Abstract: A method may include: in a used metal turbomachine blade including a root including a shank, a platform coupled to the shank and an airfoil coupled to the platform, removing the airfoil, leaving a remaining base including the platform, the shank and the root. The method may also form a radially extending opening through the platform into the shank, and insert a ceramic shank nub extending from a ceramic airfoil into the radially extending opening of the remaining base. The ceramic airfoil is fixedly attached to the remaining base. The method allows reuse of the metal shank while providing the lower cooling requirements of a ceramic airfoil.

Abstract: A method may include: in a used metal turbomachine blade including a root including a shank, a platform coupled to the shank and an airfoil coupled to the platform, removing the airfoil, leaving a remaining base including the platform, the shank and the root. The method may also form a radially extending opening through the platform into the shank, and insert a ceramic shank nub extending from a ceramic airfoil into the radially extending opening of the remaining base. The ceramic airfoil is fixedly attached to the remaining base. The method allows reuse of the metal shank while providing the lower cooling requirements of a ceramic airfoil.

Abstract: A method may include: in a used metal turbomachine blade including a root including a shank, a platform coupled to the shank and an airfoil coupled to the platform, removing the airfoil, leaving a remaining base including the platform, the shank and the root. The method may also form a radially extending opening through the platform into the shank, and insert a ceramic shank nub extending from a ceramic airfoil into the radially extending opening of the remaining base. The ceramic airfoil is fixedly attached to the remaining base. The method allows reuse of the metal shank while providing the lower cooling requirements of a ceramic airfoil.

Abstract: A method may include: in a used metal turbomachine blade including a root including a shank, a platform coupled to the shank and an airfoil coupled to the platform, removing the airfoil, leaving a remaining base including the platform, the shank and the root. The method may also form a radially extending opening through the platform into the shank, and insert a ceramic shank nub extending from a ceramic airfoil into the radially extending opening of the remaining base. The ceramic airfoil is fixedly attached to the remaining base. The method allows reuse of the metal shank while providing the lower cooling requirements of a ceramic airfoil.

Abstract: A component configured for impingement cooling includes an inner wall defining a plurality of apertures extending therethrough. Each aperture of the plurality of apertures is configured to emit a cooling fluid therethrough. The component also includes an outer wall that includes an exterior surface, an opposite interior surface, and a thickness defined therebetween. The component further includes a plurality of recesses defined in the outer wall. Each recess of the plurality of recesses extends from a recess first end to an opposite recess second end. The second recess end is defined at the interior surface, and the recess first end is positioned within the outer wall at a depth less than the thickness. Each recess is aligned with a corresponding aperture of the plurality of apertures to receive the cooling fluid therefrom.

Abstract: A component configured for impingement cooling includes an inner wall defining a plurality of apertures extending therethrough. Each aperture of the plurality of apertures is configured to emit a cooling fluid therethrough. The component also includes an outer wall that includes an exterior surface, an opposite interior surface, and a thickness defined therebetween. The component further includes a plurality of recesses defined in the outer wall. Each recess of the plurality of recesses extends from a recess first end to an opposite recess second end. The second recess end is defined at the interior surface, and the recess first end is positioned within the outer wall at a depth less than the thickness. Each recess is aligned with a corresponding aperture of the plurality of apertures to receive the cooling fluid therefrom.

Abstract: One embodiment of the present invention is a turbine nozzle segment for a turbine section of a gas turbine. The turbine nozzle segment includes an inner platform, an outer platform and an airfoil that extends therebetween. The airfoil includes a forward portion and an aft portion that is disposed downstream from the forward portion. The turbine nozzle segment further includes a fuel injection insert that extends between the inner platform and the outer platform downstream from the aft portion of the airfoil. The fuel injection insert includes a fuel circuit that extends within the fuel injection insert, and a plurality of fuel injection ports disposed within the fuel injection insert. The plurality of fuel injection ports provide for fluid communication with the fuel circuit.

Abstract: An apparatus for restricting fluid flow in a turbo-machine is disclosed. The apparatus includes: a first inter-stage sealing member configured to be located between a first rotating airfoil stage and a second rotating airfoil stage attached to a rotor shaft, the first inter-stage sealing member including a segmented structure that includes a plurality of segments forming a circumferential surface, the first inter-stage sealing member including: a first base portion including a securing mechanism configured to at least substantially radially and tangentially secure the inter-stage sealing member to an inter-stage support structure; a first axial retention mechanism; and a first axially extending sealing portion configured to be outwardly radially loaded against at least one of the first rotating airfoil stage and the second rotating airfoil stage.

Abstract: One embodiment of the present invention is a turbine nozzle segment for a turbine section of a gas turbine. The turbine nozzle segment includes an inner platform, an outer platform and an airfoil that extends therebetween. The airfoil includes a forward portion and an aft portion that is disposed downstream from the forward portion. The turbine nozzle segment further includes a fuel injection insert that extends between the inner platform and the outer platform downstream from the aft portion of the airfoil. The fuel injection insert includes a fuel circuit that extends within the fuel injection insert, and a plurality of fuel injection ports disposed within the fuel injection insert. The plurality of fuel injection ports provide for fluid communication with the fuel circuit.

Abstract: Transition ducts, hot gas path assemblies, and turbomachines are provided. A hot gas path assembly includes an outer support ring, an inner support ring, and a transition duct coupled to the outer support ring and the inner support ring. The transition duct includes a conduit defining a passage extending between an inlet and an outlet. The inlet and the outlet are generally aligned along a longitudinal axis. The transition duct further includes an airfoil disposed within the conduit and configured to alter a hot gas flow direction.

