US20140377054A1

US20140377054A1 - Nozzle film cooling with alternating compound angles

Info

Publication number: US20140377054A1
Application number: US13/924,178
Authority: US
Inventors: Luzeng ZHANG; Juan Yin; Hee Koo Moon
Original assignee: Solar Turbines Inc
Current assignee: Solar Turbines Inc
Priority date: 2013-06-21
Filing date: 2013-06-21
Publication date: 2014-12-25
Also published as: RU2016100014A; CN105339591B; WO2014205249A1; CN105339591A

Abstract

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 multiple groups of cooling apertures spaced apart and alternating in directionality such that a first grouping of cooling apertures is angled toward the first endwall, a second grouping of cooling apertures is angled toward the second endwall and spaced apart from the first grouping of cooling apertures, and a third grouping of cooling apertures are angled toward the first endwall and spaced apart from the second grouping of cooling apertures.

Description

TECHNICAL FIELD

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.

BACKGROUND

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.

SUMMARY OF THE DISCLOSURE

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.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic illustration of an exemplary gas turbine engine.

FIG. 2 is a perspective view of a nozzle segment for the gas turbine engine of FIG. 1.

FIG. 3 is a cross-sectional view of a portion of the nozzle segment of FIG. 2 showing the showerhead cooling apertures.

FIG. 4 is a detailed view of the forward cooling apertures of FIG. 2.

FIG. 5 is a detailed view of the intermediate cooling apertures of FIG. 2.

FIG. 6 is a detailed view of the aft cooling apertures of FIG. 2.

FIG. 7 is a cross-section of the airfoil of FIG. 2.

