The present disclosure generally pertains to gas turbine engines, and is more particularly directed toward nozzle segments including film cooling holes with alternating compound angles.
Gas turbine engines include compressor, combustor, and turbine sections. The turbine section is subject to high temperatures. In particular, the first stages of the turbine section are subject to such high temperatures that the first stages are often cooled with air directed from the compressor and into, inter alia, the nozzle segments and turbine blades.
A portion of the air directed into the nozzle segments may be directed through the walls of the nozzle segment airfoils and along the pressure side surface of the walls to film cool the walls. U.S. Pat. No. 7,377,743 to D. Flodman discloses a turbine nozzle that includes a mid vane mounted between a pair of end vanes in outer and inner bands. The mid vane includes a first pattern of film cooling holes configured to discharge more cooling air than each of the two end vanes having respective second patterns of film cooling holes.
The present disclosure is directed toward overcoming one or more of the problems discovered by the inventors or that is known in the art.
A nozzle segment for a nozzle ring of a gas turbine engine is disclosed. The nozzle segment includes a first endwall, a second endwall, and an airfoil extending between the first endwall and the second endwall. The airfoil includes a leading edge, a trailing edge, a pressure side wall, and a suction side wall. The leading edge extends radially from the first endwall to the second endwall. The trailing edge extends radially from the first endwall to the second endwall axially distal to the leading edge. The pressure side wall extends from the leading edge to the trailing edge. The suction side wall also extends from the leading edge to the trailing edge. The airfoil also includes a plurality of showerhead cooling apertures, a plurality of forward cooling apertures, and a plurality of intermediate cooling apertures. The plurality of showerhead cooling apertures span along the leading edge. The plurality of forward cooling apertures are grouped together proximate the plurality of showerhead cooling apertures. The plurality of intermediate cooling apertures are grouped together in the pressure side wall between the trailing edge and the plurality of forward cooling apertures. The plurality of showerhead cooling apertures, the plurality of forward cooling apertures, and the plurality of intermediate cooling apertures alternate in directionality such that the plurality of showerhead cooling apertures are angled toward the first endwall, the plurality of forward cooling apertures are angled toward the second endwall, and the plurality of intermediate cooling apertures are angled toward the first endwall.
The systems and methods disclosed herein include a nozzle segment for a nozzle ring of a gas turbine engine. In embodiments, the nozzle segment includes an upper endwall, an inner endwall, and one or more airfoils there between. Each airfoil includes spaced apart groups of cooling apertures through the pressure side wall of the airfoil. One group is angled toward the lower endwall and the next group is angled towards the upper endwall in an alternating pattern for subsequent groups of cooling holes. Alternating the directionality of the groups of cooling apertures towards the lower endwall and the upper endwall may reduce the temperatures of the lower endwall and the upper endwall, and may reduce the amount of cooling air needed to effectively cool the nozzle segment.
In addition, the disclosure may generally reference a center axis 95 of rotation of the gas turbine engine, which may be generally defined by the longitudinal axis of its shaft 120 (supported by a plurality of bearing assemblies 150). The center axis 95 may be common to or shared with various other engine concentric components. All references to radial, axial, and circumferential directions and measures refer to center axis 95, unless specified otherwise, and terms such as “inner” and “outer” generally indicate a lesser or greater radial distance from, wherein a radial 96 may be in any direction perpendicular and radiating outward from center axis 95.
A gas turbine engine 100 includes an inlet 110, a shaft 120, a compressor 200, a combustor 300, a turbine 400, an exhaust 500, and a power output coupling 600. The gas turbine engine 100 may have a single shaft or a dual shaft configuration.
