Abstract: A turbine bucket including an airfoil, the airfoil having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table 1 as contained herein wherein the Z values are non-dimensional values from approximately 0 to 1 convertible to Z distances in inches by multiplying the Z values by a height of the airfoil in inches, and wherein X and Y are distances in inches which, when connected by substantially smooth continuing arcs, define airfoil profile sections at each distance Z, the profile sections at the Z distances being joined substantially smoothly with one another to form a complete airfoil shape.

Abstract: A turbine blade that may include a tip shroud, one or more cooling cavities formed within the tip shroud, and at least one discrete structural element defined within at least one of the cooling cavities. The discrete structural element may include an element that structurally connects a floor of the cooling cavity to a ceiling of the cooling cavity that does not originate from, terminate in, or connect to an interior wall of the cooling cavity or an outer edge of the tip shroud.

Abstract: An airfoil assembly for a turbine engine, that may comprise: an airfoil that includes at least one panel opening formed therein; and an aperture panel that includes at least one cooling aperture; wherein the panel opening and the aperture panel are configured such that, upon installation of the aperture panel within the panel opening, a substantially smooth outer surface of the airfoil is formed and at least one of the cooling apertures of the aperture panel forms an exit passageway from at least one hollow internal cavity within the airfoil.

Abstract: A turbine bucket that may include an airfoil having a tip shroud, the tip shroud having an edge; wherein the edge has a profile substantially in accordance with values of X and Y in a Cartesian coordinate system set forth in Table 1 at points 1-14, wherein X and Y represent distances that may be proportionally scaled by a common multiplier which, once scaled and connected, define the profile of the edge of the tip shroud.

Abstract: A turbine blade that may include a tip shroud, the tip shroud having one or more cooling cavities defined therein, an airfoil, the airfoil having at least one coolant passage defined therein, and a plurality of exit apertures defined through an outer wall of at least one of the cooling cavities that, in use, may allow a pressurized coolant to exit at least one of the cooling cavities. The coolant passage may extend through the airfoil such that the coolant passage is in fluid communication with the cooling cavity. And, the plurality of exit apertures may include at least one round exit aperture and at least one non-round exit aperture.

Abstract: A metered cooling slot disposed in a wall comprising an outer surface that is exposed to a hot gas stream and an inner surface that defines an internal coolant chamber through which a coolant passes, the metered cooling slot comprising: a slot formed within the outer surface elongated in a first direction, the slot comprising a pair of spaced apart, opposing, slot surfaces and a base, the slot surfaces intersecting the outer surface to form a slot outlet opposite the base; and two or more metering apertures formed within the wall, each metering aperture intersecting the inner surface of the wall to form a metering aperture inlet and intersecting one of the pair of slot surfaces to form a metering aperture outlet; wherein: D represents the approximate diameter of at least two of the metering apertures; P represents the approximate distance between the center lines of at least two neighboring metering apertures; and P/D comprises a value within the range of about 4 to 6.

Abstract: A turbine blade that may include a tip shroud that comprises one or more cooling cavities formed therein. At least one of the cooling cavities may include a plurality of ribs and a first interior wall that generally opposes a second interior wall across the cooling cavity. The plurality of ribs may be configured such that some of the ribs originate at the first interior wall of the cooling cavity and extend toward the second interior wall of the cooling cavity and some of the ribs originate at the second interior wall of the cooling cavity and extend toward the first interior wall of the cooling cavity. The ribs that originate at the first interior wall of the cooling cavity and extend toward the second interior wall of the cooling cavity may be oriented and sized so that at least some of the ribs extend across at least a majority of the distance across the cooling cavity.

Abstract: The present application further describes a turbine blade that may include an airfoil, a tip shroud that is positioned at one end of the turbine blade, a coolant passage that extends through the airfoil to the tip shroud, and a continuous cooling cavity that is defined within the tip shroud and that is in fluid communication with the coolant passage. The continuous cooling cavity may form a labyrinth cooling circuit within the tip shroud.

Abstract: A metered cooling slot disposed in a wall comprising an outer surface that is exposed to a hot gas stream and an inner surface that defines an internal coolant chamber through which a coolant passes, the metered cooling slot comprising: a slot formed within the outer surface elongated in a first direction, the slot comprising a pair of spaced apart, opposing, slot surfaces and a base, the slot surfaces intersecting the outer surface to form a slot outlet opposite the base; and two or more metering apertures formed within the wall, each metering aperture intersecting the inner surface of the wall to form a metering aperture inlet and intersecting one of the pair of slot surfaces to form a metering aperture outlet; wherein: D represents the approximate diameter of at least two of the metering apertures; P represents the approximate distance between the center lines of at least two neighboring metering apertures; and P/D comprises a value within the range of about 4 to 6.

