Abstract: A core structure (10) includes a first core element (16) including a leading edge section (30), a tip section (32), and a turn section (34) joining the leading edge and tip sections (30, 32). The first core element (16) is adapted to be used to form a leading edge cooling circuit (102) in a gas turbine engine airfoil (100). The leading edge cooling circuit (102) includes a cooling fluid passage (104) having a leading edge portion (106) formed by the first core element leading edge section (30), a tip portion (108) formed by the first core element tip section (32), and a turn portion (110) formed by the first core element turn section (34). Each of the leading edge portion (106), the tip portion (108), and the turn portion (110) of the cooling fluid passage (104) are formed concurrently in the airfoil (100) by the first core element (16).

Abstract: An airfoil (10) for a gas turbine engine in which the airfoil (10) includes an internal cooling system (14) with one or more internal cavities (16) having an insert (18) contained within an aft cooling cavity (76) to form nearwall cooling channels having enhanced flow patterns is disclosed. The flow of cooling fluids in the nearwall cooling channels (20) may be controlled via a plurality of cooling fluid flow controllers (22) extending from the outer wall (12) forming the generally hollow elongated airfoil (26). In addition, heat may be extracted in the midchord region (150) via one or more heat dissipating ribs (152) extending partially between an inner surface (144) of the suction side (38) and the insert (18).

Abstract: An airfoil (10) for a gas turbine engine in which the airfoil (10) includes an internal cooling system (14) with one or more internal cavities having an insert (18) contained within an aft cooling cavity (76) to form nearwall cooling channels having enhanced flow patterns is disclosed. The flow of cooling fluids in the nearwall cooling channels may be controlled via a plurality of cooling fluid flow controllers (22) extending from the outer wall (24) forming the generally hollow elongated airfoil (26). The cooling fluid flow controllers (22) may be collected into spanwise extending rows. In at least one embodiment, the cooling fluid flow controllers (22) may be positioned within a pressure side nearwall cooling channel (48) and a suction side nearwall cooling channel (50) that are both in fluid communication with a trailing edge channel (30). The trailing edge channel (30) may also include cooling fluid flow controllers (22) extending between the outer walls (12, 13) forming the pressure and suction sides.

Abstract: An airfoil (10) for a gas turbine engine in which the airfoil (10) includes an internal cooling system (14) with one or more internal cavities (16) having an insert (18) contained therein that forms nearwall cooling channels (20) having enhanced flow patterns is disclosed. The flow of cooling fluids in the nearwall cooling channels (20) may be controlled via a plurality of cooling fluid flow controllers (22) extending from the outer wall (24) forming the generally hollow elongated airfoil (26). The cooling fluid flow controllers (22) may be collected into spanwise extending rows (28), and the internal cooling system (14) may include one or more bypass flow reducers (30) extending from the insert (18) toward the outer wall (24) to direct the cooling fluids through the channels (20) created by the cooling fluid flow controllers (22), thereby increasing the effectiveness of the internal cooling system (14).

Abstract: A core structure (10) includes a first core element (16) including a leading edge section (30), a tip section (32), and a turn section (34) joining the leading edge and tip sections (30, 32). The first core element (16) is adapted to be used to form a leading edge cooling circuit (102) in a gas turbine engine airfoil (100). The leading edge cooling circuit (102) includes a cooling fluid passage (104) having a leading edge portion (106) formed by the first core element leading edge section (30), a tip portion (108) formed by the first core element tip section (32), and a turn portion (110) formed by the first core element turn section (34). Each of the leading edge portion (106), the tip portion (108), and the turn portion (110) of the cooling fluid passage (104) are formed concurrently in the airfoil (100) by the first core element (16).

Abstract: An airfoil (10) for a gas turbine engine in which the airfoil (10) includes an internal cooling system (14) with one or more internal cavities (16) having an insert (18) contained therein that forms nearwall cooling channels (20) having enhanced flow patterns is disclosed. The flow of cooling fluids in the nearwall cooling channels (20) may be controlled via a plurality of cooling fluid flow controllers (22) extending from the outer wall (24) forming the generally hollow elongated airfoil (26). The cooling fluid flow controllers (22) may be collected into spanwise extending rows (28), and the internal cooling system (14) may include one or more bypass flow reducers (30) extending from the insert (18) toward the outer wall (24) to direct the cooling fluids through the channels (20) created by the cooling fluid flow controllers (22), thereby increasing the effectiveness of the internal cooling system (14).