DETAILED DESCRIPTION

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.
FIG. 1 is a schematic illustration of an exemplary gas turbine engine 100. Some of the surfaces have been left out or exaggerated (here and in other figures) for clarity and ease of explanation. Also, the disclosure may reference a forward and an aft direction. Generally, all references to “forward” and “aft” are associated with the flow direction of primary air (i.e., air used in the combustion process), unless specified otherwise. For example, forward is “upstream” relative to primary air flow, and aft is “downstream” relative to primary air flow.
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.
FIG. 2 is a perspective view of a nozzle segment 451 for the gas turbine engine 100 of FIG. 1. Nozzle segment 451 includes upper shroud 452, lower shroud 456, airfoil 460, and second airfoil 470. In other embodiments, nozzle segment 451 can include more or fewer airfoils. Upper shroud 452 may be located adjacent and radially inward from turbine housing 430 when nozzle segment 451 is installed in gas turbine engine 100. Upper shroud 452 includes upper endwall 453. Upper endwall 453 may be a portion or a sector of an annular shape, such as a sector of a toroid or a sector of a hollow cylinder. The toroidal shape may be defined by a cross-section with an inner edge including a convex shape. Multiple upper endwalls 453 are arranged to form the annular shape and to define the radially outer surface of the flow path through a turbine nozzle 450. Upper endwall 453 may be coaxial to center axis 95 when installed in the gas turbine engine 100.
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 FIG. 2, upper forward rail 454 extends from upper endwall 453 at an axial end of upper endwall 453. In other embodiments, upper forward rail 454 extends from upper endwall 453 near or adjacent to an axial end of upper endwall 453. Upper forward rail 454 may include a lip, protrusion or other features that may be used to secure nozzle segment 451 to turbine housing 430.
Upper aft rail 455 may also extend radially outward from upper endwall 453. In the embodiment illustrated in FIG. 2, upper aft rail 455 is ‘L’ shaped, with a first portion extending radially outward from the axial end of upper endwall 453 opposite the location of upper forward rail 454, and a second portion extending in the direction opposite the location of upper forward rail 454 extending axially beyond upper endwall 453. In other embodiments, upper aft rail 455 includes other shapes and may be located near or adjacent to the axial end of upper endwall 453 opposite the location of upper forward rail 454. Upper aft rail 455 may also include other features that may be used to secure nozzle segment 451 to turbine housing 430.
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 FIG. 2, lower forward rail 458 extends from lower endwall 457 at an axial end of lower endwall 457. In other embodiments, lower forward rail 458 extends from lower endwall 457 near or adjacent to an axial end of lower endwall 457. Lower forward rail 458 may include a lip, protrusion or other features that may be used to secure nozzle segment 451 to turbine diaphragm 440.
Lower aft rail 459 may also extend radially inward from lower endwall 457. In the embodiment illustrated in FIG. 2, lower aft rail 459 extends from lower endwall 457 near or adjacent to the axial end of lower endwall 457 opposite the location of lower forward rail 458. In other embodiments, lower aft rail 459 extends from the axial end of lower endwall 457 opposite the location of lower forward rail 458. Lower aft rail 459 may also include a lip, protrusion or other features that may be used to secure nozzle segment 451 to turbine diaphragm 440.
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 FIG. 1, nozzle segment 451 is part of the first stage turbine nozzle adjacent combustion chamber 390. In other embodiments, nozzle segment 451 is located within a turbine nozzle 450 of another stage.
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 FIGS. 3 and 6) or cooling cavities there between. Upper endwall 453, lower endwall 457, or both may include a hole, holes, or pathways for cooling air (not shown) to enter the cooling cavity 485.
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 FIG. 2, airfoil 460 includes showerhead cooling apertures 465, forward cooling apertures 466, aft cooling apertures 467, and intermediate cooling apertures 468. Showerhead cooling apertures 465 are located at leading edge 461 and may be grouped together along leading edge 461. Showerhead cooling apertures 465 may be arranged in columns. In the embodiment shown in FIG. 2, showerhead cooling apertures 465 are arranged in six columns, each column extending in the radial direction between upper endwall 453 and lower endwall 457. In other embodiments, showerhead cooling apertures 465 may be arranged in four to seven columns or may be arranged in other configurations. The portions of pressure side wall 463 and suction side wall 464 adjacent leading edge 461 may include showerhead cooling apertures 465.
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 FIG. 2, forward cooling apertures 466 are arranged in a single radial column and spaced apart radially at 3.5 pitch over diameter, the distance between the centers of adjacent apertures over the diameter of the apertures. In other embodiments, forward cooling apertures 466 are spaced apart radially from 3 to 4 pitch over diameter. Forward cooling apertures 466 may overlap with an adjacent forward cooling aperture 466 rather than align in the radial direction along the surface of pressure side wall 463.
