4365results about "Stators" patented technology

Method for manufacturing a hollow stator blade with internal stiffeners. A method for manufacturing a turbomachine stator blade, the method comprising: - the production of a fibrous blank (100) of an aerodynamically profiled structure by three-dimensional weaving of yarns, said blank comprising two skins (110, 111), an internal housing (104a) located between the skins, one or more first fibrous stiffeners (140) extending into the internal housing and one or more second fibrous stiffeners (150, 160) extending into the internal housing, the first and second fibrous stiffeners delimiting between them cavities (112, 113, 114, 115, 116, 117) in the internal housing (104a), - the insertion of soluble cores (171, 172, 173, 174, 175, 176) in the cavities of the internal housing,- the injection of a resin into the fibrous preform and the transformation of the resin into a matrix so as to obtain an intermediate stator blade part made of composite material having an internal volume between two skins and comprising stiffeners each in contact with the inner face of the skins, at least some stiffeners having one or more circulation channels, - the dissolution of the cores present in the cavities. Figure for the abbreviation: Fig. 4.

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 includes a body and an additively manufactured (AM) metal coupon that includes at least one first porous region having a first porosity and may include any number of second porous regions having a second porosity different than the first porosity. A braze material couples the metal coupon in a coupon opening in the body. The braze material infiltrates into at least one of the first porous region based at least on a characteristic of the first porosity or any second porous region based at least on a characteristic of the second porosity. The different porosities can be customized to provide the first porous region and the second porous region with the braze material therein with different physical characteristic(s).

This mixed flow turbine is provided with a turbine wheel comprising a hub and blades provided at intervals in a circumferential direction of the hub on an outer peripheral surface of the hub. The tip ends of the leading edges of the blades are positioned on an outer side in the radial direction of the turbine wheel than the hub ends of the leading edges. When a distance between the downstream end of the outer peripheral surface of the hub and a rotational axis line of the turbine wheel is Dh and the average of distances between respective positions from the hub ends to the tip ends at the leading edges and the rotational axis line is Eave, Dh / Eave>0.4 is satisfied.

The invention relates to a metal sample block with a porous area for a component and the component with the metal sample block. A component (202) includes a body (206) and an additive manufactured (AM) metal sample block (200) that includes at least one first porous region (300A) having a first porosity and can include any number of second porous regions (300B) having a second porosity different from the first porosity. A braze material (310) couples the metallic sample block (200) in a sample block opening (204) in the body (206). The braze material (310) is infiltrated into at least one of the first porous region (300A) and any second porous region (300B) based at least on a characteristic of the first porosity and at least on a characteristic of the second porosity. Different porosities can be tailored to provide different physical characteristics for the first porous region (300A) and the second porous region (300B) having the braze material (310) therein.

An airfoil is provided that includes a first end, a second end, a leading edge, a trailing edge, a pressure side and a suction side. The leading edge and the trailing edge are joined by the pressure side and the suction side to provide an exterior airfoil surface extending in a spanwise direction from the first end of the airfoil to the second end of the airfoil. The exterior airfoil surface are formed in conformance with a plurality of cross-section profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading edge and the trailing edge at the span location.

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.

An aircraft engine (10) has: a turbine (15) including a turbine rotor (15C) rotatable about a central axis (A); a scroll case (20) having an inlet (22) fluidly connected to a source of combustion gases and an outlet (23) fluidly connected to the turbine (15), and a conduit (21) extending around the central axis (A) from the inlet (22) to the outlet (23), the conduit (21) spiraling towards the central axis (A); a bearing housing (30) extending around the central axis (A); an exhaust case (15B) disposed downstream of the turbine (15); and a turbine support case (40) secured to the bearing housing (30; 130) and to the exhaust case (15B), the turbine support case (40) having spokes (44) distributed around the central axis (A) and extending along a direction having an axial component relative to the central axis (A), the spokes (44) extending through the conduit (21) of the scroll case (20) and radially supported by the bearing housing (30).

The invention relates to an aircraft turbomachine subassembly comprising an annular duct (36) extending concentrically about a longitudinal axis and in which an air stream flows from upstream to downstream, the annular duct comprising a duct portion (28) defining a swan neck having an upstream inlet section and a downstream outlet section. A plurality of radial arms (B1, B2) extend circumferentially in the duct portion (28). One or several intermediate walls (P1, P2) extend between two circumferentially adjacent arms (B1, B2) which form a pair of circumferentially adjacent arms among the plurality of circumferentially adjacent arms so as to locally separate, between the two arms of the pair of arms, the annular duct portion (28) into a radially inner portion (Ci) and a radially outer portion (Ce) further away from the longitudinal axis than the radially inner portion.