The compressor 200 includes a compressor rotor assembly 210, compressor stationary vanes (stators) 250, and inlet guide vanes 255. The compressor rotor assembly 210 mechanically couples to shaft 120. As illustrated, the compressor rotor assembly 210 is an axial flow rotor assembly. The compressor rotor assembly 210 includes one or more compressor disk assemblies 220. Each compressor disk assembly 220 includes a compressor rotor disk that is circumferentially populated with compressor rotor blades. Stators 250 axially follow each of the compressor disk assemblies 220. Each compressor disk assembly 220 paired with the adjacent stators 250 that follow the compressor disk assembly 220 is considered a compressor stage. Compressor 200 includes multiple compressor stages. Inlet guide vanes 255 axially precede the compressor stages.
The combustor 300 includes one or more fuel injectors 310 and includes one or more combustion chambers 390.
The turbine 400 includes a turbine rotor assembly 410 and turbine nozzles 450. The turbine rotor assembly 410 mechanically couples to the shaft 120. As illustrated, the turbine rotor assembly 410 is an axial flow rotor assembly. The turbine rotor assembly 410 includes one or more turbine disk assemblies 420. Each turbine disk assembly 420 includes a turbine disk that is circumferentially populated with turbine blades. A turbine nozzle 450 or nozzle ring axially precedes each of the turbine disk assemblies 420. Each turbine nozzle 450 includes multiple nozzle segments 451 grouped together to form a ring. Each turbine disk assembly 420 paired with the adjacent turbine nozzle 450 that precede the turbine disk assembly 420 is considered a turbine stage. Turbine 400 includes multiple turbine stages.
The turbine 400 may also include a turbine housing 430 and turbine diaphragms 440. Turbine housing 430 may be located radially outward from turbine rotor assembly 410 and turbine nozzles 450. Turbine housing 430 may include one or more cylindrical shapes. Each nozzle segment 451 may be configured to attach, couple to, or hang from turbine housing 430. Each turbine diaphragm 440 may axially precede each turbine disk assembly 420 and may be adjacent a turbine disk. Each turbine diaphragm 440 may also be located radially inward from a turbine nozzle 450. Each nozzle segment 451 may also be configured to attach or couple to a turbine diaphragm 440.
The exhaust 500 includes an exhaust diffuser 520 and an exhaust collector 550. The power output coupling 600 may be located at an end of shaft 120.
Upper shroud 452 may also include upper forward rail 454 and upper aft rail 455. Upper forward rail 454 extends radially outward from upper endwall 453. In the embodiment illustrated in
Upper aft rail 455 may also extend radially outward from upper endwall 453. In the embodiment illustrated in
Lower shroud 456 is located radially inward from upper shroud 452. Lower shroud 456 may also be located adjacent and radially outward from turbine diaphragm 440 when nozzle segment 451 is installed in gas turbine engine 100. Lower shroud 456 includes lower endwall 457. Lower endwall 457 may be a portion or a sector of an annular shape, such as a toroid or a hollow cylinder. The toroidal shape may be defined by a cross-section with an outer edge including a convex shape. Multiple lower endwalls 457 are arranged to form the annular shape and to define the radially inner surface of the flow path through a turbine nozzle 450. Lower endwall 457 may be coaxial to upper endwall 453 and center axis 95 when installed in the gas turbine engine 100.
