Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: For controlling a target system, e.g. a gas or wind turbine or another system, operational data of a plurality of source systems are used. The operational data of the source systems are received and are distinguished by source system specific identifiers. By a neural network a neural model is trained on the basis of the received operational data of the source systems taking into account the source system specific identifiers, where a first neural model component is trained on properties shared by the source systems and a second neural model component is trained on properties varying between the source systems. After receiving operational data of the target system, the trained neural model is further trained on the basis of the operational data of the target system, where a further training of the second neural model component is given preference over a further training of the first neural model component.

Abstract: 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.

Abstract: Length adjustable braces (56, 56A-D) attached between a gas turbine exhaust diffuser (40) and an exhaust casing (34) along a horizontal joint (50) between upper and lower halves (40A, 40B) of the outer diffuser shell. Brace lengths are adjusted to align bolt holes (52A, 52B) in respective bolt bosses (43A, 43B) on the upper and lower halves of the shell. The braces may be turnbuckles (56, 67) 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 (62) or eye fittings (72). The fittings may be configured to support both tension (58) and compression (59) of each brace. Two opposed fittings (70A, 70B) across the joint may be configured for insertion of a respective clevis bolt (63A, 63C) in both fittings from the same side.

Abstract: An emissions control system (10) for a gas turbine engine (12) for reducing CO emissions at partial load of the gas turbine engine (12) is disclosed. In at least one embodiment, the emissions control system (10) may be formed from one or more compressed air exhausts (14) for exhausting compressed air into a hot gas pathway contained within a channel (18) formed by a transition (20) extending from a combustor (22) to a turbine assembly (24). In at least one embodiment, a plurality of seals (26) may extend circumferentially around the transition (20) providing a seal (26) between the transition (20) and a component of a downstream turbine assembly (24). One or more compressed air exhausts (14) may be positioned between adjacent seals (26). The compressed air exhausts (14) may be, but are not limited to being, channels (30), orifices (34) and metered spaces (32) having various shapes and configurations.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: For controlling a target system, operational data of a plurality of source systems are used. The data of the source systems are received and are distinguished by source system specific identifiers. By a neural network, a neural model is trained on the basis of the received operational data of the source systems taking into account the source system specific identifiers, where a first neural model component is trained on properties shared by the source systems and a second neural model component is trained on properties varying between the source systems. After receiving operational data of the target system, the trained neural model is further trained on the basis of the operational data of the target system, where a further training of the second neural model component is given preference over a further training of the first neural model component. The target system is controlled by the further trained neural network.

Abstract: A technique for improving the thermal protection against oxidation for a component in a gas turbine engine, for example, blades, row 1 vanes and row 2 vanes. The technique includes spraying a thin layer of alloy 230 on a base substrate of the component at those locations on the component where thermal protection against oxidation is desired. A metal bond coat layer is then deposited on the alloy 230 layer and a thermal barrier coating is deposited on the bond coat layer. The chromium, molybdenum, iron and tungsten in alloy 230 provide superior oxidation resistance, and the addition of lanthanum in the alloy 230 helps tailor thermal expansion with the thermal barrier coating resulting in higher spallation life.

Abstract: 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.

Abstract: 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.

Abstract: Length adjustable braces (56, 56A-D) attached between a gas turbine exhaust diffuser (40) and an exhaust casing (34) along a horizontal joint (50) between upper and lower halves (40A, 40B) of the outer diffuser shell. Brace lengths are adjusted to align bolt holes (52A, 52B) in respective bolt bosses (43A, 43B) on the upper and lower halves of the shell. The braces may be turnbuckles (56, 67) 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 (62) or eye fittings (72). The fittings may be configured to support both tension (58) and compression (59) of each brace. Two opposed fittings (70A, 70B) across the joint may be configured for insertion of a respective clevis bolt (63A, 63C) in both fittings from the same side.

Abstract: Manufacture of a gas turbine exhaust diffuser shell (40A/40B) to achieve a final cross-sectional shell geometry by forming an opening (76) in the shell to receive a strut shield collar (46); forming a compensating outward bowing (78) of the shell around the opening that departs from a desired final shell geometry in an amount and shape that compensates for a welding shrinkage when welding the collar in the opening; and welding the collar in the opening. This produces the desired shell geometry after the welding. The collar may be welded proximate an edge (74) of the diffuser shell, such as along an intersection of an axial plane with the diffuser shell. A multi-bolt flange (68) may be welded to or otherwise formed along this edge for assembling an annular exhaust diffuser duct (38A-B, 40A/B) in an exhaust section (20) of a gas turbine engine.

Abstract: 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.

Abstract: A method for casting a collar (44, 46) for a heat shield (36) of a strut (32) in a gas turbine exhaust section (20). A casting geometry (60, 70) is defined with extra wall thickness (56, 68) in an area of wall curvature (53, 54), which provides a flow path beyond a final geometry of the collar to facilitate a flow of molten metal in the mold (63, 64). The extra thickness is removed after casting, leaving the collar in its final geometry, which may have uniform wall thickness (T, T2). The extra thickness in the casting geometry may be provided by increased radius (R3) in the wall curvature and/or by casting feed portals (66, 68) that span the wall curvature between a tubular portion (50) and a flange (52) of the collar.

Abstract: Manufacture of an arcuate diffuser shell (38A/38B) assembled from an axially forward portion (38A) and an axially aft portion (38B), the two portions welded to respective sides of an arcuate flange (58A) via two respective pairs of circumferential welds (70A/70B and 72A/72B or 80/84 and 82/86 or 80/88 and 82/90). Each pair of welds comprises first and second welds on opposed surfaces (58, 74) of the shell. The first and second welds compensate each other with respect to welding process shrinkage, eliminating weld warping (68) of the shell. The first and second welds may have equal cross sectional areas or equal masses over a circumferential span of the flange.

Abstract: 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.