An aircraft engine, has: a turbine; a scroll case having an inlet connected to a source of combustion gases and an outlet connected to the turbine, and a conduit extending from the inlet to the outlet; a bearing housing including a support flange; an exhaust case downstream of the turbine; and a turbine support case secured to the bearing housing and to the exhaust case, the turbine support case having spokes extending along a direction having an axial component, the spokes extending through the scroll case and radially supported by the bearing housing, a spoke having a distal end secured to the support flange via: one or more fasteners, and a retaining member at the distal end, the retaining member defining an abutment face facing an axial direction and circumferentially overlapping the support flange, a portion of the support flange located axially between the distal end and the abutment face.

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 stator device provided by the invention is used for the open type rotor engine, and comprises a stator blade base which is arranged on the open type rotor engine; the plurality of stator blades are mounted on the stator blade base and are arranged around the engine rotating shaft in the circumferential direction of the engine rotating shaft; the telescopic rod device is provided with a telescopic rod connected with the stator blade base, and the telescopic rod stretches out and draws back to enable the stator blade to be close to the engine rotating shaft or away from the engine rotating shaft in the radial direction of the engine rotating shaft; and the control device is connected with the telescopic rod device through a data line, and sends an instruction to the telescopic rod device to control the telescopic rod to stretch out and draw back, so that the stator blades are close to the engine rotating shaft or far away from the engine rotating shaft in the radial direction of the engine rotating shaft.

An aircraft engine, has: a turbine including a turbine rotor rotatable about a central axis; a scroll case having an inlet fluidly 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, the conduit spiraling towards the central axis; a bearing housing extending around the central axis; an exhaust case disposed downstream of the turbine; and a turbine support case secured to the bearing housing and to the exhaust case, the turbine support case having spokes distributed around the central axis and extending along a direction having an axial component relative to the central axis, the spokes extending through the conduit of the scroll case and radially supported by the bearing housing.

A gas turbine engine includes, among other things, a propulsor section including a rotor and blades, 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 gear train.

Aircraft turbine engine including two coaxial annular walls defining between them a main annular duct for the flow of a primary airflow; —a rotor blading extending radially through the first duct; —an annular separator arranged downstream of the rotor blading and configured to divide the primary airflow into two to form the secondary air flows; —first variable guide vanes which are distributed around the shaft and each comprise a leading edge which is located upstream of the separator and trailing edges which are located in the secondary airflows; —and fixed guide vanes which are distributed around the shaft in the external airflow and downstream of the first variable vanes.

A blade for a fan of a gas turbine engine includes an airfoil section that spans along a radial direction between first and second radial ends, and a dovetail at the second end of the airfoil section. The airfoil section and the dovetail are comprised of a fiber-reinforced composite that includes inter-section fiber plies that each extend from the airfoil section into the dovetail and wedges that are interleaved with the inter-section fiber plies in the dovetail. The dovetail is asymmetric about the radial direction and has first and second circumferential sides that are sloped to face radially outwardly and form non-zero angles with the radial direction. One of the non-zero angles is 10° or less and a sum of the non-zero angles is from 35° to 45°.

An air / oil heat exchanger arranged in a vein. The exchanger includes internal parietal corridors and / or external parietal corridors and central corridors, the radial height of the internal and / or external parietal corridors increasing from the inlet of the exchanger to its outlet.

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

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 guide vane assembly for a nacelle of a gas turbine engine includes a forward vane and an aft vane. The aft vane is located aft of the forward vane and forward of a plurality of fan blades. The forward vane defines a fixed pitch angle and the aft vane is movable between a first pitch angle and a second pitch angle.

A turbofan engine defining an axial direction and a longitudinal centerline along the axial direction is provided. The turbofan engine includes: a fan section having a fan, the fan comprising a plurality of fan blades; a turbomachine drivingly coupled to the fan, the turbomachine comprising a compressor section with a low pressure compressor, a turbine section with a low pressure turbine, a reduction gearbox, and an outer casing, the low pressure turbine drivingly coupled to the low pressure compressor across the reduction gearbox; an outer nacelle surrounding the fan and at least a portion of the turbomachine; an outlet guide vane extending between the turbomachine and the outer nacelle at a location downstream of the plurality of fan blades, the outlet guide vane defining a base and a tip and being forward swept from the base to the tip.

A hollow vane assembly including an open body including an interior; at least one cover support structure formed in said open body proximate the interior; a cover brazed to the open body to form at least one flow passage; and the open body and the cover configured as two piece substantially symmetrical halves.