Abstract: An airfoil (10) for a gas turbine engine in which the airfoil (10) includes an internal cooling system (14) with one or more internal cavities having an insert (18) contained within an aft cooling cavity (76) to form nearwall cooling channels having enhanced flow patterns is disclosed. The flow of cooling fluids in the nearwall cooling channels may be controlled via a plurality of cooling fluid flow controllers (22) extending from the outer wall (24) forming the generally hollow elongated airfoil (26). The cooling fluid flow controllers (22) may be collected into spanwise extending rows. In at least one embodiment, the cooling fluid flow controllers (22) may be positioned within a pressure side nearwall cooling channel (48) and a suction side nearwall cooling channel (50) that are both in fluid communication with a trailing edge channel (30). The trailing edge channel (30) may also include cooling fluid flow controllers (22) extending between the outer walls (12, 13) forming the pressure and suction sides.

Abstract: A compressor airfoil tip clearance optimization system for reducing a gap between a tip of a compressor airfoil and a radially adjacent component of a turbine engine is disclosed. The turbine engine may include ID and OD flowpath boundaries configured to minimize compressor airfoil tip clearances during turbine engine operation in cooperation with one or more clearance reduction systems that are configured to move the rotor assembly axially to reduce tip clearance. The configurations of the ID and OD flowpath boundaries enhance the effectiveness of the axial movement of the rotor assembly, which includes movement of the ID flowpath boundary. During operation of the turbine engine, the rotor assembly may be moved axially to increase the efficiency of the turbine engine.

Abstract: A compressor airfoil tip clearance optimization system for reducing a gap between a tip of a compressor airfoil and a radially adjacent component of a turbine engine is disclosed. The turbine engine may include ID and OD flowpath boundaries configured to minimize compressor airfoil tip clearances during turbine engine operation in cooperation with one or more clearance reduction systems that are configured to move the rotor assembly axially to reduce tip clearance. The configurations of the ID and OD flowpath boundaries enhance the effectiveness of the axial movement of the rotor assembly, which includes movement of the ID flowpath boundary. During operation of the turbine engine, the rotor assembly may be moved axially to increase the efficiency of the turbine engine.

Abstract: An inner mounting ring (20) for gas turbine flow path components such as shroud ring segments (24). The inner ring (20) may be mounted to an outer ring (22) on radially slidable mounts (26, 28) that maintain the two rings (20, 22) in coaxial relationship, but allows them to thermally expand at different rates. This allows matching of the radial expansion rate of the inner ring (20) to that of the turbine blade tips (32), thus providing reduced clearance (33) between the turbine blade tips (32) and the inner surface of the shroud ring segments (24) under all engine operating conditions. The inner ring (20) may be made of a material with a lower coefficient of thermal expansion than that of the outer ring (22).

Abstract: An inner mounting ring (20) for gas turbine flow path components such as shroud ring segments (24). The inner ring (20) may be mounted to an outer ring (22) on radially slidable mounts (26, 28) that maintain the two rings (20, 22) in coaxial relationship, but allows them to thermally expand at different rates. This allows matching of the radial expansion rate of the inner ring (20) to that of the turbine blade tips (32), thus providing reduced clearance (33) between the turbine blade tips (32) and the inner surface of the shroud ring segments (24) under all engine operating conditions. The inner ring (20) may be made of a material with a lower coefficient of thermal expansion than that of the outer ring (22).

Abstract: A seal system for an intersection between a turbine stator and a turbine rotor to seal cooling fluids. The seal system may be formed from a seal base extending from the turbine stator, an arm extending radially outward from the turbine rotor and toward the seal base but terminating short of the seal base thereby creating a gap between the seal base and the arm. The seal system may include a honeycomb shaped seal attached to the seal base and extending radially inward from the seal base toward the arm. An outer sealing surface of the seal may be nonparallel with a longitudinal axis about which the turbine rotor rotates thereby enabling the distance of the gap to be reduced with axial movement of the turbine rotor.