Abstract: Various embodiments of the invention include turbine nozzles and systems employing such nozzles. Various particular embodiments include a turbine nozzle having: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and at least one endwall connected with the airfoil along the suction side, pressure side, trailing edge and the leading edge, the at least one endwall including a non-axisymmetric contour proximate a junction between the endwall and the leading edge of the airfoil.

Abstract: Embodiments of the present disclosure provide a cooling structure for a stationary blade, including: an endwall coupled to a radial end of an airfoil; a chamber positioned within the endwall and radially displaced from a radially outer end of the trailing edge of the airfoil, wherein the chamber includes a pair of opposing chamber walls, one of the pair of opposing chamber walls being positioned proximal to the pressure side surface of the airfoil and the other of the pair of opposing chamber walls being positioned proximal to the suction side surface and the trailing edge of the airfoil, and wherein the cooling fluid in the chamber is in thermal communication with least a portion of the endwall positioned proximal to the pressure side surface and the trailing edge of the airfoil; and a plurality of thermally conductive fixtures positioned within the chamber.

Abstract: Various embodiments of the invention include turbine nozzles and systems employing such nozzles. Various particular embodiments include a turbine nozzle having: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and at least one endwall connected with the airfoil along the suction side, pressure side, trailing edge and the leading edge, the at least one endwall including a non-axisymmetric contour proximate a junction between the endwall and the leading edge of the airfoil.

Abstract: Embodiments of the present disclosure provide a cooling structure for a stationary blade, including: an endwall coupled to a radial end of an airfoil; a chamber positioned within the endwall and radially displaced from a radially outer end of the trailing edge of the airfoil, wherein the chamber includes a pair of opposing chamber walls, one of the pair of opposing chamber walls being positioned proximal to the pressure side surface of the airfoil and the other of the pair of opposing chamber walls being positioned proximal to the suction side surface and the trailing edge of the airfoil, and wherein the cooling fluid in the chamber is in thermal communication with least a portion of the endwall positioned proximal to the pressure side surface and the trailing edge of the airfoil; and a plurality of thermally conductive fixtures positioned within the chamber.

Abstract: A turbomachine component includes a body having a first end that extends to a second end. One of the first and second ends includes a mounting element, and a mounting component. A cover plate is arranged at the one of the first and second ends to establish an interface region. The cover plate includes a mounting member configured to align with the mounting element, and a mounting portion configured to align with the mounting element. A fastener member is configured and disposed to cooperate with the mounting element and the mounting member to constrain the cover plate to the body along at least two axes with the interface region being devoid of a metallurgical bond.

Abstract: The present application provides an impingement cooling system for use with a contoured surface. The impingement cooling system may include an impingement plenum and an impingement plate with a linear shape facing the contoured surface. The impingement surface may include a number of projected area thereon with a number of impingement holes having varying sizes and varying spacings.

Abstract: The present application provides a seal carrier for use about a number of flow orifices of a platform of a turbine nozzle. The seal carrier may include an inner surface facing the platform with the inner surface having a number of slots therein aligning with the flow orifices of the platform and an opposed outer surface with a seal positioned about the outer surface.

Abstract: The present application provides a compartmentalized cooling system for providing a cooling flow in a turbine with a flow of combustion gases therein. The compartmentalized cooling system may include a turbine nozzle and a cooling baffle. The turbine nozzle may include an airfoil insert and a nozzle outer sidewall. The cooling baffle may include a high pressure cooling passage in communication with the airfoil insert in a first circuit and an impingement plate positioned about the nozzle outer sidewall in a second circuit.

Abstract: The present application provides an inner nozzle platform. The inner nozzle platform may include a platform cavity, an impingement plenum positioned within the platform cavity, a retention plate positioned on a first side of the impingement plenum, and a compliant seal positioned on a second side of the impingement plenum.

Abstract: The present application provides a turbine nozzle including an airfoil shape. The airfoil shape may have a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table 1. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances in inches by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil in inches. The X and Y values are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The airfoil profile sections at Z distances may be joined smoothly with one another to form a complete airfoil shape.

Abstract: The present application provides a method of installing an impingement cooling assembly in an inner platform of an airfoil of a turbine nozzle. The method may include the steps of positioning an insert within a cavity of the airfoil, positioning a core exit cover about an opening of the cavity, positioning an impingement plenum within a platform cavity, inserting an unfixed spoolie through an assembly port of the impingement plenum and into an airflow cavity of the insert, and closing the assembly port.

Abstract: The present application provides an inner nozzle platform. The inner nozzle platform may include a platform cavity, an impingement plenum positioned within the platform cavity, a retention plate positioned on a first side of the impingement plenum, and a compliant seal positioned on a second side of the impingement plenum.

Abstract: The present application provides a method of installing an impingement cooling assembly in an inner platform of an airfoil of a turbine nozzle. The method may include the steps of positioning an insert within a cavity of the airfoil, positioning a core exit cover about an opening of the cavity, positioning an impingement plenum within a platform cavity, inserting an unfixed spoolie through an assembly port of the impingement plenum and into an airflow cavity of the insert, and closing the assembly port.

Abstract: A turbomachine component includes a body having a first end that extends to a second end. One of the first and second ends includes a mounting element, and a mounting component. A cover plate is arranged at the one of the first and second ends to establish an interface region. The cover plate includes a mounting member configured to align with the mounting element, and a mounting portion configured to align with the mounting element. A fastener member is configured and disposed to cooperate with the mounting element and the mounting member to constrain the cover plate to the body along at least two axes with the interface region being devoid of a metallurgical bond.

Abstract: The present application provides an impingement cooling system for use with a contoured surface. The impingement cooling system may include an impingement plenum and an impingement plate with a linear shape facing the contoured surface. The impingement surface may include a number of projected area thereon with a number of impingement holes having varying sizes and varying spacings.

Abstract: A device for sealing a gas path in a turbine includes a first shroud segment and a slot in a surface of the first shroud segment. A barrier extends inside the slot, and a bypass channel in the slot provides fluid communication between the barrier and the slot to the gas path in the turbine. A method for sealing a gas path in a turbine includes placing a barrier between a first slot in a first shroud segment and a second slot in a second shroud segment and flowing a fluid between the barrier and the first slot to the gas path in the turbine, wherein the fluid flows through a first bypass channel in the first slot.

Abstract: The present application provides a seal carrier for use about a number of flow orifices of a platform of a turbine nozzle. The seal carrier may include an inner surface facing the platform with the inner surface having a number of slots therein aligning with the flow orifices of the platform and an opposed outer surface with a seal positioned about the outer surface.

Abstract: The present application provides a compartmentalized cooling system for providing a cooling flow in a turbine with a flow of combustion gases therein. The compartmentalized cooling system may include a turbine nozzle and a cooling baffle. The turbine nozzle may include an airfoil insert and a nozzle outer sidewall. The cooling baffle may include a high pressure cooling passage in communication with the airfoil insert in a first circuit and an impingement plate positioned about the nozzle outer sidewall in a second circuit.

Abstract: The present application provides a turbine nozzle including an airfoil shape. The airfoil shape may have a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table 1. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances in inches by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil in inches. The X and Y values are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The airfoil profile sections at Z distances may be joined smoothly with one another to form a complete airfoil shape.