Abstract: A flow inducer assembly and a method for cooling turbine blades of a gas turbine engine are presented. The gas turbine engine includes a rotor disk having circumferentially distributed disk grooves and turbine blades. Each turbine blade includes a blade root inserted into blade mounting section of the disk groove. Seal plates are attached to an aft side circumference of the rotor disk. The flow inducer assembly is integrated to each seal plate at a side facing away from the rotor disk. The flow inducer assembly is configured to function as a paddle due to rotation of the rotor disk and the seal plate therewith during operation of the gas turbine engine to drive ambient air as a cooling fluid into the disk cavity and enter inside of the turbine blade from the blade root for cooling the turbine blade.

Abstract: A blade (10) for a turbine engine that includes an internal cooling system (56) formed from at least one cavity (58) positioned within a generally elongated airfoil (12). A squealer tip (36) and at least one densified oxide dispersion strengthened layer (38) extend radially from a radially outer tip cap (70) of the blade (10), the tip cap (70) having a tip cap upper surface (50).

Abstract: The invention relates to a method for starting up a gas turbine engine of a combined cycle power plant. The method includes applying load to the gas turbine engine and increasing the load until a predetermined combustor firing temperature is reached, while keeping the adjustable inlet guide vanes in a start position adapted to reduce the mass flow of air into the compressor; further increasing the load of the gas turbine engine while opening the adjustable inlet guide vanes and keeping the predetermined combustor firing temperature constant until the inlet guide vanes reach an end position adapted to increase the mass flow of air into the compressor; further increasing the load of the gas turbine engine while keeping the adjustable inlet guide vanes in the end position until a predetermined load of the gas turbine engine is reached.

Abstract: A flow inducer assembly and a method for cooling turbine blades of a gas turbine engine are presented. The gas turbine engine includes a rotor disk having circumferentially distributed disk grooves and turbine blades. Each turbine blade includes a blade root inserted into blade mounting section of the disk groove. Seal plates are attached to aft side circumference of the rotor disk. The flow inducer assembly is integrated to each seal plate at a side facing away from the rotor disk. The flow inducer assembly is configured to function as a paddle due to rotation of the rotor disk and the seal plate therewith during operation of the gas turbine engine to drive ambient air as a cooling fluid into the disk cavity and enter inside of the turbine blade from blade root for cooling the turbine blade.

Abstract: The invention relates to a method for starting up a gas turbine engine of a combined cycle power plant. The method includes applying load to the gas turbine engine and increasing the load until a predetermined combustor firing temperature is reached, while keeping the adjustable inlet guide vanes in a start position adapted to reduce the mass flow of air into the compressor; further increasing the load of the gas turbine engine while opening the adjustable inlet guide vanes and keeping the predetermined combustor firing temperature constant until the inlet guide vanes reach an end position adapted to increase the mass flow of air into the compressor; further increasing the load of the gas turbine engine while keeping the adjustable inlet guide vanes in the end position until a predetermined load of the gas turbine engine is reached.

Abstract: A cooling system configured to cool aspects of the turbine engine between a compressor and a turbine assembly is disclosed. In at least one embodiment, the cooling system may include one or more mid-frame cooling channels extending from an inlet through one or more mid-frame torque discs positioned downstream of the compressor and upstream of the turbine assembly. The inlet may be positioned to receive compressor bleed air. The mid-frame cooling channel may be positioned in a radially outer portion of the mid-frame torque disc to provide cooling to outer aspects of the mid-frame torque disc such that conventional, low cost materials may be used to form the mid-frame torque disc rather than high cost materials with capacity to withstand higher temperatures. The cooling fluid routed through the mid-frame cooling channel in the mid-frame torque disc may be exhausted into a cooling system (10) for the downstream turbine assembly.

Abstract: A blade (10) for a turbine engine that includes an internal cooling system (56) formed from at least one cavity (58) positioned within a generally elongated airfoil (12). A squealer tip (36) and at least one densified oxide dispersion strengthened layer (38) extend radially from a radially outer tip cap (70) of the blade (10), the tip cap (70) having a tip cap upper surface (50).

