Siemens Energy, Inc.

Siemens Aktiengesellschaft

Mrinal Munshi

Ching-Pang Lee

Yevgeniy P Shteyman

Steffen Udluft

Jae Y Um

Matthew R Porter

John W Finneran

Jonathan M Leagon

Daryl J Graber

Humberto A Zuniga

Petya M Georgieva

Matthew H Lang

Adam C Pela

Paul Bradley Davis

Steven James Vance

Gm Salam Azad

James Morin

Harshawardhan S Bhide

Shivashankar Vinay Amatiganahally

Eric L Roach

Todd F Young

Douglas R Roth

Benjamin P Swanson

Thomas N Silvey

Siegmund DÜLL

A gas turbine engine including an outer case and an exhaust gas passage defined within the outer case for conducting an exhaust gas flow from a turbine section of the gas turbine engine. A cooling channel is associated with an outer surface of the outer case, the cooling channel having a channel inlet and a channel outlet. An air duct structure is provided and includes an inlet end in fluid communication with the channel outlet and includes an outlet end in fluid communication with an area of reduced pressure relative to the air duct structure inlet end. An exit cavity is located at the air duct structure outlet end, wherein the exit cavity effects a reduced pressure at the outlet end to draw air from the cooling channel into the air duct.

Gas turbine outer case active ambient cooling including air exhaust into sub-ambient cavity

An attachment system for attaching at least one exhaust diffuser downstream from a turbine assembly in a gas turbine engine is disclosed. The attachment system may include at least one attachment flange extending from a downstream edge and attached to a spring plate diffuser support structure and at least one attachment flange extending from side edges of the exhaust diffuser to couple sections of the exhaust diffuser together. The diffuser may also include a thermal barrier/cooling system for controlling a temperature of an outer case of the gas turbine engine. The thermal barrier/cooling system may form a flow path for an ambient air flow cooling.

Gas turbine engine with multiple component exhaust diffuser operating in conjunction with an outer case ambient external cooling system

A component in a gas turbine engine includes an airfoil extending radially outwardly from a platform associated with the airfoil. The airfoil includes opposed pressure and suction sidewalls, which converge at a first location defined at a leading edge of the airfoil and at a second location defined at a trailing edge of the airfoil opposed from the leading edge. The component includes a built-up surface adjacent to the leading edge at an intersection between the pressure sidewall and the platform, and at least one cooling passage at least partially within the built-up surface at the intersection between the pressure sidewall and the platform. The at least one cooling passage is in fluid communication with a main cooling channel within the airfoil and has an outlet at the platform for providing cooling fluid directly from the main cooling channel to the platform.

Airfoil having built-up surface with embedded cooling passage

A gas turbine vane containment cap is attached to an inboard surface of the vane inner shroud by penetrating flat weld filler, formed in a root gap between the cap and inner shroud. A semi-circular bead weld filler is formed outboard the penetrating weld filler closer to the vane exterior. Vane containment caps so welded are more resistant to in-service cracking along weldments.

Crack resistant turbine vane and method for vane containment cap attachment

Length adjustable braces (A-D) attached between a gas turbine exhaust diffuser () and an exhaust casing () along a horizontal joint () between upper and lower halves (A,B) of the outer diffuser shell. Brace lengths are adjusted to align bolt holes (A,B) in respective bolt bosses (A,B) on the upper and lower halves of the shell. The braces may be turnbuckles () welded to or releasably attached at one end to the shell and at the other end to the casing. Exemplary fittings on the diffuser shell and casing for the brace ends may be clevis fittings () or eye fittings (). The fittings may be configured to support both tension () and compression () of each brace. Two opposed fittings (A,B) across the joint may be configured for insertion of a respective clevis bolt (A,C) in both fittings from the same side.

Adjustable bracing apparatus and assembly method for gas turbine exhaust diffuser

A support ring for a row of vanes in an engine section of a gas turbine engine includes an annular main body portion for providing structural support for a row of vanes in the engine section, an aft hook, a forward wall, and a strong back plate. The aft hook extends from an aft side of the main body portion and is coupled to an outer engine casing for structurally supporting the support ring in the engine section. The forward wall extends generally radially outwardly from a forward side of the main body portion. The strong back plate spans between the forward wall and the aft hook and effects a reduction in dynamic displacement between the forward wall and the aft hook during operation of the engine.

