1706results about "Leakage prevention" patented technology

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 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 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 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.

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.

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.

The invention provides an integrated turbine outer ring sector block structure combined by metal and ceramic components, and belongs to the technical field of aero-engines. The integrated turbine outer ring sector block structure comprises a metal supporting sector section fixedly connected with a turbine case shell; the outer side of the ceramic-based outer ring sector section is provided with a front mounting edge and a rear mounting edge, the ceramic-based outer ring sector block is arranged in an inner cavity of the metal supporting sector block in the radial direction, and the metal supporting sector block and the ceramic-based outer ring sector block are limited in the radial direction and the circumferential direction through a front pin shaft and a rear pin shaft which are arranged in the axial direction; the damping reeds are arranged in the inner cavities of the metal supporting sector blocks, convex structures are formed on the damping reeds, and the convex structures are in elastic pre-pressing fit with the cylindrical surfaces of the ceramic-based outer ring sector blocks; the front clamping ring is arranged in a plug pin hole of the metal supporting sector block; the elastic sealing ring is arranged between the metal supporting sector block and the front side of the ceramic-based outer ring sector block, and the ceramic-based outer ring sector block is pre-pressed on the rear end face of the metal supporting sector block; and the rear sealing piece is arranged on the rear pin shaft and is used for realizing axial limiting and gap sealing of the rear bolt.

A gas turbine engine includes a turbine section, a turbine case, an inner case, and a damping seal baffle. The turbine section extends along a rotational axis of the gas turbine engine. The turbine case includes a cantilevered inner wall. The cantilevered inner wall extends circumferentially about the rotational axis. The cantilevered inner wall extends between and to an upstream axial end and a downstream axial end. The upstream axial end is disposed at and downstream of the turbine section. The cantilevered inner wall forms a first sealing surface and a second sealing surface. The inner case extends and the damping seal baffle extend circumferentially about the rotational axis. The damping seal baffle is mounted to the inner case. The damping seal baffle is disposed in contact with one or both of the first sealing surface or the second sealing surface.

A turbine engine rotor has a central shaft and a disk stack having a plurality of disks encircling the shaft. A seal has a first member having, 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. The seal has a second member having, in central axial cross-section: a forward-facing surface contacting an aft-facing surface of the shaft; and a rear-facing surface contacting the forward surface of said portion extending radially inward from the sleeve.

The invention belongs to the technical field of environmental barrier coatings, and particularly relates to an environmental barrier coating for an outer ring block of an engine and a preparation method of the environmental barrier coating. The environmental barrier coating sequentially comprises a silicon layer, a double-ytterbium silicate layer, a single-ytterbium silicate layer, a transition layer and an abradable layer from inside to outside, and each of the transition layer and the abradable layer comprises a mixed oxide containing aluminum oxide, lanthanum oxide, magnesium oxide and neodymium oxide. The mass ratio of aluminum oxide to lanthanum oxide to magnesium oxide to neodymium oxide in the transition layer is (50-70): (15-25): (10-15): (8-10), and the mass ratio of lanthanum oxide to magnesium oxide to neodymium oxide in the abradable layer is (40-60): (20-30): (15-20): (5-8). The environmental barrier coating has significantly improved abradability, corrosion resistance, life in high temperature service environments.

A brush seal for use with a ceramic matrix composite (CMC) component of a gas turbine engine may feed a cooling flow to the CMC component, which has a hot side configured for exposure to a hot gas path of the gas turbine engine and an opposing cold side having a cavity for feeding a cooling flow to at least one film cooling hole of the CMC component. The brush seal includes: an outer diameter (OD) backing plate; an inner diameter (ID) backing plate; a plurality of brush seal bristles sandwiched between the OD backing plate and the ID backing plate; at least one hole forming a passage 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 configured to extend into the cavity of the CMC component.

The application discloses a turbine rim sealing gap passive adjustment structure and a gas turbine with the same. The turbine rim sealing gap passive adjustment structure comprises a casing, a rotating disc, a static disc, a rotating disc sealing ring arranged on an end surface of the rotating disc, a static disc sealing ring arranged on an end surface of the static disc, a containing groove arranged on the static disc, the rotating disc sealing ring extending into the containing groove, a first gap formed between the rotating disc sealing ring and the static disc sealing ring, an annular platform arranged on an inner circumferential surface of the containing groove, a second gap formed between the annular platform and the rotating disc sealing ring, a moving blade, a static blade, an adjustment cavity arranged on the end surface of the static disc, an adjustment chamber arranged in the adjustment cavity, a plurality of grids arranged on the adjustment cavity, the adjustment chamber and the containing groove being communicated through a guide hole, a jet hole arranged between every two adjacent grids, and a third gap formed between the grid and the rotating disc sealing ring. The turbine rim sealing gap passive adjustment structure has the advantages of good sealing effect and the like.

