Abstract: An assembly for a gas turbine engine includes a seal and a flow diverter. The flow diverter is disposed adjacent the seal to direct a secondary gas flow that passes across the seal away from a rotor cavity such that the secondary gas flow travels back toward a main gas flow path of the gas turbine engine.

Abstract: A seal strip (60A, 60B, 60C, 60F, 60G, 60H, 60J) with an imperforate width-spanning portion (64) first and second rounded edges (65, 66) and one or more strip-thickening elements (67, 68, 74, 76, 82, 84, 90, 91) between the rounded edges. The strip-thickening elements may have transverse slots (80, 86, 88, 93L, 93R, 72, 78) for increased flexibility of the strip. The strip-thickening elements may also have perforations (92) or gaps (69) to admit coolant and/or to reduce weight. Folded embodiments (60A, 60B) may have dimples (70) on the width-spanning portion (64) to limit bending of the folded portions (67, 68, 74, 76). The seal may be slidably mounted in opposed slots (58A, 58B) in respective adjacent turbine components (54A, 54B), filling a width (W) of the slots, and a side of the seal may be cooled by compressed air (48).

Abstract: Respective seals (54, 78) for the upper and lower spans (48A, 48B) of an exit frame (48) of a turbine combustion system transition piece (28). Each seal has a first strip (55, 79) and a second strip (66, 88) of a sealing material. The two strips of each seal are attached together along a common edge. The second strip is flexible, generally parallel to the first strip, and has a bead (72, 90) along its free edge. This forms a spring clamp that clamps a rail (68, 86) of the exit frame between the bead and the first strip of each seal. A tab extends axially aft from the first strip of each seal for insertion into a circumferential slot (58, 82) in a turbine inlet support structure (52, 76), thus sealing the transition piece (46) to the turbine inlet for efficient turbine operation.

Abstract: A seal strip (54) with a central relatively thin portion (68) and first and second thicker side portions (70, 72) that may be wedge-shaped adjacent the central portion. Each side portion may be formed of a linear array of base-in prisms (56), where each prism includes a base adjacent and normal to the central portion, and a thickness tapering distally toward an adjacent edge of the seal strip. The base-in prisms of each side portion may be separated by transverse slots (55) along the length of the strip. The transverse slots of the first side portion may be unaligned with the transverse slots of the second side portion along the length of the strip. A retention pin (58) may extend normally from an end of the seal strip. The seal strip may be mounted in tapered slots (49) of a gas turbine transition exit frame (48).

Abstract: A seal strip (60A, 60B, 60C, 60F, 60G, 60H, 60J) with an imperforate width-spanning portion (64) first and second rounded edges (65, 66) and one or more strip-thickening elements (67, 68, 74, 76, 82, 84, 90, 91) between the rounded edges. The strip-thickening elements may have transverse slots (80, 86, 88, 93, 93R, 72, 78) for increased flexibility of the strip. The strip-thickening elements may also have perforations (92) or gaps (69) to admit coolant and/or to reduce weight. Folded embodiments (60A, 60B) may have dimples (70) on the width-spanning portion (64) to limit bending of the folded portions (67, 68, 74, 76). The seal may be slidably mounted in opposed slots (58A, 58B) in respective adjacent turbine components (54A, 54B), filling a width (W) of the slots, and a side of the seal may be cooled by compressed air (48).

Abstract: Respective seals (54, 78) for the upper and lower spans (48A, 48B) of an exit frame (48) of a turbine combustion system transition piece (28). Each seal has a first strip (55, 79) and a second strip (66, 88) of a sealing material. The two strips of each seal are attached together along a common edge. The second strip is flexible, generally parallel to the first strip, and has a bead (72, 90) along its free edge. This forms a spring clamp that clamps a rail (68, 86) of the exit frame between the bead and the first strip of each seal. A tab extends axially aft from the first strip of each seal for insertion into a circumferential slot (58, 82) in a turbine inlet support structure (52, 76), thus sealing the transition piece (46) to the turbine inlet for efficient turbine operation.

Abstract: A seal strip (54) with a central relatively thin portion (68) and first and second thicker side portions (70, 72) that may be wedge-shaped adjacent the central portion. Each side portion may be formed of a linear array of base-in prisms (56), where each prism includes a base adjacent and normal to the central portion, and a thickness tapering distally toward an adjacent edge of the seal strip. The base-in prisms of each side portion may be separated by transverse slots (55) along the length of the strip. The transverse slots of the first side portion may be unaligned with the transverse slots of the second side portion along the length of the strip. A retention pin (58) may extend normally from an end of the seal strip. The seal strip may be mounted in tapered slots (49) of a gas turbine transition exit frame (48).

Abstract: A triple acting radial seal used as an interstage seal assembly in a gas turbine engine, where the seal assembly includes an interstage seal support extending from a stationary inner shroud of a vane ring, the interstage seal support includes a larger annular radial inward facing groove in which an outer annular floating seal assembly is secured for radial displacement, and the outer annular floating seal assembly includes a smaller annular radial inward facing groove in which an inner annular floating seal assembly is secured also for radial displacement. A compliant seal is secured to the inner annular floating seal assembly. The outer annular floating seal assembly encapsulates the inner annular floating seal assembly which is made from a very low alpha material in order to reduce thermal stress.

Abstract: A pressure vessel test rig for testing a component under high pressure and temperature, the pressure vessel having a flange with an opening so that instrumentation lines can pass from inside the vessel to outside, and a sealing plate with grooves that provide a seal for the instrumentation lines when a cover plate is secured to the flange. A removable instrumentation hold-down tool is attached to the flange and includes a number of spring biased clamps that secure the instrumentation lines to the flange during changes to the component within the vessel.

Abstract: A combustion liner having enhanced heat transfer over at least a portion of the liner is disclosed. The combustion liner comprises, amongst other features, an outer surface wherein at least a region of the outer surface is textured in a substantially uniform pattern such that the surface area exposed to a passing cooling medium is increased, thereby increasing the heat transfer through the combustion liner. Multiple embodiments of textured patterns are also disclosed.

Abstract: A combustion liner having enhanced heat transfer over at least a portion of the liner is disclosed. The combustion liner comprises, amongst other features, an outer surface wherein at least a region of the outer surface is textured in a substantially uniform pattern such that the surface area exposed to a passing cooling medium is increased, thereby increasing the heat transfer through the combustion liner. Multiple embodiments of textured patterns are also disclosed.

Abstract: A gas turbine combustion system having reduced emissions and improved flame stability at multiple load conditions is disclosed. The improved combustion system accomplishes this through complete premixing, a plurality of fuel injector locations, combustor geometry, and precise three dimensional staging between fuel injectors. Axial, radial, and circumferential fuel staging is utilized including fuel injection proximate air swirlers. Furthermore, strong recirculation zones are established proximate the introduction of fuel and air premixture from different stages to the combustion zone. The combination of the strong recirculation zones, efficient premixing, and staged fuel flow thereby provide the opportunity to produce low emissions combustion at various load conditions.

Abstract: A gas turbine combustion system having reduced emissions and improved flame stability at multiple load conditions is disclosed. The improved combustion system accomplishes this through complete premixing, a plurality of fuel injector locations, combustor geometry, and precise three dimensional staging between fuel injectors. Axial, radial, and circumferential fuel staging is utilized including fuel injection proximate air swirlers. Furthermore, strong recirculation zones are established proximate the introduction of fuel and air premixture from different stages to the combustion zone. The combination of the strong recirculation zones, efficient premixing, and staged fuel flow thereby provide the opportunity to produce low emissions combustion at various load conditions.