Abstract: An airfoil assembly for a turbine engine, that may comprise: an airfoil that includes at least one panel opening formed therein; and an aperture panel that includes at least one cooling aperture; wherein the panel opening and the aperture panel are configured such that, upon installation of the aperture panel within the panel opening, a substantially smooth outer surface of the airfoil is formed and at least one of the cooling apertures of the aperture panel forms an exit passageway from at least one hollow internal cavity within the airfoil.

Abstract: Methods and apparatuses for film cooling of one or more turbine components are provided. A cooling gas flow passage provides a cooling gas to an turbine component with the hot gas path of a turbine. The cooling gas flow passage includes at least one feed aperture operable to receive cooling gas from at least one cooling gas compartment associated with a turbine component. The cooling gas flow passage also includes at least one slot with a converging portion having a first opening and a second opening, where the first opening receives the cooling gas from the feed aperture, and the second opening provides the cooling gas to at least a portion of an outer surface of the turbine component.

Abstract: A turbine bucket including an airfoil, the airfoil having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table 1 as contained herein wherein the Z values are non-dimensional values from approximately 0 to 1 convertible to Z distances in inches by multiplying the Z values by a height of the airfoil in inches, and wherein X and Y are distances in inches which, when connected by substantially smooth continuing arcs, define airfoil profile sections at each distance Z, the profile sections at the Z distances being joined substantially smoothly with one another to form a complete airfoil shape.

Abstract: A turbine bucket that may include an airfoil having a tip shroud, the tip shroud having an edge; wherein the edge has a profile substantially in accordance with values of X and Y in a Cartesian coordinate system set forth in Table 1 at points 1-14, wherein X and Y represent distances that may be proportionally scaled by a common multiplier which, once scaled and connected, define the profile of the edge of the tip shroud.

Abstract: A turbine blade that may include a tip shroud, one or more cooling cavities formed within the tip shroud, and at least one discrete structural element defined within at least one of the cooling cavities. The discrete structural element may include an element that structurally connects a floor of the cooling cavity to a ceiling of the cooling cavity that does not originate from, terminate in, or connect to an interior wall of the cooling cavity or an outer edge of the tip shroud.

Abstract: A turbine blade that may include a tip shroud that comprises one or more cooling cavities formed therein. At least one of the cooling cavities may include a plurality of ribs and a first interior wall that generally opposes a second interior wall across the cooling cavity. The plurality of ribs may be configured such that some of the ribs originate at the first interior wall of the cooling cavity and extend toward the second interior wall of the cooling cavity and some of the ribs originate at the second interior wall of the cooling cavity and extend toward the first interior wall of the cooling cavity. The ribs that originate at the first interior wall of the cooling cavity and extend toward the second interior wall of the cooling cavity may be oriented and sized so that at least some of the ribs extend across at least a majority of the distance across the cooling cavity.

Abstract: The present application further describes a turbine blade that may include an airfoil a tip shroud that is positioned at one end of the turbine blade, a coolant passage that extends through the airfoil to the tip shroud, and a continuous cooling cavity that is defined within the tip shroud and that is in fluid communication with the coolant passage. The continuous cooling cavity may form a labyrinth cooling circuit within the tip shroud.

Abstract: A turbine blade that includes a tip shroud and one or more cooling cavities formed within the tip shroud. At least one of the cooling cavities may include a plurality of ribs and a first interior wall that generally opposes a second interior wall across the cooling cavity. The ribs may be configured such that some of the ribs originate on the first interior wall of the cooling cavity and extend toward the second interior wall of the cooling cavity and some of the ribs originate on the second interior wall of the cooling cavity and extend toward the first interior wall of the cooling cavity. And, the ribs that originate on the first interior wall of the cooling cavity and the ribs that originate on the second interior wall of the cooling cavity may have an alternating arrangement.

Abstract: A turbine blade that may include a tip shroud, the tip shroud having one or more cooling cavities defined therein, an airfoil, the airfoil having at least one coolant passage defined therein, and a plurality of exit apertures defined through an outer wall of at least one of the cooling cavities that, in use, may allow a pressurized coolant to exit at least one of the cooling cavities. The coolant passage may extend through the airfoil such that the coolant passage is in fluid communication with the cooling cavity. And, the plurality of exit apertures may include at least one round exit aperture and at least one non-round exit aperture.

Abstract: A localized directional impingement cooling is used to reduce the metal temperatures on highly stressed regions of the tip shroud.