Lower shroud 456 may also include lower forward rail 458 and lower aft rail 459. Lower forward rail 458 extends radially inward from lower endwall 457. In the embodiment illustrated in
Lower aft rail 459 may also extend radially inward from lower endwall 457. In the embodiment illustrated in
Airfoil 460 extends between upper endwall 453 and lower endwall 457. Airfoil 460 includes leading edge 461, trailing edge 462, pressure side wall 463, and suction side wall 464. Leading edge 461 extends from upper endwall 453 adjacent an axial end of upper endwall 453 to lower endwall 457 adjacent an axial end of lower endwall 457. Leading edge 461 may be located near upper forward rail 454 and lower forward rail 458. Trailing edge 462 extends from upper endwall 453 distal to leading edge 461, adjacent the axial end of upper endwall 453 opposite the location of leading edge 461 and from lower endwall 457 adjacent the axial end of upper endwall 453 opposite or axial distal to the location of leading edge 461. When nozzle segment 451 is installed in gas turbine engine 100, leading edge 461, upper forward rail 454, and lower forward rail 458 may be located axially forward and upstream of trailing edge 462, upper aft rail 455, and lower aft rail 459. Leading edge 461 may be the point at the upstream end of airfoil 460 with the maximum curvature and trailing edge 462 may be the point at the downstream end of airfoil 460 with maximum curvature. In the embodiment illustrated in
Pressure side wall 463 may span from leading edge 461 to trailing edge 462 between upper endwall 453 and lower endwall 457. Pressure side wall 463 may include a concave shape. Pressure side wall 463 may also include a pressure side surface 469, the outer surface of pressure side wall 463, with a concave shape. Suction side wall 464 may also span from leading edge 461 to trailing edge 462 between upper endwall 453 and lower endwall 457. Suction side wall 464 may include a convex shape. Leading edge 461, trailing edge 462, pressure side wall 463 and suction side wall 464 may form a cooling cavity 485 (illustrated in
Airfoil 460 may also include multiple groupings of film cooling holes or apertures. Each cooling hole or aperture may be a channel extending through a wall of the airfoil, such as the pressure side wall 463. In the embodiment illustrated in
Forward cooling apertures 466 may be grouped together and located within the third of pressure side wall 463 that is adjacent leading edge 461. Forward cooling apertures 466 may be proximate showerhead cooling apertures 465. In embodiments, forward cooling apertures 466 are located from ⅛ to ¼ of the length of pressure side wall 463 from showerhead cooling apertures 465. In other embodiments, forward cooling apertures 466 are located 1/6 of the length of pressure side wall 463 from showerhead cooling apertures 465. In yet other embodiments, forward cooling apertures 466 are located at least ⅛ of the length of pressure side wall 463 from showerhead cooling apertures 465. Forward cooling apertures 466 may be grouped together between upper endwall 453 and lower endwall 457. In the embodiment illustrated in
Aft cooling apertures 467 may be grouped together and located within the third of pressure side wall 463 that is adjacent to trailing edge 462. Aft cooling apertures 467 may be proximate trailing edge 462. In embodiments, aft cooling apertures are located from ⅛ to ¼ of the length of pressure side wall 463 from trailing edge 462. In other embodiments, aft cooling apertures 467 are located 1/6 of the length of pressure side wall 463 from trailing edge 462. In yet other embodiments, aft cooling apertures 467 are located at least ⅛ of the length of pressure side wall 463 from trailing edge 462. Aft cooling apertures 467 may be arranged radially between upper endwall 453 and lower endwall 457. In the embodiment illustrated in
Intermediate cooling apertures 468 may be grouped together and located within the middle third of pressure side wall 463. Intermediate cooling apertures 468 may be between forward cooling apertures 466 and trailing edge 462. Intermediate cooling apertures 468 may also be between forward cooling apertures 466 and aft cooling apertures 467. In some embodiments, intermediate cooling apertures 468 are located from ¼ to ⅜ of the length of pressure side wall 463 from forward cooling apertures 466 and 1/4 to 3/8 of the length of pressure side wall 463 from aft cooling apertures 467. In other embodiments, intermediate cooling apertures 468 are located 1/3 of the length of pressure side wall 463 from forward cooling apertures 466 and 1/3 of the length of pressure side wall 463 from aft cooling apertures 467. In yet other embodiments, intermediate cooling apertures 468 are located at least ⅛ of the length of pressure side wall 463 from forward cooling apertures 466 and at least ⅛ of the length of pressure side wall 463 from aft cooling apertures 467. Intermediate cooling apertures 468 may be arranged radially between upper endwall 453 and lower endwall 457. In the embodiment illustrated in
While the embodiment illustrated in
Airfoil 460 may further include slots 483. Slots 483 may be located on pressure side wall 463 and may be adjacent trailing edge 462. Slots 483 may be rectangular and may be aligned in the radial direction between upper endwall 453 and lower endwall 457. Slots 483 may extend from cooling cavity 485 to trailing edge 462.
