1935results about "Stators" patented technology

A power turbine blade for a gas turbine engine includes at least one of a third-stage blade, a fourth-stage blade, or a fifth-stage blade. The power turbine blade includes a platform and an airfoil that extends in a spanwise direction from the platform to a tip shroud. The airfoil has an external airfoil surface formed in conformance with multiple cross-section airfoil profiles described by a set of Cartesian coordinates.

The invention provides a pneumatic design method for a high-load turbine rear rectification blade grid. The pneumatic design method comprises the steps of pneumatic design of a rectification supporting plate and pneumatic design of a meridian flow channel. The rectification supporting plates are high in load, the meridian flow channels are designed by being matched with blade-shaped channels of the rectification supporting plates under a certain number of rectification supporting plates, and the problem of large-scale separation flow caused by high load is solved from the aspect of meridian flow channel molded line design; the structure is compact, the total pressure loss is low, the weight of the whole machine can be reduced to a certain extent, and the manufacturability is improved; the method can be applied to the pneumatic design of the high thrust-weight ratio aero-engine turbine rear rectification cascade and the pneumatic design of the gas turbine engine core engine test turbine rear rectification cascade.

A gas turbine engine includes, among other things, a propulsor section including a rotor and blades, a geared architecture including a gear train, a compressor section and a turbine section. A static structure includes a first case and a second case. The first case includes a first engine mount location. The second case includes a second engine mount location. The first engine mount location is axially near the geared architecture.

An assembled diaphragm for steam turbine and a production device thereof are provided, which relate to the technical field of steam turbine manufacturing. The diaphragm includes diaphragm inner rings and diaphragm outer rings, the diaphragm inner rings and the diaphragm outer rings are semicircular, and blades are arranged between the diaphragm inner rings and the diaphragm outer rings, two the diaphragm inner rings are symmetrically connected through a connecting assembly, and two the diaphragm outer rings are symmetrically connected through multiple connecting assemblies. Two semicircular diaphragm inner rings are connected into a complete circular diaphragm inner ring through multiple connecting assemblies, and two semicircular diaphragm outer rings are connected into a complete circular diaphragm outer ring through multiple connecting assemblies, and then the blade is welded and fixed between the complete circular diaphragm inner ring and the circular diaphragm outer ring.

A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine comprising: a turbomachine comprising, in serial flow order, a compressor section, a combustion section, and a turbine section, the turbomachine defining a working gas flow path and a fan duct flow path; a primary fan driven by the turbomachine, the primary fan defining a primary fan tip radius R1 and a primary fan hub radius R2; a secondary fan downstream of the primary fan and driven by the turbomachine, at least a portion of airflow from the primary fan being configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R3 and a secondary fan hub radius R4, wherein the secondary fan is configured to provide, during operation, a fan duct airflow through the fan duct flow path to generate a fan duct thrust, wherein the fan duct thrust is equal to a %Fn of a total engine thrust during operation of the gas turbine engine at rated speed during standard day operating conditions 3S ; wherein a ratio of R1 to R3 is equal to wherein EFP is between 1.5 and 11, wherein RqR Prim.‑Fan is a ratio of R2 to R1, and wherein RqR Sec.‑Fan is a ratio of R4 to R3.

An aircraft engine, has: a turbine including a turbine rotor rotatable about a central axis; a scroll case having an inlet connected to a source of combustion gases and an outlet fluidly connected to the turbine, and a conduit extending around the central axis from the inlet to the outlet; a bearing housing; a turbine support case secured to the bearing housing, the scroll case disposed axially between the turbine support case and the bearing housing, the turbine support case having spokes extending through the scroll case and radially supported by the bearing housing; and heat shields distributed around the central axis in circumferential alignment with the spokes and extending axially through the scroll case for thermally shielding the spokes, a heat shield of the heat shields disposed around a spoke of the spokes, the heat shield defining an air gap surrounding the spoke inside the scroll case.

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.

An air / oil heat exchanger arranged in a vein. The exchanger includes parietal corridors and central passages, the radial height of the parietal corridors is less than the radial height of the central corridors at the inlet of the exchanger.

