Abstract: A turbine vane for a gas turbine engine may include a composite airfoil structure. The composite airfoil structure may have an opening. The turbine vane may include a spar. The spar may have a body, which may be disposed within the opening. A standoff structure may be disposed within the opening. In some non-limiting embodiments, a cooling air gap may be defined between the body and an internal surface of the composite airfoil structure.

Abstract: A fireshield fastener hood and method of bonding a cover to a substrate therefor.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with composite grooves and vertically projecting rows of ridges in planform patterns, establishing upper and lower wear zones. The lower wear zone reduces, redirects and/or blocks blade tip downstream airflow leakage, while the upper wear zone is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone. An elongated first ridge in the lower wear zone terminates in a continuous surface plateau. A plurality of second ridges or nibs, separated by grooves, project from the plateau, forming the upper wear zone. Each of the second ridges has a planform cross section smaller than the plateau planform cross section and a height smaller than the first ridge height. Some embodiments of the second ridges have spacing, planform cross sections, heights and separating groove dimensions selected for shearing when contacted by turbine blade tips.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with composite grooves and vertically projecting rows of stepped first ridges in planform patterns, to reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. Each stepped first ridge has a first portion proximal the substrate surface with a pair of first opposed lateral walls terminating in a plateau, and a second portion terminating in a ridge tip. These ridge or rib embodiments have first lower and second upper wear zones. The lower zone, which at and below first portion height, optimizes engine airflow characteristics, while the upper zone, between the plateau and the second portion ridge is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone.

Abstract: One aspect of the present disclosure includes a turbine vane assembly comprising a vane made from ceramic matrix composite material having an outer wall extending between a leading edge and a trailing edge and between a first end and an opposing second end; an endwall made at least partially from a ceramic matrix composite material configured to engage the first end of the vane; and a retaining region including corresponding bi-cast grooves formed adjacent the first end of the vane and a receiving aperture formed in the endwall; wherein a bond is formed in the retaining region to join the vane and endwall together.

Abstract: The present application relates to a stator blade sector configured to be fixed to a housing of an axial turbomachine, the sector having a plurality of blades with platforms juxtaposed, so as to describe an arc of a circle. At least one of the platforms comprises on its outer face a fixing screw and at least one other platform has no fixing screws, the platforms being fixed together at their adjacent edges. The application also relates to a stator or portion of stator having a housing forming a generally circular wall and blade sectors arranged along the wall. The housing includes several parts connected to each other by longitudinal flanges. Platforms with no fixing screws are located opposite the flanges.

Abstract: A CMC turbine stator blade 1 includes a blade 2 of a ceramic matrix composite material formed or a metallic material, and a band 5 formed of a ceramic matrix composite material and supporting the blade 2. The blade 2 includes a first fitting portion 3 to which the band 5 is fitted outside. The band 5 includes a second fitting portion 7 to which the first fitting portion 3 is fitted inside. A flexible wire 8 is disposed between the fitting portions 3, 7 to fix the blade 2 and the band 5. A first groove 3a in the outer peripheral surface of the first fitting portion 3 extends in the front-to-rear direction of the blade 2. A second groove 7a in the inner peripheral surface of the second fitting portion 7a extends along the first groove 3a. The wire 8 is disposed between the grooves 3a,7a.

Abstract: A composite material and method for manufacturing the material, the material including: a plurality of plies layered one on top of the other; and one or more through-thickness fibers which join one or more of the plurality of plies to one another; wherein the one or more through-thickness fibers form a boundary which delineates one or more discrete regions of the material.

Abstract: The invention relates to a method for producing a component from a TiAl alloy, wherein the component is shaped by forging, in particular isothermal forging, and is subsequently subjected to at least one heat treatment, wherein in the first heat treatment the temperature is between 1100 and 1200° C. and is maintained for 6 to 10 hours and then the component is cooled.

Abstract: A turbine engine vane is made of composite material comprises a hollow airfoil with an internal passage extending along its entire length, an inner platform, and an outer platform. The airfoil is extended on the outside of the outer platform by an upstream attachment portion and a downstream attachment portion for mounting the vane in a casing. The attachment portions are formed by an attachment extending from the outer longitudinal ends of the pressure side wall and the suction side wall of the airfoil with continuity of the fiber reinforcing texture of the composite material between the walls and the attachment tabs.

Abstract: A damper for a gas turbine engine is provided. The damper comprises a first portion, a second portion, and a surface. The second portion angularly extends from the first portion to define an interior corner therebetween. The surface bounds the first portion, the second portion, and the interior corner. The damper further comprises a grain composition containing columnar grains and equiaxed grains. The equiaxed grains are disposed adjacent to the surface such that a volume of the columnar grains is less than a volume of the equiaxed grains in the grain composition.

