Abstract: A turbine airfoil including a central cavity defined by an outer wall including pressure and suction sides extending between and joined at leading and trailing edges, and a chordal axis extends generally centrally between the pressure and suction sides. Rib structures located in the central cavity define radial central channels extending across the chordal axis. Radial near wall passages are defined between the rib structures and each of the pressure and suction sides of the outer wall. The radial near wall passages are each open to an adjacent central channel along a radial extent of both the near wall passages and the adjacent central channel to define a radial flow pass associated with each central channel. The flow passes are connected in series to form a serpentine cooling path extending in the direction of the chordal axis.

Abstract: A turbine airfoil cooling system including a low pressure cooling system and a high pressure cooling system for a turbine airfoil of a gas turbine engine is disclosed. In at least one embodiment, the low pressure cooling system may be an ambient air cooling system, and the high pressure cooling system may be a compressor bleed air cooling system. In at least one embodiment, the compressor bleed air cooling system in communication with a high pressure subsystem that may be a snubber cooling system positioned within a snubber. A delivery system including a movable air supply tube may be used to separate the low and high pressure cooling subsystems. The delivery system may enable high pressure cooling air to be passed to the snubber cooling system separate from low pressure cooling fluid supplied by the low pressure cooling system to other portions of the turbine airfoil cooling system.

Abstract: An airfoil attachment system (10) for a modular turbine vane (12) of a gas turbine engine (14) including an outer attachment system (10) with forward and aft radially extending axial hooks (20, 22) configured to be coupled directly to a vane carrier (24) to increase structural integrity of the modular vane (12) is disclosed. The airfoil attachment system (10) may also include one or more midshroud outer supports (26) positioned between the forward and aft radially extending axial hooks (20, 22) to reduce circumferential rocking movement of the airfoil vane back and forth between the suction and pressure sides (28, 30) of the vane (12). The modular turbine airfoil vane (12) may be positioned between adjacent shrouds (32, 34) forming first and second joints (88, 102). A first sealing system (104) may be placed at the first joint (88), and a second sealing system (110) may be placed at the second joint (102) to limit hot gas ingestion.

Abstract: A turbine airfoil including a central cavity defined by an outer wall including pressure and suction sides extending between and joined at leading and trailing edges, and a chordal axis extends generally centrally between the pressure and suction sides. Rib structures located in the central cavity define radial central channels extending across the chordal axis. Radial near wall passages are defined between the rib structures and each of the pressure and suction sides of the outer wall. The radial near wall passages are each open to an adjacent central channel along a radial extent of both the near wall passages and the adjacent central channel to define a radial flow pass associated with each central channel. The flow passes are connected in series to form a serpentine cooling path extending in the direction of the chordal axis.

Abstract: A turbine airfoil usable in a turbine engine and having at least one snubber with a snubber cooling system positioned therein and in communication with an airfoil cooling system is disclosed. The snubber may extend from the outer housing of the airfoil toward an adjacent turbine airfoil positioned within a row of airfoils. The snubber cooling system may include an inner cooling channel separated from an outer cooling channel by an inner wall. The inner wall may include a plurality of impingement cooling orifices that direct impingement fluid against an outer wall defining the outer cooling channel. In one embodiment, the cooling fluids may be exhausted from the snubber, and in another embodiment, the cooling fluids may be returned to the airfoil cooling system. Flow guides may be positioned in the outer cooling channel, which may reduce cross-flow by the impingement orifices, thereby increasing effectiveness.

Abstract: A turbine airfoil cooling system including a low pressure cooling system and a high pressure cooling system for a turbine airfoil of a gas turbine engine is disclosed. In at least one embodiment, the low pressure cooling system may be an ambient air cooling system, and the high pressure cooling system may be a compressor bleed air cooling system. In at least one embodiment, the compressor bleed air cooling system in communication with a high pressure subsystem that may be a snubber cooling system positioned within a snubber. A delivery system including a movable air supply tube may be used to separate the low and high pressure cooling subsystems. The delivery system may enable high pressure cooling air to be passed to the snubber cooling system separate from low pressure cooling fluid supplied by the low pressure cooling system to other portions of the turbine airfoil cooling system.

Abstract: A turbine airfoil usable in a turbine engine and having at least one snubber with a snubber cooling system positioned therein and in communication with an airfoil cooling system is disclosed. The snubber may extend from the outer housing of the airfoil toward an adjacent turbine airfoil positioned within a row of airfoils. The snubber cooling system may include an inner cooling channel separated from an outer cooling channel by an inner wall. The inner wall may include a plurality of impingement cooling orifices that direct impingement fluid against an outer wall defining the outer cooling channel. In one embodiment, the cooling fluids may be exhausted from the snubber, and in another embodiment, the cooling fluids may be returned to the airfoil cooling system. Flow guides may be positioned in the outer cooling channel, which may reduce cross-flow by the impingement orifices, thereby increasing effectiveness.

Abstract: A turbine blade assembly includes a turbine blade having a pressure sidewall and an opposed suction sidewall and a first snubber assembly associated with one of the pressure sidewall and the suction sidewall. The first snubber assembly includes a first base portion extending outwardly from the one of the pressure sidewall and the suction sidewall, and a first snubber portion. The first base portion is integrally cast with the turbine blade and includes first connection structure. The first snubber portion is bicast onto the first base portion and includes second connection structure that interacts with the first connection structure to substantially prevent separational movement between the first base portion and the first snubber portion.