An assembly is provided for an aircraft powerplant. This assembly includes an electric machine, an electric machine controller, an electric cable, a conduit and an air circuit. The electric machine controller is configured to control operation of the electric machine. The electric cable electrically couples the electric machine controller to the electric machine. The conduit extends longitudinally from a first end of the conduit to a second end of the conduit. The electric cable extends longitudinally through a bore of the conduit. The air circuit is configured to cool the electric cable within the conduit using a flow of air. The air circuit includes the bore of the conduit.

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.

An assembly is provided for a turbine engine. This assembly includes an engine structure, and the engine structure includes a plenum, a reverse flow combustor, a nozzle structure and a cooling circuit. The reverse flow combustor is disposed in the plenum and includes a combustion chamber. The nozzle structure is arranged at an outlet from the combustion chamber. The nozzle structure includes a first platform, a second platform and a plurality of vanes arranged circumferentially about an axis. The vanes extend across the flowpath from the first platform to the second platform. The cooling circuit extends in the first platform and between a circuit inlet into the cooling circuit and a circuit outlet from the cooling circuit. The circuit inlet is fluidly coupled with and disposed along the plenum. The circuit outlet is fluidly coupled with and disposed along the flowpath.

A swirl recovery vane (SRV) includes a trunnion configured to pivotally connect to an exterior surface of an inner platform of an SRV structure system. The SRV includes a body extending radially from the trunnion to an unshrouded distal vane tip of the vane. The SRV includes a heat exchanger that includes a gas inlet configured to couple a bleed air valve of a compressor section of a gas generator to a gas channel. The heat exchanger includes the gas channel adapted to: enable gas to flow into the gas inlet as cool gas and out through a gas outlet as warm gas; and extend into the body of the vane to enable heat transfer from a heat source within the vane to the cool gas. The heat exchanger includes the gas outlet configured to output the warm gas from the channel to an ambient environment outside of the vane.

Gas turbine engines with inlet guide vanes are described herein. The inlet guide vanes have throat solidity (TS), variable throat solidity (VTS), and span throat solidity (STS) values within particular ranges.

The invention belongs to the related technical field of steam turbine water corrosion prevention, and discloses a steam turbine movable blade additional liquid film water corrosion prevention structure, a steam turbine and a design method of the steam turbine movable blade additional liquid film water corrosion prevention structure, and the steam turbine movable blade water corrosion prevention structure comprises a movable blade conveying channel and a nozzle seam which are arranged on a movable blade; the movable blade conveying channel extends from the blade root to the blade top in the movable blade, the multiple nozzle seams are formed in the windward side of the movable blade and connected with the movable blade conveying channel, and the movable blade conveying channel is used for conveying liquid water from the blade root. And liquid water is conveyed to the nozzle seam to be sprayed out under the centrifugal force effect of the movable blade, so that a liquid film is formed on the windward side surface of the movable blade, and the water corrosion prevention effect is achieved. A layer of liquid film can be actively formed on the surface of the windward side of the movable blade to buffer high-speed liquid drop impact, an active liquid film protection technology is formed, a traditional passive defense technology is improved and innovated, and the defects of passive defense are overcome.

An open rotor engine includes a core engine, a plurality of guide vanes positioned within or extending from the core engine; and a pitch change assembly operably coupled to the plurality of guide vanes. The pitch change assembly includes one or more actuators configured to change a pitch angle of respective ones of the plurality of guide vanes, and a plurality of linkage arms that are respectively movable by actuation of at least one of the one or more actuators. The plurality of linkage arms are directly or indirectly coupled to a corresponding one of the plurality of guide vanes. The plurality of linkage arms may have a length that differs from one another, and such length may orient a displacement or a range of motion of the respective linkage arm to an envelope of rotation about a guide vane axis that differs as between the plurality of guide vanes.

An apparatus comprising a plurality of power feeder cables (136) for transferring power between an electrical machine (128) located within an engine core of a gas turbine engine (102) and power controllers (132) located on a fan casing (138) of the gas turbine engine (102). A conduit housing (134) defines an interior for enclosing the plurality of power feeder cables (136). A first connector (140) connects a first point on the conduit housing (134) to the engine core of the gas turbine engine (102). A second connector (142) connects a second point on the conduit housing (134) to a trailing edge flange of the fan casing (138) of the gas turbine engine (102).