A CMC airfoil includes core fiber plies that define a radial tube that circumscribes an internal cavity, and skin fiber plies that define an exterior of the airfoil section. There is a filler element in the trailing edge region aft of the internal cavity and sandwiched between the skin fiber plies on the pressure side and the skin fiber plies on the suction side. At least one cooling passage includes at least one inlet orifice section that opens to the internal cavity and extends through the core fiber plies, a radially-elongated multi-dimensional plenum connected with the at least one inlet orifice section, a plurality of outlet orifice sections extending through the trailing edge, and a plurality of intermediate passage sections bound by at least one of the skin plies and the filler element and connecting the radially-elongated multi-dimensional plenum with the plurality of outlet orifice sections.

A stator vane includes a platform that defines an opening. The stator vane further includes an airfoil that has a leading edge, a trailing edge, a suction side wall, and a pressure side wall. The airfoil extends radially between a base and a tip. At least one of the base or the tip includes a protrusion. The protrusion extends into the opening of the platform such that the platform surrounds the protrusion of the airfoil. The stator vane further includes a braze joint disposed between and fixedly coupling the platform and the protrusion of the airfoil. The stator vane further includes a cooling circuit defined in at least one of the protrusion or the platform to cool the braze joint.

The invention provides a rear casing guide vane and a design method thereof, and relates to the technical field of power machinery. The rear casing guide vane design method comprises the steps that S01, according to guide vane design information, a vane profile is designed, and a designed vane profile is obtained; the blade profile design comprises an inlet airflow angle, an outlet airflow angle, the maximum thickness and the axial chord length, and the position of the maximum thickness is within the range of 0.2-0.3 of the axial chord length; s02, generating a geometric model of the designed blade profile, and performing numerical calculation on the geometric model; s03, analyzing a numerical calculation result, and if an analysis result does not meet the design requirement, carrying out iterative design on the blade profile until the design requirement is met; and S04, variable working condition analysis is conducted on the blade profiles meeting the design requirements. The rear casing guide vane design method can adapt to an inlet large pre-rotation angle rear casing, the total pressure loss of the rear casing is reduced, and the economical efficiency is improved.

A tail cone ventilation system including a tail cone case defining a tail cone interior and a tail cone exterior, the tail cone having a forward portion and an aft portion separated axially along a tail cone axis; a distribution manifold located within the tail cone interior proximate the forward portion, wherein the distribution manifold comprises nozzles arranged radially around the axis, the nozzles configured to direct a cooling air over at least one electronic component within the tail cone interior and along an inner surface of the tail cone case; an air inlet fluidly coupled with the distribution manifold through ducting, the air inlet located externally from the tail cone interior; and a tail cone discharge located proximate the tail cone aft portion, the tail cone discharge being fluidly coupled with the distribution manifold.

The invention relates to a propulsion unit which comprises: - an outer casing (102); - a hub; - a pusher propeller; and - a fixed flow straightener comprising at least one stator blade (114) with variable pitch about a pitch axis (Y'), each variable-pitch stator blade (114) having a pitch angle defined as the angle between any plane perpendicular to the main axis (X) and a chord line (706) connecting a leading edge of the stator blade (114) to a trailing edge of the stator blade (114), this chord line (706) being taken at 75% of an outer radius of the stator blade (114) with respect to the main axis (X). The leading edge of at least one stator blade (114) has a dihedral angle (D'BA') at 95% of the outer radius with an absolute value greater than 1°, preferably greater than 3°, more preferably greater than 10°, for at least one value of the pitch angle within the range [70°; 90°].

A rotor system for a turbine engine. The rotor system includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades that rotate. The stator assembly includes a plurality of stator vanes arranged circumferentially about the stator assembly and includes at least one pair of non-uniform gaps between adjacent stator vanes. The plurality of stator vanes includes a first group of stator vanes having a first non-uniform gap between adjacent stator vanes, a second group of stator vanes having a second non-uniform gap between adjacent stator vanes, and a third group of stator vanes having a uniform spacing between adjacent stator vanes. The first non-uniform gap is positioned 180° from the second non-uniform gap. The plurality of rotor blades directs air through the plurality of stator vanes.

The invention discloses a turbine for testing, an engine testing device and a method for determining a pneumatic instability boundary of a compression system, which can determine the pneumatic instability boundary of the compression system in a complete machine environment, do not cause extra burden to the turbine and compression system components, and do not need to modify a tester and surrounding equipment on a large scale. The turbine for testing is used for replacing an original turbine in a complete machine surge test of an engine, and the engine testing device is used for the complete machine surge test of the engine. Stator blades of the test turbine are arranged to be rotatable blades, so that the circulation sectional area of the test turbine is reduced by rotating the stator blades, and the compression system can enter a surge state.