Outer vane support ring including a strong back plate in a compressor section of a gas turbine engine

A method of repairing a turbine blade having a radially extending outer wall defining an internal cavity width and a blade tip. The method comprises removing at least a portion of the blade tip to form a repair surface and providing a tip cap having a radially outer side with an outer width that may be less than the internal cavity width, and having a radially inner side with an inner width that is substantially equal to or greater than the internal cavity width. The tip cap is positioned at the repair surface, and the tip cap is welded to the repair surface using a ductile welding material. A cap peripheral portion is formed by build-up welding around the tip cap, and a squealer portion is formed by build-up welding on the cap peripheral portion.

Method of providing a turbine blade tip repair

A film cooling structure formed in a component wall of a turbine engine and a method of making the film cooling structure. The film cooling structure includes a plurality of individual diffusion sections formed in the wall, each diffusions section including a single cooling passage for directing cooling air toward a protuberance of a wall defining the diffusion section. The film cooling structure may be formed with a masking template including apertures defining shapes of a plurality of to-be-formed diffusion sections in the wall. A masking material can be applied to the wall into the apertures in the masking template so as to block outlets of cooling passages exposed through the apertures. The masking template can be removed and a material may be applied on the outer surface of the wall such that the material defines the diffusion sections once the masking material is removed.

Component wall having diffusion sections for cooling in a turbine engine

A method of servicing an airfoil for use in a gas turbine engine. The airfoil assembly is defined by a base material and includes an airfoil and a platform from which the airfoil extends. A predetermined amount of the base material is removed from the airfoil assembly proximate to a fillet area of the airfoil assembly via water jet material removal. The fillet area comprises a junction between the airfoil and the platform and is located at an intersection between the airfoil and the platform. A remainder of the base material comprising base material of the airfoil assembly other than proximate to the fillet area is left intact.

Method of servicing an airfoil assembly for use in a gas turbine engine

A support ring for a row of vanes in an engine section of a gas turbine engine includes an annular main body portion to which a row of vanes is affixed for providing structural support for the vanes in the engine section, and an aft hook extending from an aft side of the main body portion with reference to a direction of air flow through the engine section. The aft hook is coupled to an outer engine casing for structurally supporting the support ring in the engine section. The support ring does not include a forward hook having a flange that extends axially from a forward or aft side of the forward hook with reference to the direction of air flow through the engine section.

Vane outer support ring with no forward hook in a compressor section of a gas turbine engine

A thermal barrier/cooling system for controlling a temperature of an outer case of a gas turbine engine. The thermal barrier/cooling system includes an internal insulating layer supported on an inner case surface, the internal insulating layer extending circumferentially along the inner case surface and providing a thermal resistance to radiated energy from structure located radially inwardly from the outer case. The thermal barrier/cooling system further includes a convective cooling channel defined by a panel structure located in radially spaced relation to an outer case surface of the outer case and extending around the circumference of the outer case surface. The convective cooling channel forms a flow path for an ambient air flow cooling the outer case surface.

Gas turbine engine with outer case ambient external cooling system

A component wall in a turbine engine. The component wall includes a substrate, a trench, and a plurality of cooling passages. The substrate has a first surface and a second surface opposed from the first surface. The trench is located in the second surface and is defined by a bottom surface between the first and second surfaces, a first sidewall, and a second sidewall spaced from the first sidewall. The first sidewall extends radially outwardly continuously from the bottom surface of the trench to the second surface. The first sidewall includes a plurality of first protuberances extending toward the second sidewall. The cooling passages extend through the substrate from the first surface to the bottom surface of the trench. Outlets of the cooling passages are arranged within the trench such that cooling air exiting the cooling passages is directed toward respective ones of the first protuberances.

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Orlando, FL, United States of America

Mrinal Munshi