In the embodiment illustrated in
The various components of nozzle segment 451 including upper shroud 452, lower shroud 456, airfoil 460, and second airfoil 470 may be integrally cast or metalurgically bonded to form a unitary or one piece assembly thereof.
In accordance with embodiments of this invention, the spaced apart groups of cooling apertures, showerhead cooling apertures 465, forward cooling apertures 466, intermediate cooling apertures 468, and aft cooling apertures 467, alternate in directionality, being angled or partially angled at lower endwall 457 or upper endwall 453. The directionality or angle of the apertures directs cooling air in a selected direction. In the embodiment illustrated in
In embodiments that include the second intermediate cooling apertures and showerhead cooling apertures 465 angled toward lower endwall 457, second intermediate cooling apertures would be the grouping after intermediate cooling apertures 468 and would be angled towards upper endwall 453 and aft cooling apertures 467 would be the grouping after the second intermediate cooling apertures and would be angled toward lower endwall 457.
Each showerhead cooling aperture 465 may also be angled towards the lower endwall 457 or the upper endwall 453 relative to the direction normal to leading edge 461 at the location where the showerhead cooling aperture 465 is located.
As illustrated in
Referring to
Referring to
Referring to
In embodiments not including intermediate cooling apertures 468 or embodiments including second intermediate cooling apertures, aft compound angle 487 may be angled toward lower endwall 457 relative to the flow direction or reference line 482. In one of these embodiments, aft compound angle 487 is from fifteen to forty-five degrees. In another of these embodiments, aft compound angle 487 is approximately thirty degrees.
Aft cooling apertures 467 may also include an aft injection angle 442. Aft injection angle 442 may be the component of the angle of aft cooling apertures 467 in the plane perpendicular to pressure side surface 469. Aft injection angle 442 may be measured relative to a line extending toward trailing edge 462 and tangent to pressure side surface 469 at the location of each aft cooling aperture 467. In one embodiment, aft injection angle 442 is from fifteen to fifty degrees. In another embodiment, aft injection angle 442 is approximately thirty degrees.
Intermediate cooling apertures 468 may also include an intermediate injection angle 443. Intermediate injection angle 443 may be the component of the angle of intermediate cooling apertures 468 in the plane perpendicular to pressure side surface 469. Intermediate injection angle 443 may be measured relative to a line extending toward trailing edge 462 and tangent to pressure side surface 469 at the location of each intermediate cooling aperture 468. In one embodiment, intermediate injection angle 443 is from fifteen to fifty degrees. In another embodiment, intermediate injection angle 443 is approximately thirty degrees.
Cooling cavity 485 may be a single cavity or may be subdivided into multiple cavities. In the embodiment illustrated in
Each forward cooling aperture 466 may include forward inlet end 493 adjacent cooling cavity 485 and forward outlet end 494 adjacent or at pressure side surface 469. Each intermediate cooling aperture 468 may include intermediate inlet end 497 adjacent cooling cavity 485 and intermediate outlet end 498 adjacent or at pressure side surface 469. Each aft cooling aperture 467 may include aft inlet end 495 adjacent cooling cavity 485 and aft outlet end 496 adjacent or at pressure side surface 469.
The compound angles may be determined by the positions of the inlet ends and the outlet ends of the apertures relative to lower endwall 457 and upper endwall 453, while the injection angle may be determined by the positions of the inlet ends and the outlet ends relative to leading edge 461 and trailing edge 462.
