces residual stress in the metal sample block (200), which reduces related challenges to couple the metal sample block (200) to the body (206) of the component (202), particularly in the case of the use of superalloy materials.

A component including a body, and an additively manufactured (AM) metal coupon including a first section and a stress-relief section, is disclosed. The stress-relief section includes a porous region having a porosity between 2% to 50% open space volume to total volume of the stress-relief section. The stress-relief section has at least one of thermal conductivity, tensile strength, ductility, or fatigue strength less than first section. A braze material couples the AM metal coupon to a coupon opening in the body. The stress-relief section creates a low-stress zone in the metal coupon that can be located near or adjacent to a high-stress zone. The stress-relief section reduces residual stress in the metal coupon, which reduces the related challenges of coupling the metal coupon to the body of the component, especially where superalloy materials are used.

The present invention relates to an assembly (12) for a turbomachine, comprising: —an outer casing (15) made of half-shells; —an inner casing (14); —a plurality of guide vanes (13) mounted between the outer casing (15) and the inner casing (14), —a structural casing comprising a flange (23) which is radially inside the inner casing (14); and—a system (24) for blocking the inner casing (14) on the structural casing, comprising a generally annular tab (25) fixed to one of the inner casing (14) and the structural casing and a complementary chute (26) fixed to the other of the inner casing (14) and the structural casing, the chute (26) being configured to receive the tab (25) and to block it along the axis (X) with respect to the structural casing.

The invention relates to the technical field of aero-engines, in particular to a sealing and cooling structure and method for adjusting the axial force of an aero-engine rotor, which comprises a turbine disc provided with a rear baffle, and further comprises a low-pressure turbine guide vane arranged opposite to the turbine disc, a low guide front inner supporting ring and a low guide rear inner supporting ring are arranged in the middle of the low-pressure turbine guide vane; labyrinth teeth correspondingly matched with the low-guide front inner supporting plate are formed on the rear baffle, radial labyrinth tooth gaps are formed, and when airflow in an air collection cavity enters a rotor cavity from a low-pressure pre-rotation nozzle, the airflow is embedded into an air supply channel from a front pre-rotation hole at the same time; one part of airflow entering the air supply channel enters the disc rear cavity, and the other part of airflow enters the rotor cavity. By adjusting the sealing labyrinth structure, it is ensured that the axial force of the rotor does not exceed the allowable maximum value in the large-state working process and is not subjected to light load in the small-state working process, meanwhile, sealing gas used behind a turbine disc is ensured, turbine hot gas is prevented from flowing into an inner disc cavity from a main channel, and it is ensured that an engine works reliably.

An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.

The invention relates to a magnetic suspension air separation expansion machine system and an operation method thereof. The system comprises a base, a magnetic suspension bearing assembly, a rotor assembly and a control module. The magnetic suspension bearing assembly is arranged on the base, comprises an expansion end radial magnetic suspension bearing, a brake end radial magnetic suspension bearing and an axial thrust magnetic suspension bearing, and is used for supporting radial and axial forces of the rotor assembly; the rotor assembly is arranged in the machine body and connected with the machine body through the magnetic suspension bearing assembly, and non-contact suspension is achieved. The control module is electrically connected with the magnetic suspension bearing assembly and used for controlling the magnetic suspension bearing assembly to suspend the rotor. The operation method comprises the steps of self-checking before starting, electrifying to suspend the rotor, starting the air compressor to pressurize, adjusting flow to drive the expansion impeller, carrying out split-flow cooling on gas, and collecting data in real time to maintain stable suspension of the rotor. The device has the advantages of high efficiency, stability, energy conservation and the like and is suitable for the field of air separation.