Abstract: An apparatus for cooling an airfoil is provided. The airfoil includes an upper airfoil section, a lower airfoil section, at least one cooling passage, and a transition section. The at least one cooling passage is defined at least partially within the lower airfoil section. The at least one cooling passage is configured to flow a cooling medium therethrough, cooling at least a portion of the airfoil. The transition section is disposed between the upper airfoil section and the lower airfoil section and has an outer surface. The outer surface defines at least one cooling hole. The at least one cooling hole is fluidly connected to the at least one cooling passage. At least a portion of the cooling medium is exhausted through the at least one cooling hole.

Abstract: A device for reducing secondary airflow in a gas turbine is disclosed. The device includes an inter-stage sealing member located between a plurality of first turbine buckets attached to a first rotor disk, and a plurality of second turbine buckets attached to a second rotor disk. The first rotor disk and the second rotor disk are rotatable about a central axis. The inter-stage sealing member is configured to be attached in a fixed position relative to the first rotor disk and the second rotor disk, and to contact the plurality of first buckets and the plurality of second buckets in a sealing engagement.

Abstract: An axial retention device for a turbine is disclosed. The axial retention device includes a latch associated with a mating surface of one of a turbine component and a support structure. The latch has an outward bias and includes an axial load surface. The axial retention device further includes a pocket defined in a mating surface of the other of the turbine component and the support structure. The pocket is configured to accept the latch therein and includes a mating axial load surface. Engagement of the latch and the pocket allows the axial load surface and the mating axial load surface to interact, preventing axial movement of the turbine component with respect to the support structure in at least one direction.

Abstract: An apparatus for restricting fluid flow in a turbo-machine is disclosed. The apparatus includes: a first inter-stage sealing member configured to be located between a first rotating airfoil stage and a second rotating airfoil stage attached to a rotor shaft, the first inter-stage sealing member including a segmented structure that includes a plurality of segments forming a circumferential surface, the first inter-stage sealing member including: a first base portion including a securing mechanism configured to at least substantially radially and tangentially secure the inter-stage sealing member to an inter-stage support structure; a first axial retention mechanism; and a first axially extending sealing portion configured to be outwardly radially loaded against at least one of the first rotating airfoil stage and the second rotating airfoil stage.

Abstract: A seal member for a hot gas path component in a gas turbine system is provided. The seal member includes a first seal pin, a second seal pin spaced from the first seal pin, and at least one cross-member. The at least one cross-member has a first end and a second end, and may be joined at the first end to the first seal pin and at the second end to the second seal pin. The at least one cross-member is configured to force the first and second seal pins apart when the seal member is subjected to a centrifugal force.

Abstract: An apparatus for cooling an airfoil is provided. The airfoil includes an upper airfoil section, a lower airfoil section, at least one cooling passage, and a transition section. The at least one cooling passage is defined at least partially within the lower airfoil section. The at least one cooling passage is configured to flow a cooling medium therethrough, cooling at least a portion of the airfoil. The transition section is disposed between the upper airfoil section and the lower airfoil section and has an outer surface. The outer surface defines at least one cooling hole. The at least one cooling hole is fluidly connected to the at least one cooling passage. At least a portion of the cooling medium is exhausted through the at least one cooling hole.

Abstract: A device for reducing secondary airflow in a gas turbine is disclosed. The device includes an inter-stage sealing member located between a plurality of first turbine buckets attached to a first rotor disk, and a plurality of second turbine buckets attached to a second rotor disk. The first rotor disk and the second rotor disk are rotatable about a central axis. The inter-stage sealing member is configured to be attached in a fixed position relative to the first rotor disk and the second rotor disk, and to contact the plurality of first buckets and the plurality of second buckets in a sealing engagement.

Abstract: Embodiments of the invention can provide systems and methods for using a combustion dynamics tuning algorithm with a multi-can combustor. According to one embodiment of the invention, a method for controlling a gas turbine engine with an engine model can be implemented for an engine comprising multiple cans. The method can include obtaining operating frequency information associated with multiple cans of the engine. In addition, the method can include determining variation between operating frequency information of at least two cans. Furthermore, the method can include determining a median value based at least in part on the variation. Moreover, the method can include inputting the median value to an engine model, wherein based at least in part on the median value, the engine model determines an engine control action. In addition, the method can include outputting a control command to implement the engine control action.

Abstract: Embodiments of the invention can provide systems and methods for using a combustion dynamics tuning algorithm with a multi-can combustor. According to one embodiment of the invention, a method for controlling a gas turbine engine with an engine model can be implemented for an engine comprising multiple cans. The method can include obtaining operating frequency information associated with multiple cans of the engine. In addition, the method can include determining variation between operating frequency information of at least two cans. Furthermore, the method can include determining a median value based at least in part on the variation. Moreover, the method can include inputting the median value to an engine model, wherein based at least in part on the median value, the engine model determines an engine control action. In addition, the method can include outputting a control command to implement the engine control action.

Abstract: A damper and seal system for a stage of a turbine that includes inner shrouds disposed circumferentially of a hot gas path through the turbine stage and shroud body(s) for supporting the inner shroud(s). A damper block engages a backside surface of the inner shroud and a damping mechanism is carried by the shroud body and connected to the damper block for applying a load to the damper block and inner shroud through the engagement of the block with the backside surface of the inner shroud, thereby damping vibratory movement of the inner shroud. The seal system includes at least one primary, integral seal and at least one secondary, non-integral seal to limit axial and radial hot gas leakage through the stage.