Abstract: A collar seal is configured for a turbine nozzle vane. The seal includes a retainer having a circumferential airfoil contour conforming with the airfoil contour of the vane. A flexible leaf is surrounded by the retainer and fixedly joined thereto. A woven sheath encases the leaf and is fixedly joined to the retainer. In an exemplary embodiment, the collar seal surrounds one end of a ceramic turbine vane mounted in a metal supporting band.

Abstract: A low-pressure compressor exit stator assembly for a gas turbine includes a multiple of vanes mounted between a full hoop outer shroud and an inner shroud. The outer shroud and inner shroud are mounted to the intermediate case structure. A seal is located between the outer shroud and a bleed duct structure to accommodate thermal growth mismatch.

Abstract: There is disclosed a gas turbine shroud structure comprising: a shroud support component 12 attached to an inner surface of a gas turbine casing 3; a shroud segment 14 divided in a peripheral direction and supported by the inner surface of the shroud support component; and a heat-resistant restricting spring 18 held between the shroud segment and shroud support component to urge the shroud segment inwards in a radial direction. The shroud segment 14 is formed of a ceramic composite material, and includes a coating layer 15 having heat insulation and impact absorption effect on an inner surface of the segment.

Abstract: The clearances between an array of turbine blades and its surrounding blade track may be controlled by an expansion control material system supporting the blade tracks. The blade track support hoop is placed in tension by the expansion control material placed therein and the expansion control material is placed in compression.

Abstract: A gas turbine frame has inner and outer annular bands, respectively, joined together by generally radially extending struts therebetween. A radially outer conical support arm extends radially outwardly from the outer band and a radially inner conical support arm extends radially inwardly from the inner band. Circumferentially spaced apart inner and outer openings are disposed in the inner and outer conical support arms, respectively. Each of the struts has at least one radially extending hollow passage which extends through the inner and outer bands. The frame is a single piece integral casting. The inner and outer conical support arms have an equal number of the inner and outer circumferentially spaced apart openings. The inner circumferentially spaced apart openings are equi-angularly spaced apart and the outer circumferentially spaced apart openings are equi-angularly spaced apart.

Abstract: A gas turbine frame has inner and outer annular bands, respectively, joined together by generally radially extending struts therebetween. A radially outer conical support arm extends radially outwardly from the outer band and a radially inner conical support arm extends radially inwardly from the inner band. Circumferentially spaced apart inner and outer openings are disposed in the inner and outer conical support arms, respectively. Each of the struts has at least one radially extending hollow passage which extends through the inner and outer bands. The frame is a single piece integral casting. The inner and outer conical support arms have an equal number of the inner and outer circumferentially spaced apart openings. The inner circumferentially spaced apart openings are equi-angularly spaced apart and the outer circumferentially spaced apart openings are equi-angularly spaced apart.

Abstract: A segmented vane support structure for use in a gas turbine engine having an engine casing, includes a single piece inner ring and separated front outer and rear outer rings. The vane segments circumferentially abut to form a stator ring which is clamped at the respective opposed outer ends by the front outer and rear outer rings therebetween onto the inner ring. The front outer and rear outer rings are axially restrained to the engine casing by a retaining ring which is fitted in an inner annular groove of the engine casing. Each vane segment has a lug member at its outer diameter which radially and slidably engages in a slot of the engine casing to provide angular positioning of the vane segments within the engine casing and to transmit circumferential vane load into the engine casing. This support structure arrangement transmits circumferential loading to the engine casing and isolates radial loading from the engine casing caused by thermal growth changes of the vane segments.

Abstract: A segmented vane support structure for use in a gas turbine engine having a engine casing, includes a single piece inner ring and separated front outer and rear outer rings. The vane segments circumferentially abut to form a stator ring which is clamped at the respective opposed outer ends by the front outer and rear outer rings therebetween onto the inner ring. The front outer and rear outer rings are axially restrained to the engine casing by a retaining ring which is fitted in an inner annular groove of the engine casing. Each vane segment has a lug member at its outer diameter which radially and slidably engages in a slot of the engine casing to provide angular positioning of the vane segments within the engine casing and to transmit circumferential vane load into the engine casing. This support structure arrangement transmits circumferential loading to the engine casing and isolates radial loading from the engine casing caused by thermal growth changes of the vane segments.

