Abstract: A mounting system for a gas turbine engine includes a mounting linkage assembly and a tangential link positioned generally transverse to the mounting linkage assembly. The mounting linkage assembly reacts at least a thrust load. The tangential link reacts at least a vertical load, a side load, and a torque load of the gas turbine engine.

Abstract: A core case section for a gas turbine engine a multitude of discreet radial extending 2.5 bleed ducts defined in part by a rear structural wall.

Abstract: A combustor assembly for a turbine engine includes a combustor liner, a flow sleeve and a plenum ring. The flow sleeve surrounds the combustor liner to form an annulus radially between the combustor liner and the flow sleeve. The flow sleeve has a plurality of rows of cooling holes. The plenum ring radially surrounds the combustor liner in the annulus. The plenum ring has a plurality of bypass tubes for directing axial air flow and radial flow chambers for directing radial air flow.

Abstract: A turbine shroud includes a ceramic shroud ring configured to surround a plurality of turbine rotor blades, a plurality of slots circumferentially distributed around the ceramic shroud ring, a forward metallic support ring, a plurality of tabs attached to a forward edge of the forward metallic support ring configured to engage a turbine support case, and a plurality of tabs attached to an aft edge of the forward metallic support ring received by the slots of the ceramic shroud ring. Only two axially extending radial surfaces of each of the tabs attached to the aft edge of the forward metallic support ring are configured to contact the slots of the ceramic shroud ring.

Abstract: A rotor blade for a gas turbine engine includes an airfoil that extends in span between a root and a tip region. A leading edge and a trailing edge of the airfoil section extend between a chord line of the airfoil. A sweep angle is defined at the leading edge of the airfoil section, and a dihedral angle is defined relative to the chord line of the airfoil section. The sweep angle and the dihedral angle are localized at the tip region of the airfoil section.

Abstract: A gas turbine engine is provided with turbine sealing structures including knife edge seals which extend at an angle relative to an axial center line of the engine. Each knife edge seal is associated with a control pocket defined between a radially inner surface and a spaced radially outer surface. The control pockets and their associated knife edge seals create a difficult flow path to prevent leakage into radially inner portions of the turbine section.

Abstract: According to the present invention, an annular face seal arrangement for a gas turbine engine is provided that includes a mounting ring, a mounting member, a rotor, and a stator. The mounting ring has a width extending between a first axial end and a second axial end. The mounting member has a clamp portion and a biasing portion. The clamp portion extends axially between a first clamping surface and a second clamping surface. The biasing portion includes a first segment and a second segment. The second segment has a width and a rotor contact surface. The rotor has a rotor seal surface and a clamp portion having a width. The stator has a stator seal surface that is aligned with the rotor seal surface. The mounting ring is disposed radially inside of the rotor and radially inside of at least part of the biasing portion, and is disposed in contact with the second clamping surface of the clamp portion. The rotor contact surface of the biasing portion is disposed in contact with the rotor.

Abstract: A powder metallurgy method includes the steps of forming a member, such as a work piece or an aerospace component, from a titanium alloy powder. The average size of a carbide phase in the titanium alloy powder is controlled in order to control an average size of a carbide phase in the member. In one example, an amount of carbon within the titanium alloy and size of the carbide phase are selected to provide a desirable balance of good hot workability, resisting formation of an alpha-titanium phase within the member and a desired level of fatigue performance.

Abstract: A low density abradable coating for use as a seal material is formed by depositing a metal bond coat followed by depositing a low density, porous, abradable metal seal layer on the bond coat. The seal layer is formed by co-depositing metallic sponge particles and metal precursor particles. The metal precursor particles decompose during heat treatment leaving fine pores distributed throughout the microstructure.

Abstract: A compressor rotor seal includes a bare radial inward end of a rotor vane in proximity to a rotor surface coated with an abrasive material.

Abstract: An abrasive coating for rotor shafts that interact with cantilevered vanes to form an abradable air seal in a turbine engine. The abrasive coating including a metal bond coat layer on the rotor shaft, and an abrasive top coating bond coat layer for contact with vanes during operation of the rotor shaft, the abrasive coating including a plurality of abrasive grit particles in a matrix. the abrasive grit particles are selected from the group consisting of cubic boron nitride (CBN), zirconia, alumina, silicon carbide, diamond and mixtures thereof.

Abstract: An abrasive coating on a rotor shaft interacts with cantilevered vanes to form an abradable air seal in a turbine engine. The abrasive coating includes a metal bond coat, and an abrasive layer containing a plurality of abrasive CBN grit particles in a ceramic matrix.

Abstract: Trivalent chromium conversion coatings are provided on a metal substrate wherein the trivalent chromium conversion coating has a halogen content of 1 atom % maximum.

Abstract: A process for forming a dense abrading thermally insulating coating on a rotor shaft in a gas turbine engine is described. The process comprises fixturing the rotor shaft to allow it to rotate about its axis and plasma spraying the coating on the rotor shaft. The coating comprises a zirconia based ceramic top coat layer on a metallic bond coat.

Abstract: An abrasive coating on a rotor shaft interacts with cantilevered vanes to form an abradable air seal in a turbine engine. The abrasive coating includes a metal bond coat, and an abrasive layer containing a plurality of abrasive grit particles in a ceramic matrix. The grit particles may be selected from cubic boron nitride (CBN), zirconia, alumina, silicon carbide, diamond and mixtures thereof.

Abstract: A seal for a gas turbine engine component having an airfoil with a radial outward end and a radial inward end, one of which is bare metal. A seal member is adjacent to the bare metal end of the airfoil wherein the seal member is coated with a abrasive layer having a laser engraved surface with the top of the surface has less than about 5% of the surface area of the base of the abrasive layer.

Abstract: An abrasive coating is formed on a rotor shaft that rotates with respect to cantilevered vanes. The coating is formed by thermal spray techniques and comprises a ceramic layer on a metal bond coat. The coating surface is roughened by crush grinding or by grit blasting to increase the abradability of the surface.

Abstract: A split disk assembly for a gas turbine engine includes a forward disk section and an aft disk section, the aft disk section engageable with the forward disk section to retain a multitude of rotor blades therebetween.

Abstract: A method for using a heat exchange system in operating equipment in which a working fluid is utilized in providing selected operations thereof, including for use in lubricating systems for aircraft turbofan engine equipment, the heat exchange system for providing air and working fluid heat exchanges to cool the working fluid at selectively variable rates in the operating equipment developed airstreams. A heat exchanger core is provided in a controlled air flow duct system opening at its entrance to those airstreams and having its outlet end opening downstream in those airstreams.

Abstract: A case for a gas turbine engine includes a single-piece case that includes a combustor case portion and a turbine case portion that is integrally formed with the combustor case portion as one piece. In one example, the single-piece case includes a transition duct/mid-turbine frame case portion is integrally formed with the combustor and turbine case portions.