In the embodiment illustrated in
In the embodiment illustrated in
In the embodiment illustrated in
One or more of the above components (or their subcomponents) may be made from stainless steel and/or durable, high temperature materials known as “superalloys”. A superalloy, or high-performance alloy, is an alloy that exhibits excellent mechanical strength and creep resistance at high temperatures, good surface stability, and corrosion and oxidation resistance. Superalloys may include materials such as HASTELLOY, alloy x, INCONEL, WASPALOY, RENE alloys, HAYNES alloys, alloy 188, alloy 230, INCOLOY, MP98T, TMS alloys, and CMSX single crystal alloys.
Gas turbine engines may be suited for any number of industrial applications such as various aspects of the oil and gas industry (including transmission, gathering, storage, withdrawal, and lifting of oil and natural gas), the power generation industry, cogeneration, aerospace, and other transportation industries.
Referring to
Once compressed air 10 leaves the compressor 200, it enters the combustor 300, where it is diffused and fuel is added. Air 10 and fuel are injected into the combustion chamber 390 via fuel injector 310 and combusted. Energy is extracted from the combustion reaction via the turbine 400 by each stage of the series of turbine disk assemblies 420. Exhaust gas 90 may then be diffused in exhaust diffuser 520, collected and redirected. Exhaust gas 90 exits the system via an exhaust collector 550 and may be further processed (e.g., to reduce harmful emissions, and/or to recover heat from the exhaust gas 90).
Operating efficiency of a gas turbine engine generally increases with a higher combustion temperature. Thus, there is a trend in gas turbine engines to increase the temperatures. Gas reaching forward stages of a turbine from a combustion chamber 390 may be 1000 degrees Fahrenheit or more. To operate at such high temperatures a portion of the compressed air from the compressor 200, cooling air, may be diverted through internal passages or chambers to cool various components of a turbine including turbine nozzle segments such as nozzle segment 451. However, the use of cooling air may reduce the operating efficiency of the gas turbine engine.
Alternating the direction of groupings of cooling apertures such as showerhead cooling apertures 465, forward cooling apertures 466, intermediate cooling apertures 468, and aft cooling apertures 467, to direct cooling air towards upper endwall 453 of upper shroud 452 and lower endwall 457 of lower shroud 456 may reduce the temperatures of upper endwall 453 and lower endwall 457, which may improve the operating life of nozzle segment 451.
The first order cooling or initial use of the cooling air exiting showerhead cooling apertures 465, forward cooling apertures 466, intermediate cooling apertures 468, and aft cooling apertures 467 may be to film cool pressure side wall 463. The second order cooling or second use of the cooling air may be to reduce the temperatures of upper endwall 453 and lower endwall 457.
The cooling air may be directed through turbine housing 430, turbine diaphragm 440, or both and into cooling cavity 485. The cooling air may then be directed through the cooling apertures including showerhead cooling apertures 465, forward cooling apertures 466, intermediate cooling apertures 468, and aft cooling apertures 467. The cooling air may also be used for cooling airfoil 460 internally prior to passing through the cooling apertures. The multiple uses of the cooling air that may include the first order film cooling, the second order endwall cooling, and the internal cooling may reduce the amount of air needed to effectively cool nozzle segment 451. Reducing the amount of air needed to cool nozzle segment 451 may improve or increase the efficiency of gas turbine engine 100.
The cooling apertures of second airfoil 470 may be used in the same or a similar manner as the cooling apertures of airfoil 460 resulting in a further reduction of the temperatures of upper endwall 453 and lower endwall 457, as well as the reduction in the amount of cooling air needed to effectively cool each nozzle segment 451.
The preceding detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. The described embodiments are not limited to use in conjunction with a particular type of gas turbine engine. Hence, although the present disclosure, for convenience of explanation, depicts and describes a particular nozzle segment, it will be appreciated that the nozzle segment in accordance with this disclosure can be implemented in various other configurations, can be used with various other types of gas turbine engines, and can be used in other types of machines. Furthermore, there is no intention to be bound by any theory presented in the preceding background or detailed description. It is also understood that the illustrations may include exaggerated dimensions to better illustrate the referenced items shown, and are not consider limiting unless expressly stated as such.