Powder compositions are having core-shell structures for use in forming environmental barrier coatings (EBCs) and / or abradable coatings by atmospheric plasma spraying. The shell compositions and thicknesses are selected to provide inner core particle (20) (silicate or phosphate particle) protection from the plasma and plume environments during atmospheric plasma spraying and avoid undesired modification to the particle chemistry during the deposition process. Shell coats (10) can be designed to survive the plasma and plume environments during atmospheric plasma spraying. Alternatively, shell coats (10) can be designed to be consumed during atmospheric plasma spraying.

A method for laser treating a substrate of a piston seal ring (PSR) (20) for a gas turbine engine includes: deforming the substrate to place a portion (310) of a first surface (30) of the substrate in tensile stress; and while holding the substrate deformed, laser treating the portion (310). After the laser treating but before release of the deformation, the first surface (30) may remain in tensile stress. Release of the deformation then at least partially cancels the tensile stress.

A turbine shroud assembly adapted for use with a gas turbine engine includes a carrier segment, a blade track segment, and a seal system. The carrier segment arranged circumferentially at least partway around an axis. The blade track segment is coupled to the carrier segment and defines a portion of a gas path of the gas turbine engine. The seal system is arranged radially between the carrier segment and the blade track segment to block gases from flowing between the carrier segment and the blade track segment.

A gas turbine engine includes a compressor section and a turbine section in axial flow arrangement defining an axially extending, longitudinal centerline, and arranged as a rotor and a stator. A seal assembly defines a primary seal at an interface of the rotor and the stator and separates an inlet plenum from an outlet plenum. The seal assembly includes one or more aspiration conduits fluidly connected to the primary seal. One or more flingers are positioned at least partially between the seal assembly and the inlet plenum that define one or more passageways fluidly connecting the inlet plenum to the one or more aspiration conduits. The flingers are positioned to direct at least a portion of an airflow within the inlet plenum away from the one or more passageways.

A seal assembly for a gas turbine engine may include a seal carrier extending about an assembly axis. A contacting-type seal member may be secured along the seal carrier. At least one spring carrier may extend from the seal carrier. At least one spring member may be received in a spring pocket of the respective at least one spring carrier. The at least one spring member may be arranged to bias a seal face of the seal member against a seal land at an interface. The at least one spring carrier may extend past the seal face relative to the assembly axis. A method of sealing for a gas turbine engine is also disclosed.

A gas turbine engine disposed within a nacelle is provided that includes an open rotor propulsion system, compressor, combustion, and turbine sections, a lubrication system, and a turbine active clearance control system. The turbine active clearance control system includes first and second valves, a heat exchanger, and a nacelle air inlet device. The heat exchanger receives air from the air inlet device, accepts a lubricant flow therethrough, and permits heat transfer between the air flow and the lubricant to produce a conditioned air flow. The first valve receives conditioned air flow from the heat exchanger and may pass the conditioned air flow to the second valve. The second valve receives a second air flow from the air inlet device. The turbine active clearance control system may operate in a heating mode in which conditioned air flow, or a combined air flow is provided to the turbine case clearance control segment.

A turbine engine is provided. The gas turbine engine defines a radial direction, and includes: a rotor; a stator comprising a carrier; a seal assembly disposed between the rotor and the stator, the seal assembly comprising a seal segment having a seal face forming a fluid bearing with the rotor; and a seal support assembly comprising a torsional spring extension extending from the carrier, from the seal segment, or both, to bias the seal segment along the radial direction.

The disclosure describes a ceramic matrix composite (CMC) component having a cooling cavity. The CMC component has a base having a bottom inner surface and a top outer surface, in which the base is made from a plurality of ceramic fiber plies within a matrix. A cooling cavity is provided within the base in which the cooling cavity has an opening in the top outer surface of the base. A cover plate made from one or more ceramic fiber plies within a matrix is provided to close the opening of the cooling cavity. The cover plate is bonded to the base by simultaneous densification of the base and cooling cavity.

A compressor is provided for use in domestic appliances. The compressor includes: an impeller including an inlet, an outlet, a hub defining a rotational axis of the impeller, a plurality of impeller blades that extend from the hub, and a shroud that at least partially surrounds the impeller blades; a housing that at least partially surrounds the shroud and that is coaxially arranged with respect to the rotational axis of the impeller to allow the impeller to rotate within the housing; a leakage flow path defined between the shroud and the housing leading from a leakage flow path inlet to a reinjection port, and a labyrinth seal arrangement provided in the leakage flow path. The labyrinth seal arrangement includes one or more shroud rings interdigitated with one or more housing rings so as to form a tortuous flow path for fluid flowing through the leakage flow path.

A gas turbine engine includes a compressor section, a combustion section, and a turbine section arranged in serial flow order and defining a working gas flowpath, a compressor of the compressor section comprising an aft-most compressor stage; a stage of stator vanes located downstream of the aft-most compressor stage; a stator case including a seal pad; and a spool drivingly coupled to the compressor, the spool and the stator case together defining a rotor cavity in fluid communication with the working gas flowpath, the spool comprising a seal tooth assembly, the seal tooth assembly including a seal support extension, a seal tooth extending from the seal support extension toward the seal pad, and a dampener operable with the seal support extension.