The invention relates to a stator part (20) of a turbine engine, comprising a platform (22), a blade (24, 26) extending radially relative to a central axis (A), and a fin (28) extending radially from a fin root (44) to a fin tip (46), the fin comprising a lower side (48) and an upper side (50), each point (100) of the lower side or of the upper side defining a radial axis (Ar) passing through the point, each plane (Pr) that includes the radial axis defining a section (S) of the lower side or of the upper side, an angle defined in the plane between the root profile and a tangent to the section at an intersection (104) of the section and of the root profile being less than or equal to 45 degrees, the section being located between the root profile and the tangent.

A double-flow turbomachine has an exhaust casing through which a hot primary flow passes and surrounded by a cold secondary flow, and a cone carried by the exhaust casing. The exhaust casing has a hollow radial arm passing through the primary flow to convey air into the cone in order to form a cooling flow for components located in the cone. The cone has an outer wall and an inner wall which runs alongside the outer wall and is radially spaced therefrom to delimit a thermal insulation space through which the cooling flow passes from upstream to downstream before it is discharged through an outlet opening at the end of the cone. It has an inner cover surrounding a component to channel the cooling flow so that it runs along the component to cool it before passing through the thermal insulation space to be discharged there.

An integrated fixed blade type variable nozzle turbocharger relates to the technical field of superchargers and comprises a turbine shell, an installation disc is fixedly arranged in the turbine shell, a plurality of fixed blades abutting against the plane of the end of the turbine shell are arranged on the end face of the installation disc in a surrounding mode, and a gas flowing channel is formed through the area between every two adjacent fixed blades. A plurality of variable blades which are arranged in a swinging mode and used for adjusting the size of the gas flowing channel are further arranged on the end face of the installation disc in a surrounding mode, and a driving disc for driving the variable blades to swing synchronously is rotationally arranged on the other end face of the installation disc. The problems that in the prior art, due to the fact that the diameter of a supporting sleeve or a spacer sleeve is large, when the turbocharger is low in speed and small in flow, vacuum exists on the back of the supporting sleeve or the spacer sleeve, and gas flow separation is caused are solved; the problem that an existing variable nozzle assembly relates to a large number of parts, the connecting position of a supporting sleeve is prone to being seriously deformed and even fractured under long-time high-low-temperature cold and hot impact, and the variable nozzle is clamped or stuck is solved.

The invention provides a gas collection structure, a turbine pump and a rocket engine. The gas collecting structure comprises a gas collecting ring and a nozzle disc, the gas collecting ring comprises an annular body and a gas inlet formed in the outer wall of the annular body, and a gas collecting groove communicated with the gas inlet is formed in one side of the annular body; the nozzle disc covers a notch of the gas collecting groove, a plurality of nozzles are arranged on the nozzle disc, inlet ends of the nozzles are communicated with the gas inlet, outlet ends of the nozzles are configured to be communicated with the turbine, the nozzles comprise a group of first nozzles and a group of second nozzles, and the second nozzles are far away from the gas inlet relative to the first nozzles; the cross-sectional area of the inlet end of the first nozzle is smaller than that of the inlet end of the second nozzle, and the cross-sectional area of the outlet end of the first nozzle is smaller than that of the outlet end of the second nozzle, so that the airflow attack angles of the first nozzle and the second nozzle at the outlet ends are equal or the difference value of the airflow attack angles is smaller than a design threshold value. The circumferential uniformity of airflow sprayed out of the nozzle disc can be improved, and the turbine efficiency is improved.

The invention relates to the technical field of aero-engines, and discloses a turbine guider flow adjusting structure and method based on porous combined jet flow. A first jet flow assembly and a second jet flow assembly are arranged on the blade back of a guide blade corresponding to a turbine guider blade grid channel in the upstream direction and the downstream direction in the incoming flow direction correspondingly; when the first jet flow assembly sprays air flow towards the upstream of the incoming flow direction, a backflow area is formed on the downstream of the first jet flow assembly due to the penetrating effect of the air flow on the incoming main flow, low-speed fluid in the backflow area can block a main flow channel of the turbine guider, the area of a pneumatic throat is reduced, and the throttling effect is achieved. When the second jet flow assembly is used for jetting airflow towards the downstream of the incoming flow direction, the airflow can be supplemented into the main flow, and the outlet flow of the high-pressure turbine guider is further increased; therefore, the adjusting range of the same cold air flow to the outlet flow of the high-pressure turbine is widened, and the pneumatic and cooling efficiency of the turbine is improved.