A slurry for use to form or repair a silicon carbide coating is provided. In one aspect, the slurry includes solid particles and a carbonaceous resin. The solid particles include silicon carbide particles, silicon particles, and carbon particles. A method of fabricating a silicon carbide coating is also provided. In one aspect, the method includes applying the slurry, heating the slurry, and forming the silicon carbide coating from the solid particles and the carbonaceous resin. A method of repairing a silicon carbide coating is also provided. In one aspect, the method includes applying the slurry to a damaged region of the silicon carbide coating, heating the slurry, and repairing the silicon carbide coating with the solid particles and the carbonaceous resin in the damaged region.

A system includes a gas turbine engine, including a compressor, a combustor, and a turbine. The system includes an intercooler configured to receive a discharge air from the compressor and to cool the discharge air. The system includes a first conduit to divert a portion of the cooled discharge air from the intercooler or a location downstream of the intercooler to a bearing located within the turbine to pressurize the bearing to block leakage of fluid from the bearing.

An apparatus for an aircraft powerplant includes an impeller rotor and a seal runner. The impeller rotor includes a plurality of fluid circuits. Each of the fluid circuits includes a first outlet passage, a second outlet passage and an inlet passage fluidly coupled to the first outlet passage and the second outlet passage in parallel. The seal runner includes a bore, an inner guide surface and an outer land surface. The seal runner extends circumferentially about the axis and radially between the inner guide surface and the outer land surface. The bore is fluidly coupled to the first outlet passage of each fluid circuit. The inner guide surface forms a radial outer peripheral boundary of the bore. The inner guide surface radially diverges away from the axis as the inner guide surface extends axially away from the impeller rotor and to a distal end of the seal runner.

The invention discloses a cylindrical high-pressure inner cylinder steam seal structure which is characterized in that a shrinkage fit ring is arranged outside a high-pressure inner cylinder upper half and a high-pressure inner cylinder lower half in a sleeving manner, and the high-pressure inner cylinder upper half and the high-pressure inner cylinder lower half are combined based on a horizontal split; two steam leakage grooves are respectively formed in the horizontal split surfaces of the high-pressure inner cylinder upper half and the high-pressure inner cylinder lower half; two annular grooves are formed in the end steam seal positions of the high-pressure inner cylinder upper half and the high-pressure inner cylinder lower half, and two steam seal cavities are formed respectively. Two through-flow cavities are formed in the through-flow middle positions of the high-pressure inner cylinder upper half and the high-pressure inner cylinder lower half; the steam seal cavity is communicated with the through-flow cavity through the corresponding steam leakage grooves, high-temperature and high-pressure leaked steam of the end steam seal flows back to the through-flow, and a small amount of low-temperature and low-pressure steam leaks out of the end steam seal. According to the structure disclosed by the invention, the steam leakage amount is reduced, and the heat consumption is reduced.

A containment ring assembly for a turbine casing assembly, including: a first containment ring; a second containment ring, the first containment ring being radially inward from the second containment ring, wherein a radial space is located between the first containment ring and the second containment ring, the radial space allowing deformation of the first containment ring prior to the first containment ring contacting the second containment ring.

A gas turbine engine includes a rotor that that has a seal surface, a shaft that has an annular seal channel that opens to the seal surface, and a seal disposed in the annular seal channel for sealing against the seal surface. The seal is made of a composite having carbon fibers disposed in a carbon matrix.

The embodiment of the invention discloses a labyrinth sealing assembly, a gas turbine and an active clearance control method.The labyrinth sealing assembly comprises a stator casing and a rotor, an annular deformation part is arranged on the inner circumferential wall of the stator casing, and the central axis of the deformation part and the central axis of the stator casing are coaxially arranged; the deformation part can actively deform in the radial direction of the stator casing, the rotor is provided with at least two labyrinth teeth, the position of at least one labyrinth tooth corresponds to the deformation part in the radial direction of the stator casing, the labyrinth teeth are defined as active control labyrinth teeth, and the active control labyrinth teeth are consistent with other labyrinth teeth in outer diameter and cold-state gap value. The gap between the deformation part and the active control comb teeth is adjusted through deformation of the deformation part. The labyrinth sealing assembly provided by the embodiment of the invention is mainly used for reducing leakage between the rotor and the stator and improving the efficiency of the gas turbine. Through active adjustment of the deformation part, the minimum leakage gap can be kept, and therefore energy loss is reduced.