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 FIG. 2, aft cooling apertures 467 are arranged in a single radial column and spaced apart radially at 3.5 pitch over diameter. In other embodiments, aft cooling apertures 467 are spaced apart radially from 3 to 4 pitch over diameter. Aft cooling apertures 467 may overlap with an adjacent aft cooling aperture 467 rather than align in the radial direction along the surface of pressure side wall 463.
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 FIG. 2, intermediate cooling apertures 468 are arranged in a single radial column and spaced apart radially at 3.5 pitch over diameter. In other embodiments, intermediate cooling apertures 468 are spaced apart radially from 3 to 4 pitch over diameter. Intermediate cooling apertures 468 may overlap with an adjacent intermediate cooling aperture 468 rather than being aligned in the radial direction along the surface of pressure side wall 463.
While the embodiment illustrated in FIG. 2 includes forward cooling apertures 466, aft cooling apertures 467, and intermediate cooling apertures 468, some embodiments do not include aft cooling apertures 467 and other embodiments include second intermediate cooling apertures located between intermediate cooling apertures 468 and aft cooling apertures 467. The second intermediate cooling apertures may be arranged similar to the arrangements of forward cooling apertures 466, aft cooling apertures 467, and intermediate cooling apertures 468. Other groups or columns of cooling apertures may also be included. The spacing between groups or columns of cooling apertures may be dependent on the number of groups or columns of cooling apertures located along pressure side wall 463.
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 FIG. 2, nozzle segment 451 includes second airfoil 470. Second airfoil 470 may include the same or similar features as airfoil 460 including second leading edge 471, second trailing edge (not shown), second pressure side wall 473, and second suction side wall 474. Second airfoil 470 may further include second showerhead cooling apertures 475, second forward cooling apertures 476, second aft cooling apertures 477, second intermediate cooling apertures 478, and second slots (not shown). The description of second leading edge 471, the second trailing edge, second pressure side wall 473, second suction side wall 474, second showerhead cooling apertures 475, second forward cooling apertures 476, second aft cooling apertures 477, second intermediate cooling apertures 478, and the second slots may be oriented in the same or a similar manner as leading edge 461, trailing edge 462, pressure side wall 463, suction side wall 464, showerhead cooling apertures 465, forward cooling apertures 466, aft cooling apertures 467, intermediate cooling apertures 468, and slots 483 respectively. In other embodiments, nozzle segment 451 only includes airfoil 460 and not second airfoil 470.
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 FIGS. 2-6, showerhead cooling apertures 465 are angled toward lower endwall 457, forward cooling apertures 466, the next grouping of cooling holes along pressure side wall 463, are angled toward upper endwall 453, intermediate cooling apertures 468, the following grouping of cooling holes along pressure side wall 463, are also angled at lower endwall 457, and aft cooling apertures 467, the last grouping of cooling holes, are also angled at upper endwall 453. In other embodiments, showerhead cooling apertures 465 are angled toward upper endwall 453, forward cooling apertures 466 are angled toward lower endwall 457, intermediate cooling apertures 468 are also angled at upper endwall 453, and aft cooling apertures 467 are also angled at lower endwall 457.
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.
FIG. 3 is a cross-sectional view of a portion of the nozzle segment 451 of FIG. 2 showing the showerhead cooling apertures 465. In the embodiment illustrated in FIG. 3, showerhead cooling apertures 465 may extend through a wall 444 of airfoil 460 a cooling cavity 485 towards lower endwall 457. Wall 444 may be a part of leading edge 461, pressure side wall 463, or suction side wall 464. Showerhead cooling apertures 465 may be angled relative to a reference plane 480. Reference plane 480 may be defined as a plane perpendicular to a radial extending from the nozzle axis, the axis of upper shroud 452 and lower shroud 456 and along leading edge 461. In one embodiment, showerhead angle 481, the angle of showerhead cooling apertures 465 relative to reference plane 480 is from twenty to forty-five degrees towards the lower endwall 457. In another embodiment, showerhead angle 481 is thirty degrees or within a predetermined tolerance of thirty degrees. The predetermined tolerance may be the engineering tolerance or the manufacturing tolerance. While showerhead angle 481 is directed toward lower endwall 457 in the embodiment illustrated, showerhead angle 481 may be directed toward upper endwall 453 in other embodiments.
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 FIG. 3, each showerhead cooling aperture 465 may include showerhead inlet end 491 adjacent cooling cavity 485 and showerhead outlet end 492 located at the surface of leading edge 461. Showerhead inlet end 491 may be radially closer to the upper endwall 453 than showerhead outlet end 492 and showerhead outlet end 492 may be radially closer to the lower endwall 457 than showerhead inlet end 491.
FIG. 4 is a detailed view of the forward cooling apertures 466 of FIG. 2. FIG. 5 is a detailed view of the intermediate cooling apertures 468 of FIG. 2. FIG. 6 is a detailed view of the aft cooling apertures 467 of FIG. 2. Referring to FIGS. 4, 5, and 6, forward cooling apertures 466, aft cooling apertures 467, and intermediate cooling apertures 468 may be angled relative to the flow direction of the air traveling through turbine nozzle 450 along pressure side surface 469 during operation of gas turbine engine 100. Reference line 482 illustrates the flow direction. Reference line 482 may also be defined as the intersection between pressure side surface 469 and a plane perpendicular to a radial extending from the turbine nozzle axis, the axis of upper shroud 452 and lower shroud 456, along the pressure side surface 469.
