3113results about "Stators" patented technology

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, 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 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.

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 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 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 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.

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 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.

A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.

A stator part of a turbomachine includes a platform, a blade, and a fin. The blade and the fin extend from the platform. The platform, an extrados of the blade and the fin define therebetween a gas flow channel. The channel has a section in a plane normal to an axis of the turbomachine, the section having a surface area, the surface area continuously decreasing from upstream to downstream with reference to a general gas flow direction through the turbomachine.

A turbine engine with an engine core defining an engine centerline and comprising a rotor assembly and a stator assembly. The turbine engine including a primary stage of airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor assembly. A subsidiary stage of airfoils located upstream of the primary stage of airfoils and circumferentially arranged about the engine centerline.

The invention relates to a stator blade and an open rotor engine comprising the stator blade. The stator blade is used for the rear row of the open rotor engine. The stator blade extends from the blade root part to the blade tip part, so that the noise of the open type rotor fan can be effectively reduced. A tip portion of the stator blade has a tip surface, a leading edge of the stator blade at the tip surface extending a distance from a root portion of the blade is farther than a trailing edge of the stator blade at the tip surface extending a distance from the root portion of the blade such that the tip surface is an inclined surface with a higher leading edge and a lower trailing edge when viewed from a side view.

A containment ring assembly for a turbine casing assembly, including: a first containment ring; a second containment ring, the first containment ring being radially inward from the second containment ring, wherein a radial space is located between the first containment ring and the second containment ring, the radial space allowing deformation of the first containment ring prior to the first containment ring contacting the second containment ring.

A turbine engine for an aircraft. The turbine engine includes a core turbine engine and a steam system. The steam system extracts water from combustion gases, vaporizes the water to generate steam, and injects the steam into a core air flow path of the core engine to add mass flow to core air. A core hot gas path component may be fluidly connected to a hot gas path that routes combustion gasses from the combustor. The core turbine engine may also include a turbine fluidly connected to the hot gas path to receive the combustion gases. The turbine may include a turbine airfoil. Each of the core hot gas path component and the turbine airfoil includes a combustion-gas-facing surface facing the hot gas path. A hydrophobic coating is formed on the combustion-gas-facing surface, reducing wetting of water vapor within the combustion gases on the core hot gas path component.

An assembly for an aircraft propulsion system (20) includes a propulsor rotor (38) and a guide vane structure (66). The propulsor rotor (38) is configured to rotate about an axis (22). The guide vane structure (66) includes a plurality of guide vanes (100) arranged circumferentially about the axis (22). The guide vane structure (66) is axially next to and downstream of the propulsor rotor (38). The guide vanes (100) include a first guide vane (100). The first guide vane (100) includes a camber line (108), a leading edge (110), a trailing edge (112), a leading edge section (116) and a trailing edge section (118). The first guide vane (100) extends longitudinally along the camber line (108) from the leading edge (110) to the trailing edge (112). The leading edge section (116) forms the leading edge (110). The trailing edge section (118) forms the trailing edge (112). The leading edge section (116) is configured to move longitudinally along the camber line (108) relative to the trailing edge section (118) to change a camber of the first guide vane (100).

The invention relates to an open type fan component of an open type rotor engine, the open type rotor engine and a noise reduction method. The open-type fan component comprises a stator, a plurality of fan blades and a plurality of fan blades, the rotor comprises rotor blades which are connected to a rotating shaft, and the rotor blades rotate along with the rotation of the rotating shaft; wherein the stator blade is located at the upstream of the rotor blade; the base is connected with the stator fixing force bearing device through the hydraulic executing mechanism. The open type fan component has a first state and a second state, wherein in the first state, the radial position where the radial outer end of the stator blade is located is the first position; and in the second state, the radial position where the radial outer end of the stator blade is located is a second position, and the second position is located on the radial inner side relative to the first position. By the adoption of the fan component, the advantage of low oil consumption of the engine is reserved, and the problem that BPF noise of the engine near an airport is too large is solved.

A guide vane assembly includes a guide vane and a sealing assembly slidable along a span of the guide vane relative to the guide vane and extending along a chordal axis and the span of the guide vane. The sealing assembly includes a plate configured to be disposed around at least a portion of an external surface of the guide vane. The sealing assembly further includes a manifold coupled to the plate and configured to receive pressurized fluid therein. The manifold is disposed around the plate. The sealing assembly further includes a plurality of flow tubes extending between the manifold and the plate. Each of the plurality of flow tubes are configured to receive the pressurized fluid from the manifold.

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 turbine engine with an engine core defining an engine centerline and comprising a rotor assembly and a stator assembly. The turbine engine including a primary stage of airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor assembly. A subsidiary stage of airfoils located upstream of the primary stage of airfoils and circumferentially arranged about the engine centerline.

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.

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.