Abstract: A ring segment for a gas turbine engine includes an outer panel defining a structural body for the ring segment. An outer side of an inner panel is attached to an inner side of the outer panel at an interface, and an inner side of the inner panel defines a portion of a hot gas path through the gas turbine engine. An outer side of the outer panel, opposite from the interface, is in communication with a source of cooling air. A plurality of impingement holes extend through the outer panel from the outer side to the inner side of the outer panel for directing impingement air to the outer side of the inner panel. The outer and inner panels define a plurality of flow channels at the interface for effecting convective cooling of the outer panel along the flow channels between the outer and inner panels.

Abstract: A gas turbine includes a casing, a first diaphragm assembly, a second diaphragm assembly, and first energy absorbing apparatus. The casing has a radially outer surface and a radially inner surface including an annular slot extending circumferentially therein. The first diaphragm assembly includes a first inner structure, a first outer structure and a plurality of airfoils extending between the first inner and outer structures. The second diaphragm assembly includes a second inner structure, a second outer structure and a plurality of airfoils extending between the second inner and outer structures. The first energy absorbing apparatus engages a first end portion of the first outer structure so as to absorb at least portions of unsteady aerodynamic loads and steady rotational loads generated by the first diaphragm assembly.

Abstract: A ring segment for a gas turbine engine includes an outer panel defining a structural body for the ring segment. An outer side of an inner panel is attached to an inner side of the outer panel at an interface, and an inner side of the inner panel defines a portion of a hot gas path through the gas turbine engine. An outer side of the outer panel, opposite from the interface, is in communication with a source of cooling air. A plurality of impingement holes extend through the outer panel from the outer side to the inner side of the outer panel for directing impingement air to the outer side of the inner panel. The outer and inner panels define a plurality of flow channels at the interface for effecting convective cooling of the outer panel along the flow channels between the outer and inner panels.

Abstract: A gas turbine includes a casing, a first diaphragm assembly, a second diaphragm assembly, and first energy absorbing apparatus. The casing has a radially outer surface and a radially inner surface including an annular slot extending circumferentially therein. The first diaphragm assembly includes a first inner structure, a first outer structure and a plurality of airfoils extending between the first inner and outer structures. The second diaphragm assembly includes a second inner structure, a second outer structure and a plurality of airfoils extending between the second inner and outer structures. The first energy absorbing apparatus engages a first end portion of the first outer structure so as to absorb at least portions of unsteady aerodynamic loads and steady rotational loads generated by the first diaphragm assembly.

Abstract: A turbine airfoil including a structural support system formed from chordal ribs within an internal cooling system in the airfoil. The chordal ribs forming the structural support system may be positioned in a cooling fluid supply channel in the hollow airfoil that is positioned proximate to a trailing edge cooling channel. The trailing edge cooling channel may include a plurality of pin fins forming a pin fin bank extending between inner surfaces of the pressure side wall and suction side wall. The chordal ribs may protrude from an inner surface of the outer wall of the hollow airfoil and extend in a general chordwise direction from a mid-chord region of the hollow airfoil to the trailing edge and into the pin fin bank. The structural support system reduces stresses proximate to the pin fins in the pin fin bank closest to the mid-chord region of the airfoil.

Abstract: A turbine airfoil having enhanced cooling capabilities. The turbine vane may be configured such that when a generally elongated airfoil of the turbine vane is attached to a turbine engine, a longitudinal axis of the generally elongated airfoil may be positioned nonparallel relative to a radial axis of the turbine engine in which the turbine vane is mounted. In this position, cooling orifices may be positioned in a region that is typically a dead zone in a conventional turbine vane where no cooling occurs. In one embodiment, a plurality of cooling orifices in an inner shroud of the turbine vane may be positioned between an outer edge of the inner shroud in closest proximity to a suction side of the airfoil near a leading edge of the airfoil and an intersection between the suction side and the inner shroud.

Abstract: A crack-resistant vane segment assembly member for an industrial turbine engine is disclosed. The vane segment includes at least one internally-cooled body portion and inner and outer shroud portions. The body portions include end regions characterized by blending and transition zones that ensure heat is transferred at a non-crack-inducing rate from the shroud portions to the body portions. The vane segment assembly also includes a cooling arrangement that reduces the impact of thermal gradient-induced stresses acting at the interface between the body portions and shroud portions. The cooling arrangement cooperates with the transition and blending zones to produce synergistically-enhanced crack resistance properties.

Abstract: A crack-resistant vane segment assembly member for an industrial turbine engine is disclosed. The vane segment includes at least one internally-cooled body portion and inner and outer shroud portions. The body portions include end regions characterized by blending and transition zones that ensure heat is transferred at a non-crack-inducing rate from the shroud portions to the body portions. The vane segment assembly also includes a cooling arrangement that reduces the impact of thermal gradient-induced stresses acting at the interface between the body portions and shroud portions. The cooling arrangement cooperates with the transition and blending zones to produce synergistically-enhanced crack resistance properties.