Abstract: A turbine component including a shrouded airfoil with a flow conditioner configured to direct leakage flow and coolant to be aligned with main hot gas flow is provided. The flow conditioner is positioned on a shroud base radially adjacent to the tip of the airfoil and includes a ramped radially outer surface positioned further radially inward than a radially outer surface of the shroud base. The ramped radially outer surface extends from a first edge to a second edge in a direction generally from the suction side to the pressure side of the airfoil, such that the first edge is positioned further radially inward than the second edge. Multiple coolant ejection holes are positioned on the ramped radially outer surface. The coolant ejection holes are connected fluidically to an interior of the airfoil.

Abstract: A gas turbine engine having a rotor centering cooling system for cooling struts within an exhaust diffuser and turbine case to reduce tip rub during hot restarts is disclosed. In particular, the rotor centering cooling system may be positioned within struts in the exhaust diffuser downstream from a turbine assembly for limiting thermal gradients between top and bottom struts to prevent the exhaust bearing body from becoming off-center during steady state operation as a result of the top struts becoming hotter than the bottom struts. The rotor centering cooling system may reduce the temperature at the exhaust diffuser and turbine case, thereby reducing the thermal gradient between the top and bottom struts and top and bottom of the turbine case. As such, the exhaust bearing body remains centered, thereby preventing a tighter blade tip clearance at the top of the turbine assembly than at the bottom of the assembly.

Abstract: A compressor (10) configured for use in a gas turbine engine (12) and having a rotor assembly (14) with a pumping system (16) positioned on a rotor drum (18) to counteract reverse leakage flow at a gap (20) formed between one or more stator vane tips (22) and a radially outer surface (24) of the rotor drum (18). The pumping system (16) may be from pumping components (26) positioned radially inward of one or more stator vane tips (22) to reduce, if not completely eliminate, reverse leakage flow at the stator vane tips (22). In at least one embodiment, the pumping component (26) may be formed from one or more cutouts (28) in the outer surface (24) of the rotor drum (18). In another embodiment, the pumping component (26) may be formed from at least one pumping fin (30) extending from the radially outer surface (24) of the rotor drum (18). In at least one embodiment, rows (32) of pumping components (26) may be aligned with rows (34) of stator vanes (36) within the compressor (10).

Abstract: Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with ridges projecting from the abradable surface, separated by grooves. The ridges have one or both sidewalls inclined against the opposing turbine blade tip rotational direction for redirecting and/or dissipating blade tip gap leakage airflow energy. In some embodiments the ridge tip and/or groove base have inclined profiles for redirecting airflow leakage away from the blade tip gap. In some embodiments, the inclined ridge tip profile provides a progressive wear zone that increases abradable surface area as the inclined ridge is abraded by the rotating blade tip.

Abstract: Turbine and compressor casing abradable components for turbine engines include abradable surfaces with a zonal system of forward (zone A) and rear or aft sections (zone B) surface features. The zone A surface profile comprises an array pattern of non-directional depression dimples, or upwardly projecting dimples, or both, in the abradable surface. The dimpled forward zone A surface features reduce surface solidity in a controlled manner, to help increase abradability during blade tip rubbing incidents, yet they provide sufficient material to resist incoming hot working fluid erosion of the abradable surface. In addition, the dimples provide generic forward section aerodynamic profiling to the abradable surface, compatible with different blade airfoil-camber profiles. The aft zone B surface features comprise an array pattern of ridges and grooves.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines vary localized porosity or abradability through use of holes or dimple depressions of desired polygonal profiles that are formed into the surface of otherwise monolithic abradable surfaces or rib structures. Abradable porosity within a rib is varied locally by changing any one or more of hole/depression depth, diameter, array pitch density, and/or volume. In various embodiments, localized porosity increases and corresponding abradability increases axially from the upstream or forward axial end of the abradable surface to the downstream or aft end of the surface. In this way, the forward axial end of the abradable surface has less porosity to counter hot working gas erosion of the surface, while the more aft portions of the abradable surface accommodate blade cutting and incursion with lower likelihood of blade tip wear.

Abstract: A cooling system configured to cool aspects of the turbine engine between a compressor and a turbine assembly is disclosed. In at least one embodiment, the cooling system may include one or more mid-frame cooling channels extending from an inlet through one or more mid-frame torque discs positioned downstream of the compressor and upstream of the turbine assembly. The inlet may be positioned to receive compressor bleed air. The mid-frame cooling channel may be positioned in a radially outer portion of the mid-frame torque disc to provide cooling to outer aspects of the mid-frame torque disc such that conventional, low cost materials may be used to form the mid-frame torque disc rather than high cost materials with capacity to withstand higher temperatures. The cooling fluid routed through the mid-frame cooling channel in the mid-frame torque disc may be exhausted into a cooling system (10) for the downstream turbine assembly.