Abstract: At least one airfoil shaped vane made of a low ductility material, for example a ceramic base material such as a ceramic matrix composite or an intermetallic material such as NiAl material, is releasably carried in a turbine vane assembly including inner and outer vane supports by at least one high temperature resistant compliant seal. The seal isolates the vane from at least one of the vane supports and allows independent thermal expansion and contraction of the vane in respect to the support.

Abstract: A device for positioning of nozzles of a stator stage and for cooling of rotor discs in gas turbines, of the type including nozzle segments consisting of several airfoils. Each nozzle segment is connected at its top to an outer ring to define a chamber containing cooling air and is positioned at its base on an inner ring. The nozzle segments form a circumferential array about the axis of the gas turbine. In each airfoil of the nozzle segment, there is provided at least one tube, which is inserted in a corresponding duct inside the airfoil for delivering cooling air to the first and second stage discs.

Abstract: A network switch for network communications includes at least one data port interface supporting a plurality of data ports transmitting and receiving data. A CPU interface is configured to communicate with a CPU, and an internal memory communicates with the at least one data port interface. A memory management unit is provided for communicating data from at least one data port interface and the memory. A communication channel is provided, for communicating data and messaging information between the at least one data port interface, the memory, and the memory management unit. The configuration of the network switch also includes a fast filtering process, with the fast filtering processor filtering packets coming into the at least one data port interface. Selective filter action is taken based upon a filtering result.

Abstract: An outer housing for supporting a stator assembly in a gas turbine engine, wherein the support includes two outer ring sections between which the vane segments of the stator assembly are mounted to form a ring within the cylindrical support. A central ring section is provided between the outer ring sections and is connected thereto by means of spokes. The central ring section locates the structure within the engine casing. The spokes extend at an angle to the axis of the stator assembly and define openings therebetween whereby thermal distortion between the outer rings and the central ring is attenuated.

Abstract: The present nozzle and shroud assembly mounting structure configuration increases component life and reduces maintenance by reducing internal stress between the mounting structure having a preestablished rate of thermal expansion and the nozzle and shroud assembly having a preestablished rate of thermal expansion being less than that of the mounting structure. The mounting structure includes an outer sealing portion forming a cradling member in which an annular ring member is slidably positioned. The mounting structure further includes an inner mounting portion to which a hooked end of the nozzle and shroud assembly is attached. As the inner mounting portion expands and contracts, the nozzle and shroud assembly slidably moves within the outer sealing portion.

Abstract: An airfoil and nozzle assembly including an outer shroud having a plurality of vane members attached to an inner surface and having a cantilevered end. The assembly further includes a inner shroud being formed by a plurality of segments. Each of the segments having a first end and a second end and having a recess positioned in each of the ends. The cantilevered end of the vane member being positioned in the recess. The airfoil and nozzle assembly being made from a material having a lower rate of thermal expansion than that of the components to which the airfoil and nozzle assembly is attached.

Abstract: A turbine nozzle and shroud assembly having a preestablished rate of thermal expansion is positioned in a gas turbine engine and being attached to conventional metallic components. The metallic components having a preestablished rate of thermal expansion being greater than the preestablished rate of thermal expansion of the turbine nozzle vane assembly. The turbine nozzle vane assembly includes a plurality of segmented vane defining a first vane segment and a second vane segment. Each of the first and second vane segments having a vertical portion. Each of the first vane segments and the second vane segments being positioned in functional relationship one to another within a recess formed within an outer shroud and an inner shroud. The turbine nozzle and shroud assembly provides an economical, reliable and effective ceramic component having a preestablished rate of thermal expansion being less than the preestablished rate of thermal expansion of the other component.

Abstract: A guide vane ring having guide vanes is provided on two rings which are arranged coaxially and radially at a distance from one another. These two rings are held on assigned housing sections which are subjected to thermally different expansions. One ring is rigidly connected with the guide vanes, and the other ring is connected with the free ends of the guide vanes by the compressive force of a tensioning element. In addition, the other ring is divided into several piston-ring-type elements which axially follow one another and which are each slotted approximately axially and form a cylindrical sliding surface for the whole free ends of the vanes. The tensioning element is arranged on the one group of assigned housing sections while forming a radial compressive force distributed uniformly on the elements.

Abstract: A mounting system for a stator vane assembly is disclosed for mounting a vane and its cooling system to minimize the thermal and mechanical stresses imparted to the vane structure. Platforms are attached to opposite ends of the vane, which has a hollow interior and a cooling gas distributor located within the interior of the vane. One end of the cooling gas distributor is also fixedly attached to one of the vane platforms, while the opposite end of the cooling gas distributor is attached to a casing, not to the opposite vane platform. Spaces are defined between the cooling gas distributor and the vane platform, as well as the casing and the vane platform. A gas inlet sleeve may be attached to the cooling distributor and serve as its attachment to the external casing.

