980results 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.

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 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.

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.

A turbine engine has a compression section, a combustion section, and a turbine section in a serial flow arrangement and defining a working airflow path. The combustion section has a circumferential housing, a combustor, and a fuel nozzle. The circumferential housing defines an interior. The turbine engine includes a de-swirler assembly. A de-swirler assembly couples a working airflow path of the compression section to the interior. The turbine engine includes a fuel supply system.

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.

The invention relates to a stator assembly (10) for a turbine engine compressor, comprising two successive rows of upstream stator vanes (121) and downstream stator vanes (124) defining a tandem configuration, the vanes (121, 124) extending radially between inner and outer support members, wherein, within the tandem configuration, the chord (26) of the upstream vane (121), extending between the leading edge and the trailing edge of the vane, increases with the height of the upstream vane: at 0% of the height of the upstream vane (121), the chord (26) of the upstream vane is at least 1.9 times smaller than the chord of the upstream vane at 100% of the height of the upstream vane and, at 70% of the height of the upstream vane (121), the chord (26) of the upstream vane is at least 1.2 times smaller than the chord of the upstream vane at 100% of the height of the upstream vane.

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 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 method for constructing a nozzle ring 40 for a turbocharger turbine nozzle 28 includes the steps of: providing a nozzle ring 42 in the form of an annular flat disk, the nozzle ring having a first face and an opposite second face; forming a plurality of circumferentially spaced circular bores 43 extending through the nozzle ring 42 from the first face to the second face; providing a plurality of vanes 44, each vane having a circular vane shaft 46 extending from one end of the vane; inserting the vane shafts 46 respectively into the bores 43 in the nozzle ring from said first face thereof and orienting each vane 44 to achieve a desired setting angle for the vane; and rigidly affixing the vane shafts 46 to the nozzle ring 42 to fix the vanes at the desired setting angles.

A frame (100) for directing air into an aircraft engine. The frame (100) comprises at least one section (102) for directing air through the frame, and the frame (100) also comprises composite material (120). The composite material (120) is secured to the at least one section (102), and the composite material (120) at least partially defines a mount portion (200) from the frame for attaching the frame (200) to a component part from an aircraft. This has been found to reduce stress concentrations around parts of the frame (100) which are located proximal to the mount portion (200), by allowing the forces creating such stress concentrations to be more effectively distributed about wider parts of the frame (100) where the composite material (120) is additionally present.

Electric motor comprising a squirrel-cage rotor in the form of a cage, which electrically conductively carries a plurality of rotor end rings formed as cast parts made of a casting material, wherein at least one reinforcing element made of a different material is cast into each of the cast parts, and wherein the casting material substantially forms the outermost free surface of the rotor end ring.

The rotary device T1 has an impeller 3, a first housing 6, a second housing 5, an annular flow path 62 positioned radially outside the impeller 3, a plurality of vanes V arranged in the annular flow path 62, an annular plate 52 positioned between the plurality of vanes V and the second housing 5 and supporting the plurality of vanes V, the plurality of vanes V being fixed to the annular plate 52, the annular plate 52 having a positioning surface 54a facing radially inwards and contacting the second housing 5 from a radially outer side, and an elastic body 55 arranged between the annular plate 52 and the second housing 5 and pressing the annular plate 52 and the plurality of vanes V in the axial direction.wherein the positioning surface 54a is located radially inside the elastic body 55 or is located in the same position as a radially inner end of the elastic body 55 in a radial direction.