The reverse hedging type steam seal device of the steam turbine comprises a steam seal body, high teeth and low teeth are arranged on an inner ring of the steam seal body in the radial direction, and an air seal groove channel is formed in the side, facing the low-pressure end, of the high teeth. A main cavity is formed in the air seal channel, an energy consumption cavity is formed in the left side of the main cavity, a hedging cavity is formed in the right side, and a cavity inlet is formed in the bottom. After entering the main chamber, the steam main flow leftwards forms a first steam branch flow to enter the energy consumption chamber, and rightwards forms a second steam branch flow to enter the hedging chamber; an energy consumption structure is arranged on the inner wall of the energy consumption chamber, and the energy consumption chamber is provided with an energy consumption channel for communicating the first steam branch flow with the steam main flow; the inner wall of the hedging chamber is smooth, the hedging chamber is provided with a reversely inclined hedging channel, and the second steam branch flow and the steam main flow form a reverse steam flow hedging effect through the hedging channel. According to the utility model, the leakproofness of the steam seal can be effectively improved, and the problem of movement of the rotor shaft is reduced on the basis.

The invention relates to the technical field of rotary mechanical sealing, in particular to a radial adjustable brush type sealing structure with a thermal response adjusting function, which comprises a plurality of arc-shaped sealing structures, and the end surfaces of the plurality of arc-shaped sealing structures are connected to form an annular brush type sealing structure; the arc-shaped sealing structure comprises an arc-shaped sealing case, an elastic supporting assembly and an axial limiting assembly. When the rotor radially jumps or the temperature changes, the arc-shaped sealing block can adaptively move in the guide limiting assembly to adaptively adjust the sealing gap between the brush wire bundle and the rotor, so that the sealing performance is improved, and the service life of the brush wires is prolonged.

A hydraulically supported seal including a seal plate operatively coupled with a seal, the seal being supported by a seal carrier within a housing; a piston having a piston shaft, the piston being in operative communication with the seal opposite the seal plate, the piston shaft being disposed through the housing through a piston receiver formed in the housing; a cylinder operatively coupled to the piston opposite the seal; and wherein the piston is hydraulically biased by a hydraulic fluid within the cylinder.

A containment ring assembly for a turbine casing assembly, including: a first containment ring; a second containment ring, the first containment ring being radially inward from the second containment ring, wherein a radial space is located between the first containment ring and the second containment ring, the radial space allowing deformation of the first containment ring prior to the first containment ring contacting the second containment ring.

An apparatus and method for radiation measurement are used to determine clearance of a rotating gas turbine component. The apparatus includes a probe body on a stationary component outward of the rotating component and a pair of sensor assemblies coupled to the probe body. Each sensor assembly includes a plurality of prisms coupled to the probe body Each prism of the plurality of prisms has a distinct angular orientation to direct radiation from the focusing lens to the rotating component surface off-axis with respect to an axis of the respective radiation source of the pair of radiation sources and with less directional variation than radiation from the respective radiation source.

A ceramic matrix composite (CMC) blade outer air seal (BOAS) assembly includes a CMC BOAS including a base and a pair of rails having a substantially π-shaped cross-section. The pair of rails include a first pair of aligned holes at a first end and second pair of aligned holes at a second end. A first pair of parallel metallic pins have a first end connected to a first central support pin and a second end connected to a second central support pin, respectively, wherein the first pair of parallel metallic pins are configured to engage the first pair of aligned holes. A second pair of parallel metallic pins also have a first end connected to a first central support pin and a second end connected to a second central support pin, respectively, wherein the second pair of parallel metallic pins are configured to engage the second pair of aligned holes.

