Abstract: A turbine blade has an attachment root and an airfoil. A cooling passageway system has a plurality of trunks extending from respective inlets along the root inner diameter end from a leading trunk near a first axial end to a trailing trunk near a second axial end; and a plurality of outlets along the airfoil including trailing edge outlets fed by the trailing trunk. Viewed normal to a root end-to-end centerplane: the trailing trunk has a turn passing forward and then rearward; an outside of the turn protrudes forward; and the outside of the turn has a tighter curvature than an inside of the turn.

Abstract: Embodiments described herein relate generally to a portable biometric monitoring device having a central pod, sensor electrodes, and a wearable band. The central pod can be removably coupled to the wearable band. The sensor electrodes can transfer data to an circuit on a printed circuit board (PCB) in the central pod. The circuit can be housed in the central pod and can be configured to process data transferred from the sensor electrodes. The central pod can be electrically coupled to the sensor electrodes via one or more conductive wires while the wearable band is coupled to the central pod. In some embodiments, the central pod can be electronically isolated from the sensor electrodes when the wearable band is not coupled to the central pod. In some embodiments, the one or more conductive wires are substantially encased in the wearable band.

Abstract: A turbine blade for a gas turbine engine. The turbine blade having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.

Abstract: A turbine blade has an attachment root and an airfoil. A cooling passageway system has a plurality of trunks extending from respective inlets along the root inner diameter end from a leading trunk near a first axial end to a trailing trunk near a second axial end; and a plurality of outlets along the airfoil including trailing edge outlets fed by the trailing trunk. Viewed normal to a root end-to-end centerplane: the trailing trunk has a turn passing forward and then rearward; an outside of the turn protrudes forward; and the outside of the turn has a tighter curvature than an inside of the turn.

Abstract: A turbine blade for a gas turbine engine. The turbine blade having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.

Abstract: A turbine blade for a gas turbine engine. The turbine blade having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.

Abstract: A casting core may comprise a tip flag material having a forward cavity and a first spear cavity disposed aft of the forward cavity. A trailing edge discharge material may be separated from the tip flag material and include a first row of cavities. The first row of cavities may comprise a first racetrack cavity. A second row of cavities may be disposed aft of the first row of cavities and include a second racetrack cavity. A third row of cavities may be disposed aft of the second row of cavities and include a circular cavity. A fourth row of cavities may be disposed aft of the third row of cavities and include a second spear cavity.

Abstract: A component for a gas turbine engine includes a body portion that extends between a leading edge and a trailing edge of the component. The trailing edge includes a flared region and a non-flared region. At least one discharge slot is disposed at least partially within the flared region of the component.

Abstract: A component for a gas turbine engine includes a body portion that extends between a leading edge and a trailing edge of the component. The trailing edge includes a flared region and a non-flared region. At least one discharge slot is disposed at least partially within the flared region of the component.

Abstract: A component for a gas turbine engine includes an exterior surface that provides pressure and suction sides. A cooling passage in the component includes a serpentine passageway that has first and second passes respectively configured to provide fluid flow in opposite directions from one another. The first pass includes first and second portions nested relative to one another and overlapping in a thickness direction. The first and second portions are adjacent to one another by sharing a common wall. The first portion is provided on the suction side. The second portion is provided on the pressure side.

Abstract: A casting core and/or an airfoil may comprise a tip flag cavity having a forward pedestal and a first spear pedestal disposed aft of the forward pedestal. A trailing edge discharge cavity may be separated from the tip flag cavity and include a first row of pedestals. The first row of pedestals may comprise a first racetrack pedestal. A second row of pedestals may be disposed aft of the first row of pedestals and include a second racetrack pedestal. A third row of pedestals may be disposed aft of the second row of pedestals and include a circular pedestal. A fourth row of pedestals may be disposed aft of the third row of pedestals and include a second spear pedestal.

Abstract: A casting core and/or an airfoil may comprise a tip flag cavity having a forward pedestal and a first spear pedestal disposed aft of the forward pedestal. A trailing edge discharge cavity may be separated from the tip flag cavity and include a first row of pedestals. The first row of pedestals may comprise a first racetrack pedestal. A second row of pedestals may be disposed aft of the first row of pedestals and include a second racetrack pedestal. A third row of pedestals may be disposed aft of the second row of pedestals and include a circular pedestal. A fourth row of pedestals may be disposed aft of the third row of pedestals and include a second spear pedestal.

Abstract: An airfoil of a gas turbine engine is provided including a leading edge extending in a radial direction, a tip extending in an axial direction from the leading edge, a first rib extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity within the airfoil, a second rib, the second rib and the first rib defining a serpentine cavity therein, a third rib extending axially within the tip, a flag tip cavity defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity, and a bypass aperture formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity.

Abstract: A turbine blade vibration damper is described. The damper has an axially-extending main body defining an inner surface and opposed damping surface. The main body also has an edge spanning from the inner surface to the damping surface and slotted apertures extending through the main body. The slotted apertures extend between the inner surface and damping surface. Adjacent pairs of slotted apertures have elongated shapes defining longitudinal axes that intersect at a point off the surface of the damper.

Abstract: A casting core and/or an airfoil may comprise a tip flag cavity having a forward pedestal and a first spear pedestal disposed aft of the forward pedestal. A trailing edge discharge cavity may be separated from the tip flag cavity and include a first row of pedestals. The first row of pedestals may comprise a first racetrack pedestal. A second row of pedestals may be disposed aft of the first row of pedestals and include a second racetrack pedestal. A third row of pedestals may be disposed aft of the second row of pedestals and include a circular pedestal. A fourth row of pedestals may be disposed aft of the third row of pedestals and include a second spear pedestal.

Abstract: An airfoil of a gas turbine engine is provided including a leading edge extending in a radial direction, a tip extending in an axial direction from the leading edge, a first rib extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity within the airfoil, a second rib, the second rib and the first rib defining a serpentine cavity therein, a third rib extending axially within the tip, a flag tip cavity defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity, and a bypass aperture formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity.

Abstract: A turbine blade vibration damper is described. The damper has an axially-extending main body defining an inner surface and opposed damping surface. The main body also has an edge spanning from the inner surface to the damping surface and slotted apertures extending through the main body. The slotted apertures extend between the inner surface and damping surface. Adjacent pairs of slotted apertures have elongated shapes defining longitudinal axes that intersect at a point off the surface of the damper.

Abstract: A component for a gas turbine engine includes an exterior surface that provides pressure and suction sides. A cooling passage in the component includes a serpentine passageway that has first and second passes respectively configured to provide fluid flow in opposite directions from one another. The first pass includes first and second portions nested relative to one another and overlapping in a thickness direction. The first and second portions are adjacent to one another by sharing a common wall. The first portion is provided on the suction side. The second portion is provided on the pressure side.