United Technologies Corporation

Other

United States of America as Represented by the Secretary of the Air Force

United Technologies Operations

Joseph C Burge

Richard J Tiernan, Jr

Sidney A Peerless

Michael G McCaffrey

Joel H Wagner

Shankar S Magge

Allan R Penda

Yehia M El-Aini

Daniel W Major

Daniel R Sabatino

William E Dixon

Edward Torres

Carl E Meece

Norman Paul Poire

Lee E Hensen

Gas turbine engine systems involving mechanically alterable vane throat areas are provided. In this regard, a representative vane for a gas turbine engine includes: a leading edge; a trailing edge; a suction side surface extending between the leading edge and the trailing edge; a cavity having an aperture located in the suction side surface; and a barrel located within the cavity and being moveable therein such that movement of the barrel alters an extent to which the barrel protrudes through the aperture.

Gas turbine engine systems involving mechanically alterable vane throat areas

A shroud for supporting a variable vane in a compressor section for a gas turbine engine includes a pair of components which are secured together by a threaded fastener extending into a blind hole in one of the two components. Additional material forms on the component which includes the blind hole extends beyond a pivot axis for the vane such that there is an increased length of engagement between the threaded fastener and the blind hole. The blind hole eliminates the need for multiple nuts and reduces the radial height of the shroud to save additional weight.

Small radial profile shroud for variable vane structure in a gas turbine engine

A turbine engine has a circumferentially segmented shroud within a case structure. Each shroud segment is mounted for movement between an inboard position and an outboard position. One or more springs bias the shroud segments toward their inboard positions. One or more valves are positioned to vent one or more volumes so as to counter the spring bias to shift the shroud segments to their outboard positions.

Blade clearance control

An apparatus and method for cooling an impeller for an axi-centrifugal compressor of a gas turbine engine. A curvic coupling joint is used between the aft end of the impeller and the turbine shaft. A generally annular disk-shaped shield structure is attached to the aft face of the impeller and extends from the impeller rim to the turbine shaft just aft of the curvic coupling. The shield structure is offset from the aft face of the impeller thereby forming a cavity therebetween. The shield structure has a plurality of circumferentially space apertures through which cooling air passes to first circulate through the cavity by means of radial vanes to cool the impeller, then exit through the curvic coupling joint. The cooling air is then directed afterward along the turbine shaft to be used to cool high-pressure turbine blades.

Apparatus and method for cooling Axi-Centrifugal impeller

An improved spool for a compressor of a gas turbine engine comprising spacer disks disposed between adjacent rotor disks, and a plurality of hollow tie rods that extend through the rotor disks and the spacer disks. The tie rods are tensionable to provide axial preload between the disks, and also conduct active air for thermally conditioning spool components. The spacer disks have an outer portion with a catenary shape that extends between the peripheral rims of the adjacent rotor disks. Sealing teeth located on the outer portion are preferably radially aligned with a web portion that extends radially inward from a center region on the outer portion. Each tie rod has at least one aperture through its wall for transmitting active air. In one embodiment, active air is injected through apertures in the forward stub shaft of the spool and is urged into at least one space between a rotor disk and a spacer disk, then into, along, and out of the hollow tie rods. In another embodiment active air is injected into open forward ends of the tie rods and subsequently dispersed through apertures in the tie rods. The last stage of the compressor includes a pair of mating forward and aft rotor disks with a cavity formed between the disks. A portion of the active air discharged from the tie rods is circulated in the cavity and into the hollow blades which are integral with the rotor disks. A compressor discharge pressure seal extends axially to the last stage aft rotor disk. The seal has teeth which are oriented generally axially, and preferably have a fishmouth configuration.

Gas turbine compressor spool with structural and thermal upgrades

An otological drain tube which is capable of being implanted in the ear drum to vent the differential pressure that can build up across the ear drum as the result of blocked eustachian tubes, a buildup of internal liquid pressure in the inner ear due to illness, or a change in atmospheric pressure. The otological tube comprising: a cap-plug, a central body, an end cap. And a means for permanently securing the elements together. The assembly of these elements creates a plurality of lumens of a requisite cross sectional area within the otological drain tube body. The primary object of the invention is to produce an otological drain tube which is extremely effective and economical to produce.

Otological drain tube

A rotor blade for a turbine engine rotor assembly is provided comprising a root, an airfoil, a platform, and apparatus for damping vibrations in the airfoil. The airfoil includes a pocket formed in a chordwise surface. The apparatus for damping vibrations in the blade includes a damper and a pocket lid. The damper is received within the pocket between an inner surface of the pocket and the pocket lid. The pocket lid is attached to the airfoil by conventional attachment apparatus and contoured to match the curvature of the airfoil.

Turbine engine rotor blade vibration damping device

In an assembly of spaced rotor discs, mounted to rotate about a common axis, one or more discs are provided that are removably joined to the other discs but so joined, away from potential fatigue points in the disc webs. Thus a pair of discs have a spacer arm extending therebetween in contact therewith. Further, each disc of such pair, has a flange that extends between the discs toward the flange of the other disc, which flanges (and thus the discs) are removably bolted together at a junction removed from the disc webs for greater disc durability and lower replacement costs thereof. The so joined flanges and spacer arm (which is held in a piloted joint under compressive pre-load) define an annular cavity around a pair of the discs and thus redundant structural support therebetween.

Joining means for rotor discs

The compressor customer bleed includes annular spaced walls defining the bleed manifold and staggered openings bleeding compressor air externally of the gas turbine engine avoiding a clear path for a hot particle to the most outer wall and including a thickened wall serving as a heat sink to dissipate the heat of the particle.

Compressor bleed manifold for a gas turbine engine

The segmented inner case supporting the stator vanes of the compressor is tied to an axially split outer case of a gas turbine engine by utilizing a segmented ring-like element forming a rail with bosses at either end and intermediate the ends and having a relatively thin resilient body between bosses capable of flexing in response to thermal loads. Bolts extending through the outer case are threaded into the bosses. This arrangement facilitates the assembly and disassembly of the compressor section.

Case tying means for gas turbine engine

A 'dog bone' in cross section shaped seal mounted in adjacent grooves of segmented stator vanes span the axial extent covering a single casted three rows of vanes. The outer diameter shroud of the stator vane defines a portion of an enclosed cavity subjected to leakage flow in and out of the gas path where the flow reverses at some axial extent of the seal such that the forces acting at different locations on the seal reverses. In a preferred embodiment, the bulbous ends of the 'dog bone' are solid coated metallic material.

Segmented stator vane seal

A drain tube for implanting in an ear drum is formed having a center body with a circular flange on each end with a cover member over each flange forming a chamber therewith. The cover members have openings therein and the center body has drain holes therethrough having a high L/D ratio to permit proper drain flow from an ear and prevent undesirable flow into an ear. A tab is used for insertion and a method of forming the drain holes is set forth wherein grooves are placed in the surface of a large opening and the grooves closed to form drain holes by a plug placed in the large opening.

The thermal environment to which the engine's compressor case is subjected is controlled and the coefficient of expansion selected for the stator and rotor components are suitably matched to improve engine performance by reducing radial clearances during part power engine operations. A thermal barrier around the engine case by virtue of a heat shield attached to the customary case flanges and the coaxially mounted spaced stator vanes forming an inner case defining cavities sandwiching the engine case and holds the metal temperature at a value higher than would otherwise be provided. The thermal barrier serves to isolate the outer case from the engine's fan discharge air and the engine's core gas path.

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Palm Beach Gardens, FL, United States of America

Joseph C Burge