A cooled brush seal for use with a CMC component of a gas turbine engine having a hot side exposed to a hot gas path and an opposing side having a cavity for feeding a cooling flow to at least one film cooling hole. The cooled brush seal includes: an OD backing plate; an ID backing plate; a plurality of brush seal bristles sandwiched between the OD backing plate and the ID backing plate; a first hole extending through the OD backing plate, the brush seal bristles, and the ID backing plate; and a cover plate attached to an inner surface of the ID backing plate and including a second hole circumferentially offset from the first hole and forming a second passage into the cavity, wherein the ID backing plate and the cover plate form a circumferential cooling passage extending between the first hole and the second hole.

A blade and a sealing element forming an assembly for a turbomachine turbine. The blade includes an airfoil and a platform to which the sealing element is connected. An internal channel passes through the platform so as to draw off fluid circulating in a primary duct that it delimits and to inject the thus drawn-off fluid into an inter-lip cavity delimited by the sealing element.

The invention discloses an air suspension turbine power generation and refrigeration integrated device which comprises a shell, a stator and a rotor spindle, the rotor spindle is rotationally arranged in the shell through a bearing assembly, the stator is fixed in the shell and is coaxial with the rotor spindle, and an air gap exists between the stator and the rotor spindle to form a permanent magnet synchronous generator. The rotor further comprises a first-stage volute, a second-stage volute, a first-stage radial flow turbine, a second-stage radial flow turbine, a volute connecting pipe and a cooling water jacket, the first-stage volute and the second-stage volute are fixed to the two ends of the shell respectively, and the first-stage radial flow turbine and the second-stage radial flow turbine are coaxially fixed to the two ends of the rotor spindle respectively and are arranged back to back; the first-stage radial flow turbine is located in the first-stage volute, the second-stage radial flow turbine is located in the second-stage volute, the first-stage volute is provided with a high-temperature and high-pressure air inlet and a first-stage air outlet, and the second-stage volute is provided with a second-stage air inlet and an expansion cooling air exhaust port; the device has the advantages of being capable of stably running at a high speed in a high-altitude low-pressure environment and simultaneously achieving efficient power generation and refrigeration functions.

The invention discloses a working medium self-recovery organic working medium turbine without a horizontal split. The organic working medium turbine comprises a cylinder without the horizontal split, a whole-circle partition plate group and a rotating part, the rotating part is divided into three parts along the axial direction; the whole-circle partition plate set is divided into a plurality of whole-circle partition plates arranged at intervals in the axial direction. The whole circle of partition plates and the three parts are alternately assembled in the air cylinder without the horizontal split in a vertical installation mode, and the three parts are connected into a whole after being alternately assembled. The rotating part is provided with a sealing arm and a sealing shoulder to form a balance cavity, and the pressure of the cavity is set by adopting a balance pipe; the rotating part is connected with a mechanical seal at the shaft extension part of the air cylinder without the horizontal split; according to the mechanical seal, cold spacer fluid is adopted to wash a movable and static ring friction pair, and meanwhile, the cold spacer fluid is adopted to spray a gas-liquid mixing cavity to achieve cooling, and the spacer fluid of the mechanical seal leaks to a balance cavity, is isolated by a centrifugal pressurizing ring on a rotating part and is separated and recycled in a gas-liquid header. According to the invention, external zero leakage and self-recovery of the organic working medium are effectively realized.

Disclosed herein are example variable flowpath casings for blade tip clearance control. An example casing for a turbine engine includes a first annular substrate extending along an axial direction; a second annular substrate positioned radially inward relative to the first annular substrate, the second annular substrate movably coupled to the first annular substrate; and an actuator coupled to the second annular substrate such that a force applied by the actuator moves the second annular substrate relative to the first annular substrate to adjust a tip clearance.