Abstract: A ram air fan inlet housing for containing a ram air fan rotor. The inlet housing includes a flanged surface and an interior surface. The flanged surface is perpendicular to an axis of the inlet housing and defines a flange plane at an axial end of the inlet housing. The interior surface is symmetric about the axis of the inlet housing and includes a flange section, a transition section, an outlet section, a rotor section, and an inlet section.

Abstract: A turbomachine casing assembly including a first casing element locatable radially outward of one or more rotating aerofoil elements of a turbomachine, a second casing element located radially distal to the first casing element is provided. A void is provided between a radially proximal face of the first casing element and a radially proximal face of the second casing element, at the first end of the first casing element. The first end of the first casing element is located against a corresponding first end of the second casing element by an energy absorbing element, with the first casing element; including a cantilever. The cantilever is arranged in a region extending along the first casing element from the first end; and having a pivot axis at a mid portion thereof.

Abstract: A method for manufacture of a compressor housing which is resistant to a titanium fire (burning titanium). A circular blank made of steel, or steel alloy, or superalloy which is incombustible in the presence of a titanium fire is hydroformed against a circular blank made from titanium or titanium alloy.

Abstract: A material for a gas turbine component, to be specific a titanium-aluminum-based alloy material, including at least titanium and aluminum. The material has a) in the range of room temperature, the ?/B2-Ti phase, the ?2-Ti3Al phase and the ?-TiAl phase with a proportion of the ?/B2-Ti phase of at most 5% by volume, and b) in the range of the eutectoid temperature, the ?/B2-Ti phase, the ?2-Ti3Al phase and the ?-TiAl phase, with a proportion of the ?/B2-Ti phase of at least 10% by volume.

Abstract: A component for a gas turbine, especially a blisk or a bling, whereby the component includes a rotor base (12) made of a high temperature-resistant nickel alloy and a plurality of turbine blades (14) joined to the rotor base, whereby each turbine blade includes a rotor blade (16) made of a titanium alloy and a blade root. The blade root is configured as an adapter element (18) that is made of a material that can be welded to the titanium alloy as well as to the high temperature-resistant nickel alloy and that is integrally joined to the rotor base (12) and to the rotor blade (16) fusion. A method for the production of the component is also described.

Abstract: In accordance with an exemplary embodiment, a turbine stator component includes a first endwall; a second endwall; a first stator airfoil coupled between the first and second endwalls; and a second stator airfoil adjacent to the first airfoil and coupled between the first and second endwalls. The first stator airfoil has first crystallographic primary and secondary orientations. The second stator airfoil has second crystallographic primary and secondary orientations, the first crystallographic primary and secondary orientations being different from the second crystallographic primary and secondary orientations.

Abstract: A nozzle for use in an air cycle machine has a plate. A plurality of vanes extends in a second axial direction away from the plate. The plurality of vanes extends for a height away from the plate and a width defined as the closest distance between two adjacent vanes, with a ratio of the nozzle height to the nozzle width being between 0.3563 and 0.4051. An air cycle machine and a method are also disclosed.

Abstract: A flame resistant shield includes two opposing walls between which are positioned at least: a first insulating layer, the first layer being capable of distributing heat in the plane formed by the first layer and being insulating across its thickness, a second insulating layer, one of the opposing walls which covers the first layer being produced from a refractory antioxidant material or having at least the surface intended to be exposed to flames covered with a material preventing this surface from being oxidised, the other wall being a support.

Abstract: Single-airfoil vanes each having an inner platform, an outer platform, and an airfoil are obtained by three-dimensionally weaving a fiber blank in a single piece, by shaping the fiber blank to obtain a single-piece fiber preform, and by densifying the preform with a matrix to obtain a vane of composite material forming a single piece with inner and outer platforms incorporated therein. A plurality of vanes is assembled together at an intermediate stage of densification to form a multi-airfoil composite material guide vane sector for a turbine nozzle or a compressor diffuser and the assembled-together vanes are bonded together.

Abstract: A method for armoring TiAl vanes of turbomachines is disclosed. A TiAl vane is provided onto which a mixture of at least one hard material and at least one braze material is applied so that subsequently the mixture can be brazed on the TiAl vane by an inductive heating process. A TiAl vane for a turbomachine, in particular for an aircraft engine, is also disclosed. The TiAl vane includes a TiAl main part and an armor which consists of a mixture of hard materials and braze material.

