943results about "Leakage prevention" patented technology

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 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 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.

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.

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 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.

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.

The utility model relates to the technical field of steam power generation turbines, in particular to a mixed flow type saturated steam power generation turbine which comprises a cylinder, the cylinder comprises a vertical inlet cavity and a horizontal outlet cavity which are communicated, and reinforcing ribs are arranged at the positions, corresponding to the joint of the two cavities, of the inner wall of the cylinder. A protruding structure at the bottom of the side, away from the horizontal outlet cavity, of the outer wall of the air cylinder is fixedly connected with a gearbox. The stationary blade assembly comprises a first-stage radial flow stationary blade fixedly arranged on the inner wall of the vertical inlet cavity and a second-stage axial flow stationary blade fixedly arranged on the reinforcing ribs; the rotor assembly comprises a rotating part connected with the high-speed shaft of the gear box, a first-stage radial flow vane impeller and a second-stage axial flow vane impeller, wherein the periphery of the rotating part is sleeved with the first-stage radial flow vane impeller and the second-stage axial flow vane impeller. The radial flow turbine and the axial flow turbine are combined to form the mixed-flow turbine, so that the working capacity of a single cylinder is improved, the power generation efficiency is improved, the cost is saved, and the unit size is reduced.

A sealing arrangement, for a turbine with a static casing 221 and a runner 222 that rotates relative to the casing, comprises a plurality of parallel dead-end passages 228a, 228b formed within the run

The invention provides a power plant steam turbine shaft end steam leakproof device. The power plant steam turbine shaft end steam leakproof device comprises a connecting pipe connected to the steam turbine shaft end, an air inlet pipe used for conveying steam to a steam turbine and a sealing connecting assembly used for axially fixing the connecting pipe and the air inlet pipe. Wherein the connecting pipe is provided with a first end close to the air inlet pipe, and the inner wall of the first end is fixedly connected with a first sealing ring; the air inlet pipe is provided with a second end close to the connecting pipe, the inner wall of the second end is fixedly connected with a second sealing ring, the second sealing ring is fixedly connected with an integrally-formed sealing pipe extending out of the end face of the air inlet pipe, and the sealing pipe is fixedly connected with an integrally-formed third sealing ring. The third sealing ring and the first sealing ring can be buckled and matched in a sealing mode. The power plant steam turbine shaft end steam leakproof device can effectively solve the fundamental problem that existing flange connection is prone to leakage due to lack of active sealing compensation and dependence on gasket and bolt pre-tightening.

A seal assembly (60A, 60B; 160B) includes a carbon seal (64A, 64B; 164B) that has a sealing surface (66B, 66B; 166B). A seal seat (62A, 62B; 162B) has a sealing surface (66B, 68B; 168B) and is positioned for rotation relative to the carbon seal (64A, 64B; 164B). A diamond-like carbon coating (80) at least partially forms the sealing surface (68B, 68B; 168B) on the seal seat (62A, 62B; 162B).

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.

A seal assembly for rotary machines having a wear protection assembly with an abradable covering. The wear protection assembly has one or more wear protection features that enable in-operation opening of the seal assembly by increasing a clearance between a seal rotor and a seal slider of the seal assembly. In particular, the wear protection assembly is configured to open the seal assembly and increase the clearance of the seal rotor and the seal slider during slider air-bearing wear conditions, slider counterbore wear conditions, rotor air-bearing wear conditions, or extreme vibration wear conditions. Under such or similar wear conditions (or under conditions preceding or leading up to a wear condition), the wear protection assembly yields a pressure gradient and / or a fluid flow that oppose the forces acting to decrease the clearance between the seal rotor and the seal slider.

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 invention relates to the technical field of gas turbine sealing, and particularly provides an inner-cooling type pneumatic sealing honeycomb structure based on tailor-welding and a machining method of the inner-cooling type pneumatic sealing honeycomb structure. The honeycomb structure comprises a plurality of honeycomb units which are spliced and welded with one another; each honeycomb unit comprises two welding units which are integrally formed; each welding unit comprises a node section and free sections symmetrically arranged at the two ends of the node section in the first direction. A first groove is formed in the surface of the first end, away from the free section, of the node section; the two welding units in the honeycomb unit are in mirror symmetry in the first direction, and the first ends of the two welding units are correspondingly welded; the node sections of the two honeycomb units in mirror symmetry in the first direction are correspondingly welded, and the two first grooves of the two welded node sections are spliced to form a first cooling channel. The problems that when the gas turbine sealing structure is cooled, temperature gradient exists, safety risks are likely to be generated, the punching machining difficulty is large, and the manufacturing cost is high can be solved.

A method comprises: flowing a potted component in a liquid state over a tip (214) of an airfoil (210), the tip (214) of the airfoil (210) having a coating (220) disposed thereon, the coating (220) comprising a metal plating (221) and a plurality of protrusions (222), each protrusion in the plurality of protrusions (222) extending from the metal plating (221); allowing the potted component to harden to form a hardened potted component; and removing the hardened potted component from the tip (214) of the airfoil (210).