A turbine engine rotor has a central shaft and a disk stack having disks encircling the shaft. A seal has in central axial cross-section: a rearwardly-open channel receiving a portion of one of the disks; a sleeve extending rearward from the channel and between the disk and the shaft; and a portion extending radially inward from the sleeve and having a forward surface restrained against forward movement relative to the shaft.

The invention relates to a method (400) for manufacturing a casing (102) of a turbine engine (10), the casing (102) comprising a skin (102a) having an annular shape about an axis (A) and comprising tubular vents (106) which project on an outer annular face (102e) of the skin (102a) and are oriented radially with respect to the axis, the method (400) comprising the following steps: - preparing (402) an annular skin made of a thermoplastic-based material, the skin comprising openings (102h) passing through in the radial direction, - preparing (404) tubular vents made of thermoplastic-based material, - attaching (406) the tubular vents to the skin, with the attachment step (406) comprising, for each vent (106): • welding (406a) a support (200) to the skin at one of the openings, and • welding (406b) a sleeve (202) to the support and / or the skin, with the assembly formed by the support and the sleeve forming the vent.

A cooling flow arrangement for a ceramic matrix composite (CMC) component of a turbine engine, wherein the CMC component includes a hot side configured for exposure to a hot gas path of the turbine engine and an opposing cold side. At least one cavity is provided in the CMC component, the at least one cavity having an entrance disposed on the cold side that is configured for receiving a cooling flow. A seal is disposed adjacent the at least one cavity and the cold side of the CMC component and is configured to provide a seal between the CMC component and an upstream turbine component. At least one hole extends through the seal and is positioned to provide the cooling flow to the at least one cavity of the CMC component.

A compressor of a gas turbine engine is provided. The compressor includes a casing including an elongate body having an aft surface at an aft end thereof and an outer air seal (OAS) disposed aft of the casing. The OAS includes at a forward end thereof a first forward surface, a second forward surface inboard of and recessed aftward from the first forward surface and a transitional surface radially interposed between the first and second forward surfaces. The aft surface of the casing is axially interposed between the first and second forward surfaces of the OAS.

An assembly may include a rim-rotor turbine having a hub adapted to be mounted or connected to a rotating shaft, a rim-rotor, the rim-rotor having a cooling ring defining at least one cooling channel, and a plurality of blades with each of the plurality of blades contacting the rim-rotor. A stator vane assembly is positioned upstream of the rim-rotor turbine, and has an outer shroud defining a main flowpath aligned with the plurality of blades of the rim-rotor turbine and configured to direct hot gas toward the plurality of blades of the rim-rotor turbine, and vanes in the main flowpath. A cooling guide vane assembly has an inner shroud defining a cooling flowpath aligned with the cooling ring of the rim-rotor turbine and configured to direct cooling gas toward the cooling ring of the rim-rotor turbine, and vanes in the cooling flowpath. The outer shroud of the stator vane assembly is radially inward of the inner shroud of the cooling guide vane assembly, and wherein a circumference gap is defined therebetween.

The invention provides an inter-turbine casing guider centering and sealing structure, and belongs to the technical field of aero-engine structural design, the inter-turbine casing guider centering and sealing structure comprises a guide blade, an outer ring, an inner ring, an inner ring fixing ring, a blanking cap, a bearing inner ring casing, a cover plate, a bearing seat and an inner ring centering ring; an upper edge plate of the guide blade is in centering fit with a front hanging point and a rear hanging point of the outer ring through the outer surface of a front hook and the inner surface of a rear hook; the inner ring is centered with a radial pin shaft of the guide blade; one side of the inner ring fixing ring is fixed on the inner side of the inner ring, and the inner ring centering ring is arranged between the other side of the inner ring fixing ring and the bearing inner ring casing, so that the centering of the whole guider structure is realized; the bearing seat is mounted on the inner side of the force-bearing inner ring casing, and a blanking cap is mounted on the bearing seat; and one side of the cover plate is mounted on the inner ring centering ring through a connecting piece, and the other side of the cover plate is connected with a blanking cap on the bearing seat.