Abstract: A split casing (10) of low, coefficient of expansion material, has internal upstream facing slots (12). A semicircular shroud ring (16), of high coefficient of expansion material, carries the vanes (18), and has a tongue (20) sliding within casing slot (12). Casing stop lugs (14) and shroud stop lugs (22) have mutually abuttable surfaces (70,72) to antirotate the shroud within the casing. The distance between abuttable surfaces of the casing is greater than the distance between abuttable surfaces of the shroud ring at room temperature, but the distance are the same at the long time operating temperature.

Abstract: Methods and apparatus for vane segment alignment and support in a combustion turbine. The vane segment alignment device is comprised of a rotatable, eccentric bushing and pin which is inserted into a slot of the vane segment. The eccentric bushing is further comprised of a cover plate which is peened to a splined torque plate thus holding the eccentric bushing in place against the vane segment but allowing for fine adjustments of the alignment of the vane segments. The vane segment support and alignment device provides for efficient and economical adjustment of the vane segments especially in electric generating plants where combustion turbines undergo high peak load operation. Additionally, the vane segment support and alignment device transfers torques and moments generated by aerodynamic flow from the vane segments to an inner cylinder thereby reducing the amount of misalignment due to aerodynamic drag.

Abstract: A nozzle guide vane structure for a gas turbine engine in which the vane comprises an aerofoil portion and inner and outer platforms. The non gas-contacting surfaces of the platforms are provided with forward and rearward sealing flanges sealed to adjacent engine structure to provide spaces sealed against ingress of working fluid between each pair of forward and rearward flanges. The mounting means for the vane projects into these spaces from the non gas-contacting surface.

Abstract: The invention comprises a turbine nozzle and nozzle support assembly in which a turbine nozzle support seating flange is positioned adjacent to a turbine nozzle mounting flange and pins are used to affix the turbine nozzle to the nozzle support. Retainer strips are installed into the nozzle support to trap the pins in engagement with the mounting and seating flanges. In the preferred embodiment of the invention seal means is provided for restricting the airflow through the completed turbine nozzle and nozzle support assembly.

Abstract: A turbine airfoil, such as a turbine blade or nozzle vane, is disclosed for use in a high temperature environment. The airfoil comprises a center post and at least one side piece. The side piece is attached to the center post so that, at room temperature, a clearance between the center post and the side piece permits limited relative movement between the post and side piece. This clearance space, however, is dimensioned so that at the operating temperature of the turbine, thermal expansion of the center post and/or side piece substantially eliminates the clearance as well as the limited movement between the center post and side piece.

Abstract: A turbine ring comprises an annular support in two parts and a ring of ceramic sectors for sealing purposes. The two parts of the support are interconnected by an axial groove and a sliding male part. Each part comprises an axial annular groove in which is engaged a respective axial edge portion of each sector. The cooperating internal radial faces respectively of the edge portion and of the groove are circumferential, while the radially outer radial face of each sector comprises at least one flat zone. The corresponding face of each groove may be polygonal.

Abstract: A load transference structure between a core engine and power turbine comprises a plurality of circumferentially spaced apart radially extending spoke members, each of which is attached at its radially inner extent to first and second axially spaced apart support panels. Each of the spoke members is provided with a circumferentially extending portion positioned adjacent the panels which are adapted to cooperate with a circumferentially adjacent spoke member in such a manner that radial but not circumferential load transfer is permitted to take place between them. The load transfer structure permits load transfer to take place between the core engine casing and the power turbine.

Abstract: A sensor guide tube for bridging a plenum between a turbine housing and a nozzle assembly includes an offset portion whose offset is substantially equal to half the distance which the nozzle assembly is displaced upstream for clearance restoration. After clearance restoration, the guide tube is rotated 180 degrees to align its lower end with an opening in the nozzle assembly without requiring a sharp bend in the axis of the sensor guide tube.