The application discloses a stator component and a gas turbine, and relates to the technical field of gas turbines, in particular to a stator component and a gas turbine. The stator component comprises a stator ring, a sealing element and an elastic element. The stator ring comprises a plurality of arc segments connected in sequence along the circumference of the stator ring. Two adjacent arc segments are a first arc segment and a second arc segment. The first arc segment is provided with a mounting groove at the end of the second arc segment. The opening of the mounting groove faces the second arc segment. The sealing element is arranged in the mounting groove. The sealing element abuts against the end surface of the second arc segment facing the first arc segment. The elastic element is arranged in the mounting groove. The elastic element is used for providing the sealing element with an elastic force facing the second arc segment. The stator component of the application can be maintained without affecting the sealing performance even if the sealing element is deformed, so that the sealing element can be repeatedly used, thereby reducing the maintenance cost of the stator component.

An assembly is provided for a turbine engine (20). This assembly includes a carrier (80), a control ring (82) and a bladed outer air seal (78) mounted to the carrier (80). The carrier (80) extends circumferentially around an axis (22). The carrier (80) includes an annular cavity (114), a first section (106) and a second section (108). The annular cavity (114) is formed within the carrier (80) by the first section (106) and the second section (108). The first section (106) is welded and / or brazed to the second section (108) at a sealed full-hoop bond joint (148; 150). The first section (106) is further connected to the second section (108) at a second sealed full-hoop joint (148; 150). The control ring (82) extends circumferentially around the axis (22) within the annular cavity (114). A method of manufacture is provided for an assembly for a turbine engine (20).

Gas turbine engines include an engine core having compressor, combustor, and turbine sections arranged within an engine casing. A transfer tube (404) is arranged on an exterior of the casing and configured to define a fluid conduit for fluid (410) to be transferred about the engine. A fitting (402) is connected to the transfer tube (404) and a seal assembly (406) is arranged between the fitting (402) and the transfer tube (404). The seal assembly (406) provides a fluid seal between the fluid conduit and an ambient environment (412) at a connection between the fitting (402) and the transfer tube (404). The seal assembly (406) includes a primary seal (414) arranged in a gap (413) between the fitting (402) and the transfer tube (404) and a secondary seal (416) arranged in the gap (413) between the primary seal (414) and an ambient environment (412). The secondary seal (416) is configured to expand into engagement with the fitting (402) and the transfer tube (404) when exposed to temperatures of at least 800 °F.

A primary vane segment has: an outer ring extending in a circumferential direction; and a plurality of primary vanes arranged and installed in the circumferential direction on a radially inner side of the outer ring. The primary vanes have: a vane body extending in the radial direction; and an outer shroud formed on a radially outer side of the vane body with respect to the axis. The outer shroud has: a shroud main body; and a buried portion connected to the radially outer side of the shroud main body. The buried portion has: a buried front end surface facing an upstream side of the axis; and a buried rear end surface facing a downstream side of the axis. The outer ring has an outer ring groove into which the outer shroud of each of the plurality of primary vanes enters. The buried front end surface is located at a position more on the upstream side of the axis than a leading edge of the vane body. The buried rear end surface is located at a position more on the upstream side of the axis than a trailing edge of the vane body.

A split ring seal has: a first circumferential end and a second circumferential end; an inner diameter surface and an outer diameter surface; a first axial end face and a second axial end face. The outer diameter surface has a sealing surface. The first axial end face has: a first section; a second section outboard of the first section; and an axial protrusion between the first section and the second section The second axial end face has: a first sealing surface; a second sealing surface radially outboard of the first sealing surface; a circumferential channel between the first surface and the second sealing surface; and a plurality of channels extending radially outward from the circumferential channel.

A gas turbine engine (20) includes a ceramic wall (62) for bounding an engine core gas path (C). The ceramic wall has a ceramic wall first side (64a) that faces the engine core gas path and a ceramic wall second side (64b) that faces away from the engine core gas path. There is a metallic wall (72) adjacent the ceramic wall second side. The metallic wall has a metallic wall first side (72a) that faces the ceramic wall and a metallic wall second side (72b) that faces away from the ceramic wall. The metallic wall and the ceramic wall are spaced apart such that there is a channel (76) there between. There is a seal (78) on the ceramic wall second side, and the metallic wall has at least one cooling hole (80) adjacent the seal for emitting cooling air to cool the seal.

A seal assembly for an industrial transmission includes a vertically aligned lower power shaft to exchange torque with the industrial transmission, and a transmission component arranged radially outside the power shaft. The power shaft has a greater rotational speed than the transmission component during operation. A contactless first seal between the transmission component and the power shaft includes a conical first sealing gap, which is beveled in a direction of gravity and runs from radially outside to radially inside counter to the direction of gravity. A first overflow channel runs downward to at least a large extent in the direction of gravity and communicates with the first sealing gap. The first overflow channel ends below the first sealing gap in the direction of gravity and is embodied as an annular gap with a substantially constant flow cross section such as to avoid formation of a labyrinth seal.

Device (1) for characterizing wear on a gas turbine engine comprising: - an abradable material (8) having a wear end (80) adapted to be in contact with a gaseous fluid flow (F) flowing from upstream to downstream of the gas turbine engine when the gas turbine engine is running, the abradable material (8) being electrically conductive; - a magnetic field generator (7) adapted to emit a magnetic field in the abradable material (8); - an electronic board (10) configured to calculate information enabling the characterization of the wear of the gas turbine engine from data relating to a physical response of said abradable material (8) to said magnetic field. Figure for the abbreviation: Figure 2.