Referring to FIGS. 2 and 4, forward cooling apertures 466 may be angled at a forward compound angle 486. Forward compound angle 486 may be the component of the angle of forward cooling apertures 466 in the plane of pressure side surface 469. As illustrated, forward compound angle 486 is angled toward upper endwall 453 relative to the flow direction or reference line 482. In one embodiment, forward compound angle 486 is from fifteen to forty-five degrees. In another embodiment, forward compound angle 486 is thirty degrees or within a predetermined tolerance of thirty degrees. The predetermined tolerance may be the engineering tolerance or the manufacturing tolerance. Zero degrees may be the flow direction of the direction along reference line 482 traveling from leading edge 461 to trailing edge 462. While forward compound angle 486 is directed towards upper endwall 453 in the embodiment illustrated, forward compound angle 486 is directed towards lower endwall 457 in embodiments where showerhead angle 481 is directed towards upper endwall 453.
Referring to FIGS. 2 and 5, intermediate cooling apertures 468 may be angled at an intermediate compound angle 488. Intermediate compound angle 488 may be the component of the angle of intermediate cooling apertures 468 in the plane of pressure side surface 469. As illustrated, intermediate compound angle 488 is angled toward the lower endwall 457 relative to the flow direction or reference line 482. In one embodiment, intermediate compound angle 488 is from fifteen to forty-five degrees. In another embodiment, intermediate compound angle 488 is thirty degrees or within a predetermined tolerance of thirty degrees. The predetermined tolerance may be the engineering tolerance or the manufacturing tolerance. While intermediate compound angle 488 is directed towards lower endwall 457 in the embodiment illustrated, intermediate compound angle 488 is directed towards upper endwall 453 in embodiments where forward compound angle 486 is directed towards lower endwall 457.
Referring to FIGS. 2 and 6, aft cooling apertures 467 may be angled at an aft compound angle 487. Aft compound angle 487 may be the component of the angle of aft cooling apertures 467 in the plane of pressure side surface 469. As illustrated, aft compound angle 487 is angled toward the upper endwall 453 relative to the flow direction or reference line 482 and may be similar or equal to forward compound angle 486. In one embodiment, aft compound angle 487 is from fifteen to forty-five degrees. In another embodiment, aft compound angle 487 is thirty degrees or within a predetermined tolerance of thirty degrees. The predetermined tolerance may be the engineering tolerance or the manufacturing tolerance. While aft compound angle 487 is directed towards upper endwall 453 in the embodiment illustrated, aft compound angle 487 may be directed towards lower endwall 457 in embodiments where intermediate compound angle 488 is directed towards upper endwall 453.
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.
FIG. 7 is a cross-section of the airfoil 460 of FIG. 2. Referring to FIG. 7, forward cooling apertures 466 may also include a forward injection angle 441. Forward injection angle 441 may be the component of the angle of forward cooling apertures 466 in the plane perpendicular to pressure side surface 469. Forward injection angle 441 may be measured relative to a line extending toward trailing edge 462 and tangent to pressure side surface 469 at the location of each forward cooling aperture 466. In one embodiment, forward injection angle 441 is from fifteen to fifty degrees. In another embodiment, forward injection angle 441 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 FIG. 7, cooling cavity 485 is subdivided into two cooling cavities.
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 FIG. 4, forward inlet end 493 is radially closer to lower endwall 457 and axially closer to leading edge 461 than forward outlet end 494, and forward outlet end 494 is radially closer to upper endwall 453 and axially closer to trailing edge 462 than forward inlet end 493. In other embodiments, forward inlet end 493 is radially closer to upper endwall 453 and axially closer to leading edge 461 than forward outlet end 494, and forward outlet end 494 is radially closer to lower endwall 457 and axially closer to trailing edge 462 than forward inlet end 493.
In the embodiment illustrated in FIG. 5, intermediate inlet end 497 is radially closer to upper endwall 453 and axially closer to leading edge 461 than intermediate outlet end 498, and intermediate outlet end 498 is radially closer to lower endwall 457 and axially closer to trailing edge 462 than intermediate inlet end 497. In other embodiments, intermediate inlet end 497 is radially closer to lower endwall 457 and axially closer to leading edge 461 than intermediate outlet end 498, and intermediate outlet end 498 is radially closer to upper endwall 453 and axially closer to trailing edge 462 than intermediate inlet end 497.
In the embodiment illustrated in FIG. 6, aft inlet end 495 is radially closer to lower endwall 457 and axially closer to leading edge 461 than aft outlet end 496, and aft outlet end 496 is radially closer to upper endwall 453 and axially closer to trailing edge 462 than aft inlet end 495. In other embodiments, aft inlet end 495 is radially closer to upper endwall 453 and axially closer to leading edge 461 than aft outlet end 496, and aft outlet end 496 is radially closer to lower endwall 457 and axially closer to trailing edge 462 than aft inlet end 495.
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.