Abstract: To form a turbine engine component, metal airfoils are positioned in an annular array. Outer end portions of the airfoils are embedded in a wax outer shroud ring pattern and inner end portions of the airfoils are embedded in a wax inner shroud ring pattern. A mold is formed by covering the metal airfoils and the shroud ring patterns with ceramic mold material. The wax of the shroud ring patterns is then removed from the mold to leave inner and outer shroud ring mold ring cavities. The shroud ring mold cavities are filled with molten metal which is solidified to form inner and outer shroud rings interconnecting the airfoils. To accommodate thermal expansion of the airfoils relative to the shroud rings, a slip joint is provided between at least one end portion of each of the airfoils and a shroud ring. To enable the slip joint to be formed, molten metal solidifies in the shroud ring to be formed, molten metal solidifies in the shroud ring cavities free of metallurgical bonds to the airfoils.

Abstract: Means for axially attaching a variable turbine vane array to a combustor in a gas turbine engine is disclosed. The attachment means are configured so that substantially all of the axial load on the array is transmitted through the walls of the combustor.

Abstract: A stator seal land and supporting structure suitable for use in an axial flow rotary machine is disclosed. Included are concepts for matching the thermal response of the seal land to the rotor assembly and for inhibiting rolling or twisting of the seal land under changing thermal conditions. The seal land is free to expand and contract in diameter independently of the supporting stator structure.

Abstract: A hollow thermally-conditioned turbine stator nozzle 12 is set forth to distribute and guide combustion gases from a forward inlet 15 to the turbine rotor blades 18. The nozzle includes a plurality of vanes 20 arranged annularly in the turbine. Each vane 20 has a body 22 to guide the fluid, the body being supported by outer and inner ends 34,64. The outer and inner ends each have a forward lug 44,74 and a rear lug 46,76. A floating support is provided for the vanes and includes forward shoulders 110,136 to engage the forward lugs and prevent rearward movement of the vanes and rear shoulders 114,142 having notches 116,144 to receive and confine the tangential movement of the vanes. Each vane also includes a hollow core 152 to pass a portion of fluid and reduce thermal stresses on the vane. The vane is preferably made of silicon nitride ceramic.

Abstract: An arrangement is provided for statically sealing between the inner shroud 12 of individual stator vanes of a gas turbine, and a seal housing 14 provided with a radially outwardly open channel 22 into which is received a plurality of outer seal segments 24, and which includes a radially outwardly open channel 33 throughout the length of each outer seal segment and with a plurality of inner seal segments received in the channel, with the outer seal segments being biased outwardly by the springs 36, and the inner seal segments 34 being biased outwardly into sealing contact with individual inner shrouds by the concentric inner compression springs 38 carried by the outer seal segments 24.

Abstract: An annular inlet stator vane assembly positioned upstream of the bypass and core flow passages of a turbofan jet engine is provided with energy-absorbing bendable vanes which are fixedly secured to an outer support shroud assembly and extended through slots in an inner shroud assembly. The inner ends of the vanes are releasably secured to the inner shroud by specially designed deformable retention clips which are welded to the inner vane ends and shroud. These features of the vane assembly cooperate to limit the circumferential extent of inlet vane damage caused by engine-ingested objects, and to help prevent damaged vane assembly structure from being drawn through the core passage of the engine or blocking such core passage.

Abstract: A turbine blade, more particularly a turbine nozzle vane, for a gas turbine engine, consisting of a structural metallic core enveloped at a distance by a ceramic blade jacket. Through provision a primarily highly heat-resistant, ceramic metal compound structure of the blade, the blade jacket is compressively prestressed spanwise between the blade shrouds in the inoperative condition of the turbine and where said compressive prestressing of the blade jacket, produced during assembly and which exists in the inoperative condition, is at least maintained during the operation condition by selecting for the metallic core of the blade, a material of average coefficient of thermal expansion and for the metallic inner and outer shrouds of the blade, and for the inner and outer metallic, sleeve-like components producing the axial prestress, materials of higher coefficients of thermal expansion, and for the highly thermally stressed ceramic blade jacket, a material of average coefficient of thermal expansion.

Abstract: A shroud ring (22) for a gas turbine engine having a stator structure (14) and an array of inwardly extending stator vanes (20) is disclosed. The shroud ring is circumferentially segmented to accommodate inward and outward movement of the vanes and the stator structure. The shroud ring has a segmented outer ring (28) and a segmented inner ring (30). The outer ring engages the vanes to support the vanes and to position the inner ring and an integral seal member (24) about a rotor structure. In one embodiment the vanes (20) are rotatable about a spanwise axis S. In a second embodiment, the vanes (120) are integrally attached to the shroud ring (122).