The invention relates to a turbocharging device. The turbocharging device comprises a shell; the turbocharging unit is contained in the shell, and the turbocharging unit comprises a nozzle ring assembly and a turbine assembly; the nozzle ring assembly comprises a mounting base body, a blade assembly, a shifting ring and a limiting assembly, the shifting ring is coaxially and movably connected with the mounting base body, the blade assembly is movably connected to the mounting base body, and the blade assembly is further connected with the shifting ring so as to be driven by the shifting ring to rotate; the limiting assembly comprises a first limiting body, a second limiting body and a third limiting body which are arranged on the installation base body, and the first limiting body, the second limiting body and the third limiting body abut against the inner ring of the shifting ring in the radial direction of the installation base body. The stability of the rotating center of the shifting ring can be guaranteed, eccentricity is avoided, and therefore the long-term working precision and durability of a variable-section adjusting system of the turbocharging equipment are guaranteed.

A radial inlet body of a gas turbine engine compressor is provided. The radial inlet body includes plenums arranged around opposite sides of an inlet portion defining a compressor inlet. The plenums include outer surfaces with increasing radii of curvature and a flow diverter. The flow diverter is disposed aside the inlet portion opposite a duct. The plenums are configured to direct flows from the duct around the inlet portion. The flow diverter and the plenums are configured to cooperatively encourage uniform distribution of the flows. The flow diverter includes curvilinear sides that tangentially register with and smoothly lead to the outer surfaces of the plenums and a cusp where the curvilinear sides meet. The radial inlet body further includes a front sidewall supporting the inlet portion and an aft sidewall opposite the front sidewall. The front and aft sidewalls are angled toward one another along the plenums.

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An example airfoil (123) includes: a convex panel (920, 1120); a concave panel (1020); an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140), the concave panel (1020) having an area of maximum curvature (MC) including a plurality of metering holes (910, 1010, 1110), the airfoil (123) characterized by the following equation, in which a MC metering hole area (1230) (MCA (870, 1180)), inner band (940, 1040, 1140) radius (IBR), and loading area (840, 1090) (LA) are related as follows: EQ1=IBRm1mLAm20.00001m2∗MCA8701180m20.000005m22, wherein 1.720x10-6 meters squared (m2) ≤ MCA ≤ 3.000x10-6 m2, 0.213 m ≤ IBR ≤ 0.221 m, 0.0000147 m2 ≤ LA ≤ 0.0000153 m2, and 0.351 ≤ Equation 1 ≤ 1.644.

An apparatus is provided for an aircraft propulsion system. This apparatus includes a variable guide vane. The variable area vane includes an airfoil, and the variable guide vane is configured to pivot about an axis. The airfoil extends spanwise along the axis to a tip. The airfoil extends longitudinally along a camber line between a leading edge and a trailing edge. The airfoil extends laterally between a first side and a second side. The airfoil includes a vane acoustic structure. This vane acoustic structure includes a first skin and a plurality of chambers within the airfoil. The first skin at least partially forms the first side. A first of the chambers is fluidly coupled with one or more perforations in the first skin.

The application provides an electric spark machining method for a deep and narrow ventilation groove of a guide vane and the guide vane, and the method comprises the following steps: first, electric spark high-speed small-hole machining is adopted, a plurality of bottom holes are machined on the guide vane to be machined by using a first electrode; then, a second electrode is made, the shaped electrode and corresponding clamping auxiliary components are installed to an electric spark forming machine; electric spark forming oil immersion machining is adopted, the lateral oil flushing of the guide vane to be machined is increased, the machining parameters of the electric spark forming machine are set, and the deep and narrow ventilation groove is machined. The application pre-machines part of the bottom holes in the position to be machined in the guide vane, and the method of electric spark forming machining of the shaped groove is adopted to machine the deep and narrow ventilation groove on the plurality of bottom holes, so that the oil flushing and chip removal conditions during the forming machining of the deep and narrow ventilation groove are improved, the electric spark forming machining efficiency and surface machining quality of the guide vane are improved. Moreover, the application can be applied to the machining of various deep holes, grooves and flow channels, and has good applicability.

A ring segment cooling structure with enhanced cooling performance, a ring segment (1600), and a gas turbine (1300) including the same. The ring segment cooling structure includes a cooling channel (1630) and a rib turbulator (1637). The cooling channel is formed inside a shield wall (1611) of a ring segment mounted to a turbine casing that accommodates blades (1400) and is formed in a wavy shape so as to allow cooling air to flow through the cooling channel. The rib turbulator is installed in the wavy-shaped cooling channel.