INDUSTRIAL APPLICABILITY

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 FIG. 1, a gas (typically air 10) enters the inlet 110 as a “working fluid”, and is compressed by the compressor 200. In the compressor 200, the working fluid is compressed in an annular flow path 115 by the series of compressor disk assemblies 220. In particular, the air 10 is compressed in numbered “stages”, the stages being associated with each compressor disk assembly 220. For example, “4th stage air” may be associated with the 4th compressor disk assembly 220 in the downstream or “aft” direction, going from the inlet 110 towards the exhaust 500). Likewise, each turbine disk assembly 420 may be associated with a numbered stage.
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.

Claims

What is claimed is:

1. A nozzle segment for a nozzle ring of a gas turbine engine, the nozzle segment comprising:

a first endwall;

a second endwall; and

an airfoil extending between the first endwall and the second endwall, the airfoil including

a leading edge extending radially from the first endwall to the second endwall,

a trailing edge extending radially from the first endwall to the second endwall axially distal to the leading edge,

a pressure side wall extending from the leading edge to the trailing edge,

a suction side wall extending from the leading edge to the trailing edge,

a plurality of showerhead cooling apertures spanning along the leading edge,

a plurality of forward cooling apertures grouped together in the pressure side wall proximate the plurality of showerhead cooling apertures, and

a plurality of intermediate cooling apertures 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 alternating 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.

2. The nozzle segment of claim 1, wherein each showerhead cooling aperture is angled toward the first endwall as each showerhead cooling aperture extends through a wall of the airfoil, each forward cooling aperture includes a forward compound angle from fifteen to forty-five degrees towards the second endwall relative to a flow direction of air through the nozzle segment during operation, and each intermediate cooling aperture including an intermediate compound angle from fifteen to forty-five degrees towards the first endwall relative to a flow direction of air through the nozzle segment during operation.

3. The nozzle segment of claim 2, wherein the airfoil further includes a plurality of aft cooling apertures grouped together in the pressure side wall proximate the trailing edge, each aft cooling aperture including an aft compound angle from fifteen to forty-five degrees towards the second endwall relative to a flow direction of air through the nozzle segment during operation.

4. The nozzle segment of claim 1, wherein the plurality of forward cooling apertures are arranged in a single column and the plurality of intermediate cooling apertures are arranged in a single column.

5. The nozzle segment of claim 1, wherein the first endwall is a lower endwall and the second endwall is an upper endwall located radially outward from the lower endwall.

6. The nozzle segment of claim 1, wherein each forward cooling aperture is spaced apart from an adjacent forward cooling aperture from 3 to 4 pitch over diameter and each intermediate cooling aperture is spaced apart from an adjacent intermediate cooling aperture from 3 to 4 pitch over diameter.

7. The nozzle segment of claim 1, wherein the plurality of showerhead cooling apertures is configured to direct air to film cool the leading edge and cool the first endwall, the plurality of forward cooling apertures is configured to direct air to film cool a pressure side surface of the pressure side wall and cool the second endwall, and the plurality of intermediate cooling apertures is configured to direct air to film cool the pressure side surface and cool the first endwall.

8. A gas turbine engine including the nozzle segment of claim 1, wherein the nozzle segment is located in a first stage turbine nozzle of the gas turbine engine.

9. A nozzle segment for a nozzle ring of a gas turbine engine, the nozzle segment comprising:

a first endwall;

a second endwall;

a leading edge extending radially from the first endwall to the second endwall,

a trailing edge extending radially from the first endwall to the second endwall,

a pressure side wall extending from the leading edge to the trailing edge, the pressure side wall including a pressure side surface with a concave shape, the pressure side surface being the outer surface of the pressure side wall,

a suction side wall extending from the leading edge to the trailing edge,

a cooling cavity located between the leading edge, the trailing edge, the pressure side wall, and the suction side wall,

a plurality of showerhead cooling apertures spanning along the leading edge, each showerhead cooling aperture including a showerhead inlet end adjacent the cooling cavity and a showerhead outlet end at the outer surface of the leading edge, the showerhead inlet end being radially closer to the second endwall than the showerhead outlet end and the showerhead outlet end being radially closer to the first endwall than the showerhead inlet end,

