Abstract: A bearing support assembly to support one or more bearings on a shaft. The bearing support assembly includes a bearing support frame configured to be coupled to a static frame, a plurality of ribs connected to the bearing support frame, and a bearing support connected to the plurality of ribs and configured to support a bearing of the one or more bearings. The bearing support assembly has a non-axisymmetric stiffness around a circumference of the bearing support assembly.

Abstract: A bearing support assembly to support one or more bearings on a shaft. The bearing support assembly includes a bearing support frame configured to be coupled to a static frame, a plurality of ribs connected to the bearing support frame, and a bearing support connected to the plurality of ribs and configured to support a bearing of the one or more bearings. The bearing support assembly has a non-axisymmetric stiffness around a circumference of the bearing support assembly.

Abstract: A bearing support assembly to support one or more bearings on a shaft. The bearing support assembly includes a bearing support frame configured to be coupled to a static frame, a plurality of ribs connected to the bearing support frame, and a bearing support connected to the plurality of ribs and configured to support a bearing of the one or more bearings. The bearing support assembly has a non-axisymmetric stiffness around a circumference of the bearing support assembly.

Abstract: The present application thus provides a method for reducing stress on a turbine blade wherein each of the turbine blades includes a dovetail with lockwire tab. The method may include the steps of (a) determining a starting line for a backcut relative to a lockwire tab end, (b) determining a cut angle for the backcut, and (c) removing material from the lockwire tab according to the starting line and the cut angle to form the dovetail backcut. The starting line may be positioned about 0.6 inches (about 15.24 millimeters), plus or minus 0.065 inches (about 1.65 millimeters) from the lockwire tab end along the dovetail axis.

Abstract: The present disclosure is directed to a rotor blade for a gas turbine engine. The rotor blade includes a platform having a radially inner surface and a radially outer surface. A connection portion extends radially inwardly from the radially inner surface of the platform. An airfoil extends radially outwardly from the radially outer surface of the platform to an airfoil tip. The airfoil includes a leading edge portion and a trailing edge portion. The platform, the airfoil, and the connection portion collectively define a cooling circuit extending from an inlet defined by the connection portion to one or more outlet passages at least partially defined by the trailing edge portion of the airfoil. At least one of the one or more outlet passages include an entrance and an exit radially inwardly offset from the entrance.

Abstract: A rotating airfoil component of a turbomachine, in which the component has an airfoil aligned in a spanwise direction of the component, a shank, and a platform therebetween oriented transverse to the spanwise direction. The platform has an outer radial surface adjacent the airfoil, and at least one recessed region defined in its outer radial surface. The recessed region extends opposite the spanwise direction from a platform plane that contains portions of the outer radial surface that are upstream and downstream from the recessed region. The recessed region is contiguous with an end wall of the platform and extends therefrom toward the airfoil. The recessed region defines a surface shape whose boundary is contained by the platform plane, and has a profile shape that extends from the end wall toward the airfoil. The recessed region is sized and shaped to increase the stiffness of the platform.

Abstract: A rotating airfoil component of a turbomachine, in which the component has an airfoil aligned in a spanwise direction of the component, a shank, and a platform therebetween oriented transverse to the spanwise direction. The platform has an outer radial surface adjacent the airfoil, and at least one recessed region defined in its outer radial surface. The recessed region extends opposite the spanwise direction from a platform plane that contains portions of the outer radial surface that are upstream and downstream from the recessed region. The recessed region is contiguous with an end wall of the platform and extends therefrom toward the airfoil. The recessed region defines a surface shape whose boundary is contained by the platform plane, and has a profile shape that extends from the end wall toward the airfoil. The recessed region is sized and shaped to increase the stiffness of the platform.

Abstract: According to one aspect of the invention, a turbine assembly includes a first bucket with a first slashface and a second bucket including a recess formed in a second slashface of the second bucket, wherein the second slashface is adjacent to the first slashface when the first bucket is positioned adjacent to the second bucket. The turbine assembly also includes a pin configured to be placed in the recess, wherein the pin is magnetically urged toward the first slashface to reduce fluid flow between the first and second buckets.