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An airfoil (123) of an engine nozzle segment (140, 800) includes: a convex panel (920, 1120); a concave panel (1020); an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140). The convex panel (920, 1120) includes a plurality of metering holes (910, 1010, 1110), the airfoil (123) characterized by the following equation, in which a convex panel (920, 1120) metering hole area (1230) (CVXA (860, 980)), inner band (940, 1040, 1140) radius (IBR), and loading area (840, 1090) (LA) are related as follows: EQ1=CVXAm20.000005m2+0.000005IBRmπmLAm20.00001m2, wherein 0.0000001 meters squared (m2) ≤ CVXA ≤ 0.000004 m2, 0.213 m ≤ IBR ≤ 0.221 m, 0.0000147 m2 ≤ LA ≤ 0.0000153 m2, and 4.835 ≤ Equation 1 ≤ 281.150.

A casing structure for a gas turbine engine comprising a plurality of struts defining a plurality of annular circumferentially extending airflow passages and where the strut side surfaces comprise recesses.

An assembly for a turbomachine of longitudinal axis includes a nozzle carried by an annular part which is fixed to an outer casing. The nozzle has an outer annular platform from which extends, radially outwards, a radial annular wall which carries an annular attachment tab extending longitudinally upstream from the radial annular wall. The assembly has a cooling circuit with an inner annular chamber delimited radially by the outer annular platform and the annular part and delimited downstream by the radial annular wall. An intermediate annular chamber is formed radially externally to the annular attachment tab and in fluid communication with the outer casing. Means of fluid communication between the inner annular chamber and the intermediate annular chamber is formed in the annular attachment tab of the nozzle.

Method for manufacturing a turbine blade, the method comprising: a brazing treatment for joining a brazing material to a substrate of a turbine blade by heating the substrate with the brazing material placed on it and melting the brazing material after the substrate has been subjected to a solution annealing treatment at a temperature above the solidus temperature of the brazing material; a stabilization treatment for heating the substrate material undergoing brazing; and an aging treatment to heat the substrate material undergoing stabilization treatment at a temperature below the solidus temperature of the brazing material, where the brazing treatment and the stabilization treatment are carried out by means of a sequential heat treatment.

A cooling circuit of an airfoil for a gas turbine engine includes a serpentine channel having a plurality of fluidly connected radially extending flow passages and a tip flag channel disposed adjacent to the serpentine channel. The tip flag channel includes a first tip flag passage extending radially and a second tip flag passage extending axially from a terminal end of the first tip flag passage, the second tip flag passage disposed radially outward of the serpentine channel. The second tip flag passage is in direct fluid communication with each of the first tip flag passage and a final passage of the serpentine channel.

The utility model relates to the technical field of turbochargers, in particular to a waste gas kinetic energy conversion type turbocharger nozzle ring. According to the technical scheme, an upper base is installed on a base through a locking bolt, a guide plate is installed between the upper base and the base, the guide plate is fixed to a rotating shaft and drives the rotating shaft to deflect, a deflection block is installed at the end, located on the outer side of the upper base, of the rotating shaft, a lower gear ring is installed on the outer side of the upper base, and the lower gear ring is fixed to the upper gear ring. A driving gear is arranged on the deflection block, and the lower gear ring is meshed with the driving gear when rotating, so that the deflection block drives the guide plate to rotate. The electric push rod pushes the rack to drive the upper gear ring meshed with the rack to rotate, and then the upper gear ring drives the lower gear ring connected with the upper gear ring to rotate, so that the lower gear ring is meshed with the driving gear, the angle adjustment of the guide plate is accurately realized, the waste gas flow guide effect and the kinetic energy conversion efficiency are improved, and good stability and practicability are achieved.