a plurality of forward cooling apertures in the pressure side wall grouped together and spaced apart from the plurality of showerhead cooling apertures at least ⅛ the length of the pressure side wall, each forward cooling aperture including a forward inlet end adjacent the cooling cavity and a forward outlet end adjacent the pressure side surface, the forward inlet end being radially closer to the first endwall and axially closer to the leading edge than the forward outlet end, and the forward outlet end being radially closer to the second endwall and axially closer to the trailing edge than the forward inlet end, and

a plurality of intermediate cooling apertures in the pressure side wall grouped together and spaced apart from the plurality of forward cooling apertures at least ⅛ the length of the pressure side wall, each intermediate cooling aperture including an intermediate inlet end adjacent the cooling cavity and an intermediate outlet end adjacent the pressure side surface, the intermediate inlet end being radially closer to the second endwall and axially closer to the leading edge than the intermediate outlet end, and the intermediate outlet end being radially closer to the first endwall and axially closer to the trailing edge than the intermediate inlet end.

10. The nozzle segment of claim 9, wherein the airfoil further includes a plurality of aft cooling apertures in the pressure side wall grouped together and spaced apart from the plurality of intermediate cooling apertures at least ⅛ the length of the pressure side wall, each aft cooling aperture including an aft inlet end adjacent the cooling cavity and an aft outlet end adjacent the pressure side surface, the aft inlet end being radially closer to the first endwall and axially closer to the leading edge than the aft outlet end, and the aft outlet end being radially closer to the second endwall and axially closer to the trailing edge than the aft inlet end.

11. The nozzle segment of claim 9, wherein the first endwall is a lower endwall of a lower shroud and the second endwall is an upper endwall of an upper shroud, the lower endwall being located radially inward from the upper endwall.

12. The nozzle segment of claim 9, wherein each forward cooling aperture is spaced apart from an adjacent forward cooling aperture from 3 to 4 pitch over diameter and each intermediate cooling aperture is spaced apart from an adjacent intermediate cooling aperture from 3 to 4 pitch over diameter.

13. The nozzle segment of claim 10, wherein each forward cooling aperture is spaced apart from an adjacent forward cooling aperture from 3 to 4 pitch over diameter, each intermediate cooling aperture is spaced apart from an adjacent intermediate cooling aperture from 3 to 4 pitch over diameter, and each aft cooling aperture is spaced apart from an adjacent aft cooling aperture from 3 to 4 pitch over diameter.

14. A nozzle segment for a nozzle ring of a gas turbine engine, the nozzle segment comprising:

an upper shroud including an upper endwall, the upper endwall being the shape of a sector of a toroid;

a lower shroud including a lower endwall located radially inward from the upper endwall, the lower endwall being the shape of a sector of a toroid;

an airfoil extending between the upper endwall and the lower endwall, the airfoil including

a leading edge extending radially from the upper endwall to the lower endwall,

a trailing edge extending radially from the upper endwall to the lower endwall axially distal to the leading edge,

a pressure side wall extending from the leading edge to the trailing edge, the pressure side wall including a concave shape,

a suction side wall extending from the leading edge to the trailing edge, the suction side wall including a convex shape,

a plurality of showerhead cooling apertures arranged in four to seven columns spanning along the leading edge, each showerhead cooling aperture being angled toward the lower endwall as each showerhead cooling aperture extends through a wall of the airfoil from the cooling cavity,

a plurality of forward cooling apertures arranged in a column extending radially between the upper endwall and the lower endwall and located in the third of the pressure side wall adjacent the leading edge, each forward cooling aperture extending through the pressure side wall from the cooling cavity and including a forward compound angle from fifteen to forty-five degrees towards the upper endwall and the trailing edge relative to a reference line in the plane of a pressure side surface of the pressure side wall, the reference line being defined as an intersection between the pressure side surface and a plane perpendicular to a radial extending from an axis of the upper shroud along the pressure side surface,

a plurality of intermediate cooling apertures arranged in a column extending radially between the upper endwall and the lower endwall and located in the middle third of the pressure side wall between the leading edge and the trailing edge, each intermediate cooling aperture extending through the pressure side wall from the cooling cavity and including an intermediate compound angle from fifteen to forty-five degrees towards the lower endwall and the trailing edge relative to the reference line, and

a plurality of aft cooling apertures arranged in a column extending radially between the upper endwall and the lower endwall and located in the third of the pressure side wall adjacent the trailing edge, each aft cooling aperture extending through the pressure side wall from the cooling cavity and including an aft compound angle from fifteen to forty-five degrees towards the upper endwall and the trailing edge relative to the reference line.