Abstract: A vane assembly for a gas turbine engine comprises at least one vane, a first platform and a second platform. The at least one vane has a first end and an opposite second end with the first end of the at least one vane being secured in a corresponding first aperture in the first platform, and the second end of the at least one vane being secured in a corresponding second aperture in the second platform. The at least one vane intersects the first platform at an acute angle, and the at least one vane intersects the second platform at an obtuse angle with the acute angle and the obtuse angle being supplementary to one another.

Abstract: A wind turbine is provided with turbine blades mounted for axial rotation about an axis. The blades are surrounded by a shroud to define an axial air passage. A conical ring is attached to the shroud and includes vanes for directing the airflow. Plates are attached to the shroud at a position radially outward from the shroud forming an air passage between the shroud and the plates. The plates have gaps between the adjacent plates so that air exiting the downstream opening of the shroud and air moving through the axial air passage between the shroud and the plates are mixed and a portion of the mixed air exits through the gaps. The wind turbine includes photovoltaic cells to increase the electricity generated by the wind turbine.

Abstract: A seal segment is provided for a shroud ring of a rotor of a gas turbine engine. The seal segment is positioned, in use, radially adjacent the rotor. The seal segment has two or more spaced passageways. In use, a respective support bar can be contained within each passageway. The support bars allow attachment of the seal segment to the engine casing. Each passageway has an elliptical or substantially elliptical cross-section with a minor diameter and a major diameter, the minor diameter being substantially aligned with the radial direction. The respective support bar has a correspondingly elliptical or substantially elliptical cross-section for a close fit in the passageway.

Abstract: A steam turbine facility is provided which is capable of appropriately sealing a gap between members made of cast material of Ni-based alloy even under the steam-temperature condition of 650° C. or higher, and which includes a first member and a second member whose base material is formed of cast material composed of at least one of Ni-based alloy, austenite steel, or high-chrome steel, form a space where the steam of 650° C. or higher flows. Between the first member and the second member, a metal gasket is provided which has a plurality of portions in line contact with the first member and the second member. On the first member and the second member, a first high-hardness layer which is harder than the base material, is provided at least in a portion where each of the first member and the second member is in line contact with the metal gasket.

Abstract: A centrifugal compressor comprises at least one stage suited to separate a liquid phase and a gas phase with the aid of at least one of a hydrophobic, super-hydrophobic, hydrophilic or super-hydrophilic surface layer, wherein the hydrophobic and/or super-hydrophobic surface layer is disposed on at least one of an inlet guide vane, impeller, return channel straight hub, or exiting hub bend; and the hydrophilic and/or super-hydrophilic surface is disposed on at least one of the impeller casing, diffuser casing, exiting casing bend, return channel straight hub, exiting hub bend, collection point, or drain.

Abstract: A method of fabricating a part made out of TA6Zr4DE titanium alloy including forging a blank in the alpha/beta domain to form a preform, hot die-stamping the preform to form a rough part in the beta domain of the titanium alloy, and heat treatment. During the die-stamping, the rough part is subjected throughout to local deformation c greater than or equal to 1.2, the die-stamping being terminated by immediate cooling at an initial cooling rate faster than 85° C./min. The method can be utilized to make a rotary part of a turbomachine.

Abstract: A composite aerofoil vane for a gas turbine engine extends in use in a generally radial direction across an axially- and circumferentially-extending annular duct. The vane comprises a plurality of first reinforcing fibres extending in a generally radial direction and a plurality of second reinforcing fibres extending in generally axial and circumferential directions. The second fibres are interlaced with the first fibres so as to form a more unified and integrated structure, and to prevent the first reinforcing fibres from moving relative to one another.

Abstract: A method is described for forming a lining on an inner surface of a pump casing. The lining can protect the pump casing from corrosive and abrasive fluids. The method disclosed improves the durability and reliability of the lining by introducing a compressive pre-stress to the lining when the casing is not loaded such that during normal operation, the lining can be in a stress state near neutral and avoid failures related to excessive tensile strains. The method includes securing at least a portion of the casing in a stationary position and applying an external load to deform the casing. The method further includes depositing the lining onto the inner surface of the casing, curing the lining, and removing the external load to pre-stress the lining.

Abstract: A novel ceramic matrix composite is disclosed for forming components that are operable in high temperature environments such those in gas turbine engines and the like. The ceramic matrix composite can include at least one layer of non-crimped fibers positioned substantially parallel to one another. A relatively small diameter elastic fiber can be constructed to stitch the non-crimped fibers together and a ceramic matrix may be deposited around the at least one layer of non-crimped fibers.