Abstract: A nozzle guide vane assembly for a turbomachine comprises a plurality of segments mounted in an outer casing by means of pins which locate in slots. The slots are angled relative to radial planes so that the reaction force exerted by the pin normal to the length of the slot produces a tangential and radial force. The radial force opposes the couples produced on each segment by gas loads and the tangential reaction force.

Abstract: A turbine nozzle vane suspension arrangement of the type wherein the nozzle vanes are suspended by their outer shroud segments from the outer turbine casing and are suspended by their vane ring segments from an annular vane carrier. The vane carrier has axially spaced apart and radially-extending wall sections cooperating with root-ends projecting radially from the vane ring segments. The vane ring segments and their root-ends, and the vane carrier and its wall sections define between them an annular channel for receiving cooling air. The forward wall section of the vane carrier is axially resilient and bears resiliently against the forward vane root-end to provide a seal for the annular channel. The resilient forward wall section can be flexed axially rearwardly to permit the vane ring segments to be mounted on the vane carrier.

Abstract: A case assembly for supporting stator vanes of a gas turbine engine is disclosed. Various construction details and techniques which enable support of the vanes across the flow path of the engine are discussed. The construction details and techniques which provide damping to the vanes and accommodate differential thermal expansion between the vanes and the case assembly are developed. The vane mounting system disclosed is built around the concept of enabling sliding friction between components of the case assembly to dampen vane vibrations.

Abstract: A combustion turbine vane assembly is shown having a retainer block mounting each individual cooled turbine vane. The retainer block has a pair of axially extending opposed rails for sliding receipt of complementary engaging opposed rails on the outer surface of the vane shroud. The block defines a cooling air aperture therethrough. The inner surface of the block is supported a predetermined distance from the opposed surface of the vane shroud defining a cooling fluid receiving chamber therein. The separated distance permits controlled cooling air leakage into a space between the block and the outer shroud. The radial walls of adjacent outer shrouds have complementary axial grooves machined therein and a seal member is disposed therein to bridge the gap and minimize loss of the cooling air.

Abstract: A seal member for a gas turbine engine is disclosed. Various construction details which ensure an adequate fatigue life and increase the sealing effectiveness of the member during engine operation are developed. A metal diaphragm extends from a nonrotating structure inwardly of the engine flow path to engage the stator vanes at engine operating temperature.

Abstract: A high temperature cushion for a low ductility member or article, such as of a ceramic material, comprises a flexible, fibrous core of a resilient, refractory material, and a metal foil of a high temperature material covering the core. The foil has ductility sufficient to be formed about the core, the core and foil being further characterized by the ability to withstand a temperature of at least about 1,000.degree. F. (about 540.degree. C.) Forms of such a cushion include an annular bushing and various shaped holders and interface members. Such a cushion can be included as a part of a high temperature operating article which comprises metal portions functioning as backing or holding members, a low ductility article or member such as of a ceramic material, a high temperature cushion interfacing between the metal portions and the member, and a securing member mechanically joining the metal portion with the low ductility member.

Abstract: A nozzle guide vane assembly for a gas turbine engine comprises a circumferentially extending array of angularly spaced apart aerofoils each having projections adapted to engage with structure to retain the aerofoil in its longitudinal direction. Inner and outer platform members are separate from the aerofoils and each comprises a thicker support skin and a thinner inner skin. Both skins have aerofoil shaped apertures through which the aerofoils project, the support skin retaining the aerofoil against twisting, circumferential or axial loads and the inner skin serving to define a respective boundary of the gas flow through the assembly. The aerofoil is free to slide through the apertures sufficiently to permit relative expansion in a direction longitudinal of the aerofoil, and sealing means are associated with the inner skins and provide a seal between the skins and the aerofoils.

Abstract: A stator blade for a gas turbine engine comprises a stack of laminar sections and means adapted to resiliently hold the sections together in compression to allow for differential expansion between the stack and supporting structure.

Abstract: A variable area turbine vane, for use in high temperature aircraft gas turbines, having a load carrying spar with a heat shield member surrounding the load carrying member. The heat shield member is positioned in grooves which permit spanwise and cordwise expansion of the heat shield member. Cooling air enters the hollow load bearing member and passes through holes in the load bearing member and heat shield member.

Abstract: A gas turbine consisting of alternate stators and rotors so fabricated that all stator vanes, all rotor blades and all blade tip seals may be removed through the engine case without requiring disassembly or translation of major portions of the gas turbine engine.