An air-moving device may include an aerodynamic stator. The aerodynamic stator may be positioned forward of a motor of the air-moving device and aftward of an aerodynamic rotor of the air-moving device. A control unit may be integrated in and in thermal communication with the aerodynamic stator. The aerodynamic stator may transfer heat from the control unit to thermally conductive stator vanes of the aerodynamic stator. An airflow generated by the aerodynamic rotor may facilitate heat dissipation from the thermally conductive stator vanes. The aerodynamic stator may include electrically conductive stator vanes. The electrically conductive stator vanes may provide at least one of power or control signaling to the control unit.

The invention discloses a variable geometry turbine device, which relates to the field of turbocharger technologies. The variable geometry turbine device includes: a turbine casing provided with an inlet duct, and defining an exhaust port and a bypass chamber therein; a gear rotating sleeve disposed inside the bypass chamber; a rack actuator; combined nozzle guide vanes including: a fixed nozzle fixedly disposed on a first nozzle wall surface of the turbine casing and an elastic nozzle mechanism; and a turbine. A first annular chamber and a second annular chamber are defined between the exhaust port and the bypass chamber. A first radial channel and a second radial channel are defined under the bypass chamber, and are connected to the first annular chamber and the second annular chamber respectively. A side of the gear rotating sleeve defines a first notch and a second notch configured to be selectively connected to the first radial channel and the second radial channel respectively; and opening and closing of the first radial channel and the second radial channel have a time difference. The device is a highly reliable and efficient turbine device across the entire operating range.

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.

The invention discloses an yttrium-doped gradient structure environmental barrier coating and a preparation method and application thereof, the yttrium-doped gradient structure environmental barrier coating comprises a substrate, a Si-HfO2-Y bonding layer, a Yb < 2-x > Y < x > Si2O7 intermediate layer, a y > Yb < 1.6 > Y < 0.4 > DS-(1-y) Yb < 1.6 > Y < 0.4 > MS transition layer and a Yb < 2-z > Y < z > MS environmental barrier layer which are sequentially arranged, x is 0 to 1.2, y is 0.55 to 0.85, and z is 0.2 to 1.0. In the multi-layer structure of the yttrium-doped gradient structure environmental barrier coating, a continuous yttrium-doped gradient structure is formed, so that the thermal expansion coefficient of each layer in the continuous yttrium-doped gradient structure environmental barrier coating is increased in a gradient manner, the layers have good chemical compatibility, and the continuous yttrium-doped gradient structure environmental barrier coating can stably serve for a long time in the temperature range of room temperature to 1450 DEG C.

The invention relates to a turbomachine compressor module (47) comprising an annular row of struts (46) provided with pivotable flaps (60). The struts (46) can define an inter-strut space (56) between two circumferentially adjacent struts (46, 48), in which inter-strut space (56) stator vanes (39) with variable orientation are at least partially arranged. The invention also relates to a turbomachine provided with such a module (47) and a row of rotor blades (40) directly downstream of the module.

This stator blade comprises a blade body having a cross-sectional shape that forms a blade profile. The blade body includes: a trailing end space and an adjacent space that are located between a leading edge and a trailing edge; and a plurality of trailing end cooling passages that penetrate from the trailing end space to the trailing edge. The adjacent space is located further on the leading edge side than the trailing end space. The edge of the trailing end space on the leading edge side is open and communicates with the trailing end space. A plurality of divider ribs, which are aligned in the blade height direction and divide the trailing end space in the blade height direction, and a plurality of pins are arranged in the trailing end space. The plurality of pins all both joined to a positive pressure-side inner wall surface and a negative pressure-side inner wall surface that define the trailing end space. Among the plurality of divider ribs, at least one divider rib forms an extending divider rib which extends to the interior of the adjacent space and which is joined to an inner wall surface demarcating the adjacent space.