Abstract: The disclosure relates to a steam turbine outer casing that can be at least partially cast out of a first material and has a plurality of support regions adapted to support an inner casing. The outer casing can be at least one support region that has a metal insert, made of a second material, with a flaring portion for retaining the metal insert in the casting of the outer casing. The second material has a greater hot strength than the first material. The support region can include a guide for limiting the movement of the inner casing in the outer casing.

Abstract: A vane assembly for a gas turbine engine is disclosed and includes a plurality of individual vane segments forming an annular ring. Each of the plurality of vane segments is formed of a ceramic matrix composite including at least one airfoil extending between an inner platform and an outer platform. An outer wrap is disposed about an outer periphery of the annular ring and over an interface between outer platforms of adjacent ones of the plurality of vane segments.

Abstract: In one exemplary embodiment of the present invention, a compressor housing for a forced induction system of an internal combustion engine is provided. The compressor housing includes a compressor inlet passage in fluid communication with a compressor volute configured to house a compressor wheel, the compressor inlet passage comprising a wall that is shared with the compressor volute. The compressor also includes a compressor outlet in fluid communication with the compressor volute, the compressor outlet being configured to direct a compressed gas to an intake manifold of the internal combustion engine.

Abstract: A turbomachine component of a stationary turbomachine includes a substrate made of high alloyed 10% to 18% chromium steels or titanium alloys or nickel base alloys or cobalt base alloys with a substrate surface and an erosion and corrosion resistant coating system. The coating system includes a first layer, which is deposited on the substrate surface of the turbomachine component and acts as a corrosion resistant layer and a second layer. The second layer is deposited on the first layer and acts as an erosion resistant layer, wherein the first layer is a Zr single layer and the second layer is a W/WC multilayer.

Abstract: A system may include a structure having a leading edge and a fiber-reinforced composite cover. The fiber-reinforced composite cover may include a protective portion and a retaining portion, wherein the retaining portion extends about the structure to self-retain the protective portion along the leading edge, and the protective portion is different from the retaining portion.

Abstract: A layered composite component is disclosed, the component including: a plurality of stacked layers defining a component thickness between opposed component surfaces and at least one reinforcing element extending from each of the opposed component surfaces the reinforcing elements extending at least partially through the component thickness. Another layered composite component is also disclosed, the component including: a plurality of stacked layers defining a plane and having a primary in-plane axis, wherein the at least one block array of reinforcing elements extends into the component along a reinforcing direction that is angled with respect to the plane of the stacked layers, and wherein the at least one block array is angled in the plane of the stacked layers with respect to the primary axis.

Abstract: A turbine nozzle or a compressor diffuser includes sectors made of CMC material, each having an inner platform, an outer platform, and airfoils. An abradable material support ring is made up of sectors, each presenting upstream and downstream attachment tabs. Each inner platform presents on the inside an upstream hook and a downstream hook, and the abradable material support ring is supported by the nozzle or the diffuser by mutually engaging without fastening together firstly the end portions of the upstream attachment tabs and the upstream hooks and secondly the end portions of the downstream attachment tabs and the downstream hooks.

Abstract: A steam turbine facility suppresses the possibility of vibration from occurring and prevents a drastic increase in facility cost, thereby realizing an increase in size of the facility, even if steam conditions of 650° C. or higher are adopted. In the steam turbine facility including a high-pressure turbine, an intermediate-pressure turbine, and a low-pressure turbine, the intermediate-pressure turbine is separated into a first intermediate-pressure turbine on a high-temperature and high-pressure side and a second intermediate-pressure turbine on a low-temperature and low-temperature side. At least any one of the rotors and casings of the steam-introduction-side turbines into which steam with a temperature of 650° C. or higher is introduced is formed from Ni-based alloy, and at least any one of the overall rotors and the overall casings of the turbines are constructed by joining together a plurality of rotor members or casing members by welding.

Abstract: A turbine engine fan casing of composite material including fiber reinforcement densified by a matrix is made as a single part with ribs that project from an inside face and/or from the outside face of the casing, each rib including a fiber reinforcement portion connected to the fiber reinforcement of a remainder of the casing by weaving or by stitching. On the inside, the ribs serve to mount acoustic panels to the inside face of the casing to form an assembly including a casing and acoustic panels. On the outside, the ribs enable equipment to be mounted on the casing.