Abstract: A shrouded turbine airfoil (10) with a leakage flow conditioner (12) configured to direct leakage flow to be aligned with main hot gas flow is disclosed. The leakage flow conditioner (12) may be positioned on a radially outer surface (18) of an outer shroud base (20) of the outer shroud (22) on a tip (24) of an airfoil (10). The leakage flow conditioner (12) may include a radially outer surface (28) that is positioned further radially inward than the radially outer surface (18) of the outer shroud base (20) creating a radially outward extending wall surface (30) that serves to redirect leakage flow. In at least one embodiment, the radially outward extending wall surface (30) may be aligned with a pressure side (38) of the shrouded turbine airfoil (10) to increase the efficiency of a turbine engine by redirecting leakage flow to be aligned with main hot gas flow to reduce aerodynamic loss upon re-introduction to the main gas flow.

Abstract: A turbine arrangement including a rotor and a stator surrounding the rotor and comprising guide vane segments, each guide vane segment comprising an airfoil and a radially inner vane platform. A seal arrangement includes a static seal inward from the inner vane platforms and having a radially extending face plate, first and second cylindrical seal walls extending from outer and inner ends of the annular face plate, an annular seal plate extending radially from the second cylindrical seal wall, and an angel wing extending between the first cylindrical seal wall and the annular seal plate to define a first annular cavity and a second annular cavity. Circumferentially spaced cut-outs define passages through the annular seal plate between the first and second annular cavities and are aligned with fasteners that attach the annular face plate to a support ring for supporting the inner vane platform.

Abstract: A turbine component (10) including a shrouded airfoil (32) with a flow conditioner (70, 70a, 70b) configured to direct leakage flow and coolant to be aligned with main hot gas flow is provided. The flow conditioner (70, 70a, 70b) is positioned on a shroud base (20) radially adjacent to the tip of the airfoil and includes a ramped radially outer surface (72) positioned further radially inward than a radially outer surface (25) of the shroud base (20). The ramped radially outer surface (72) extends from a first edge (74) to a second edge (76) in a direction generally from the suction side (40) to the pressure side (38) of the airfoil (32), such that the first edge (74) is positioned further radially inward than the second edge (76). Multiple coolant ejection holes (80) are positioned on the ramped radially outer surface (72). The coolant ejection holes (80) are connected fluidically to an interior (81) of the airfoil (32).

Abstract: Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with asymmetric forward and aft ridge surface area density. The forward ridges have greater surface area density than the aft ridges to compensate for greater ridge erosion in the forward zone during engine operation and reduce blade tip wear in the aft zone. Some abradable component embodiments increase forward zone ridge surface area density by incorporating wider ridges than those in the aft zone.

Abstract: A shrouded turbine airfoil (10) with a leakage flow conditioner (12) configured to direct leakage flow to be aligned with main hot gas flow is disclosed. The leakage flow conditioner (12) may be positioned on a radially outer surface (18) of an outer shroud base (20) of the outer shroud (22) on a tip (24) of an airfoil (10). The leakage flow conditioner (12) may include a radially outer surface (28) that is positioned further radially inward than the radially outer surface (18) of the outer shroud base (20) creating a radially outward extending wall surface (30) that serves to redirect leakage flow. In at least one embodiment, the radially outward extending wall surface (30) may be aligned with a pressure side (38) of the shrouded turbine airfoil (10) to increase the efficiency of a turbine engine by redirecting leakage flow to be aligned with main hot gas flow to reduce aerodynamic loss upon re-introduction to the main gas flow.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines vary localized porosity or abradability through use of holes or dimple depressions of desired polygonal profiles that are formed into the surface of otherwise monolithic abradable surfaces or rib structures. Abradable porosity within a rib is varied locally by changing any one or more of hole/depression depth, diameter, array pitch density, and/or volume. In various embodiments, localized porosity increases and corresponding abradability increases axially from the upstream or forward axial end of the abradable surface to the downstream or aft end of the surface. In this way, the forward axial end of the abradable surface has less porosity to counter hot working gas erosion of the surface, while the more aft portions of the abradable surface accommodate blade cutting and incursion with lower likelihood of blade tip wear.