15. The nozzle segment of claim 14, further comprising:

a second airfoil extending between the upper endwall and the lower endwall, the second airfoil including

a second leading edge extending radially from the upper endwall to the lower endwall,

a second trailing edge extending radially from the upper endwall to the lower endwall distal to the second leading edge,

a second pressure side wall extending from the second leading edge to the second trailing edge, the second pressure side wall including a second concave shape,

a second suction side wall extending from the second leading edge to the second trailing edge, the second suction side wall including a second convex shape,

a second cooling cavity located between the second leading edge, the second trailing edge, the second pressure side wall, and the second suction side wall,

a plurality of second showerhead cooling apertures arranged in four to seven columns spanning along the second leading edge, each second showerhead cooling aperture being angled toward the lower endwall as each second showerhead cooling aperture extends through a second wall of the second airfoil from one of the second cooling cavity,

a plurality of second forward cooling apertures arranged in a column extending radially between the upper endwall and the lower endwall and located in the third of the second pressure side wall adjacent the second leading edge, each second forward cooling aperture extending through the second pressure side wall from the second cooling cavity and including a second forward compound angle from fifteen to forty-five degrees towards the upper endwall and the second trailing edge relative to a second reference line in the plane of a second pressure side surface of the second pressure side wall, the second reference line being defined as an intersection between the second pressure side surface and a second plane perpendicular to a radial extending from the axis of the upper shroud along the second pressure side surface,

a plurality of second intermediate cooling apertures arranged in a column extending radially between the upper endwall and the lower endwall and located in the middle third of the second pressure side wall between the second leading edge and the second trailing edge, each second intermediate cooling aperture extending through the second pressure side wall from the second cooling cavity and including a second intermediate compound angle from fifteen to forty-five degrees towards the lower endwall and the second trailing edge relative to the second reference line, and

a plurality of second aft cooling apertures arranged in a column extending radially between the upper endwall and the lower endwall and located in the third of the second pressure side wall adjacent the second trailing edge, each second aft cooling aperture extending through the second pressure side wall from the second cooling cavity and including a second aft compound angle from fifteen to forty-five degrees towards the upper endwall and the second trailing edge relative to the second reference line.

16. The nozzle segment of claim 14, wherein each forward cooling aperture is spaced apart from an adjacent forward cooling aperture from 3 to 4 pitch over diameter, each intermediate cooling aperture is spaced apart from an adjacent intermediate cooling aperture from 3 to 4 pitch over diameter, and each aft cooling aperture is spaced apart from an adjacent aft cooling aperture from 3 to 4 pitch over diameter.

17. The nozzle segment of claim 14, wherein each forward cooling aperture is spaced apart from an adjacent forward cooling aperture at 3.5 pitch over diameter, each intermediate cooling aperture is spaced apart from an adjacent intermediate cooling aperture at 3.5 pitch over diameter, and each aft cooling aperture is spaced apart from an adjacent aft cooling aperture at 3.5 pitch over diameter.

18. The nozzle segment of claim 14, wherein the plurality of forward cooling apertures are spaced apart from the plurality of showerhead cooling apertures from ⅛ to ¼ the length of the pressure side wall, the plurality of intermediate cooling apertures are spaced apart from the plurality of forward cooling apertures from ¼ to ⅜ the length of the pressure side wall, and the plurality of aft cooling apertures are spaced apart from the plurality of intermediate cooling apertures from ¼ to ⅜ of the length of the pressure side wall.

19. The nozzle segment of claim 14, wherein the forward compound angle is within a predetermined tolerance of thirty degrees, the intermediate compound angle is within a predetermined tolerance of thirty degrees, and the aft compound angle is within a predetermined tolerance of thirty degrees.

20. A gas turbine engine including the nozzle segment of claim 14, wherein the nozzle segment is located in a first stage turbine nozzle of the gas turbine engine.