Abstract: An example gas turbine engine shroud includes a first annular ceramic wall having an inner side for resisting high temperature turbine engine gasses and an outer side with a plurality of radial slots. A second annular metallic wall is positioned radially outwardly of and enclosing the first annular ceramic wall and has a plurality of tabs in communication with the slot of the first annular ceramic wall. The tabs of the second annular metallic wall and slots of the first annular ceramic wall are in communication such that the first annular ceramic wall and second annular metallic wall are affixed.

Abstract: A nickel-based alloy and a turbine component are disclosed. The alloy includes, by weight, between about 0.8% and about 1.3% hafnium, between about 5.7% and about 6.4% aluminum, between about 7.0% and about 10.0% cobalt, up to about 0.1% carbon, up to about 8.7% chromium, up to about 0.6% molybdenum, up to about 9.7% tungsten, up to about 0.9% titanium, up to about 0.02% boron, up to about 0.1% manganese, up to about 0.06% silicon, up to about 0.01% phosphorus, up to about 0.004% sulfur, up to about 0.02% zirconium, up to about 1.8% niobium, up to about 0.1% vanadium, up to about 0.1% copper, up to about 0.2% iron, up to about 0.003% magnesium, up to about 0.002% oxygen, up to about 0.002% nitrogen, and a balance nickel. The turbine component is a turbine bucket, a turbine nozzle, or any other suitable turbine component including the alloy.

Abstract: A component, for example an aerofoil vane or other gas path structure in a gas turbine machine such as a gas turbine engine, in which an insulated electrical conductor path is embedded structurally integrally, and in which electromagnetic shielding, for the embedded electrical conductor path is provided structurally integrally.

Abstract: A vane structure for a gas turbine engine includes a multiple of CMC airfoil sections integrated between a CMC outer ring and a metal alloy inner ring.

Abstract: A steam turbine system including a steam turbine is provided. The steam turbine system includes a high-pressure side steam inlet device, a low-pressure side steam device, and a control device for controlling the steam turbine. An additional steam inlet device is also included arranged between the high-pressure side steam inlet device and the low-pressure side steam device. The control device control a supply of steam via the additional steam inlet device as a function of operating parameters detected at the steam turbine system.

Abstract: A vane includes a platform with a fixture overmolded by a sheath.

Abstract: The invention relates to a method for producing, repairing and/or exchanging a housing, in particular an engine housing of an aircraft engine, comprising at least two shells, between which a structural part is formed, wherein the method comprises the following step: layer-by-layer construction of the at least two shells jointly with the structural part by means of a generative manufacturing system, wherein the structural part comprises at least one porous structure and/or honeycomb structure. The invention relates furthermore to such a housing.

Abstract: There is provided a stator that can ensure safe rotation even if a width of a gap between a screw rotor and a stator is particularly smaller than that of a conventional one and can greatly increase a pumping performance of a pump. A stator of the present invention is formed of a metal and a PFA film is provided on an inner wall surface of a top end surface of a screw rotor. After coating the inner wall surface with PFA, followed by melting and remelting processes, the PFA film is formed to have a high smoothness. The stator is placed with a gap between the inner wall surface and the top end surface. If the PFA film of the present invention is provided in a predetermined manner, the width of the gap can be made significantly smaller than that of a conventional one.

Abstract: A guide vane of a condensation turbine steam turbine is provided, wherein the guide vane includes a heating resistor. The guide vane includes fiber composite material at least in some regions. The heating resistor may be embodied as a heating wire or as a heating film. A condensation steam turbine having a guide vane as described above is also provided.

Abstract: An example composite fan case assembly includes a composite fan case body disposed on an axis. The fan case body has a radially inner surface and a radially outer surface. There is at least one mounting attachment to receive a threaded member. The at least one mounting attachment includes a base component and a spool component. The base component extends through an opening in the fan case body from the radially outer surface and toward the radially inner surface of the fan case body. The base component including a base portion, a boss having a non-cylindrical profile, and a cylindrical portion. The mounting attachment also includes a spool component disposed on the radially outer surface and attached to the base component. The spool component receives at least a portion of the base component. An outer surface of the cylindrical portion and an inner surface of the spool component are threaded and correspondingly mate to lock the spool component relative to the base component.

Abstract: A method of repairing a flange of a casing for an airplane turboprop or turbojet, in which the flange includes at least one through hole for passing a bolt for fastening equipment, and the hole opens out in a face of the flange that is worn by friction against the equipment, disclosed. The method includes mounting a bearing member on the flange. The bearing member includes a radial wall covering the worn face of the flange. The wall includes a hole that comes into register with the hole in the flange and having thickness for compensating the wear of the face of the flange. The member further includes an abutment for bearing against the periphery of the flange, and being fastened to the flange by adhesive.