INNER SHROUD COOLING ARRANGEMENT IN A GAS TURBINE ENGINE
A component in a gas turbine engine includes an airfoil and a shroud. The shroud has an outer surface supporting an end of the airfoil and defines a portion of an annular gas path. The shroud includes axial edges extending between upstream and downstream edges thereof. Each of the axial edges includes a seal slot that receives a seal member extending between the shroud and an adjacent shroud. A cooling air channel extends between the upstream and downstream edges of the shroud. A cooling air supply passage extends from a cooling air chamber at an inner surface of the shroud to the cooling air channel. At least one cooling air exit passage extends from the cooling air channel to one of the axial edges. The cooling air channel is located radially between the outer surface of the shroud and the seal slot.
The present invention relates to turbine engines and, more particularly, to cooling arrangements for inner shrouds of vane segments in gas turbine engines.
BACKGROUND OF THE INVENTIONIn a turbomachine, such as a gas turbine engine, air is pressurized in a compressor section then mixed with fuel and burned in a combustor section to generate hot combustion gases. The hot combustion gases are expanded within a turbine section of the engine where energy is extracted to power the compressor section and to produce useful work, such as turning a generator to produce electricity. The hot combustion gases travel through a series of turbine stages within the turbine section. A turbine stage may include a row of stationary airfoils, i.e., vanes, followed by a row of rotating airfoils, i.e., turbine blades, where the turbine blades extract energy from the hot combustion gases for powering the compressor section and providing output power. Because the airfoils, i.e., vanes and turbine blades, are directly exposed to the hot combustion gases, they are typically provided with internal cooling channels that may feed a cooling fluid, such as compressor bleed air, through the airfoil and through various passages formed in structure associated with the vanes and/or blades.
One type of stationary airfoil in a turbine engine is provided as a component of a stator vane segment. The stator vane segment may include a radially inner shroud, a radially outer shroud, and one or more airfoils extending between the inner and outer shrouds. Hot combustion gases, or working gases, may be supplied from a combustor section and pass through passages defined between adjacent airfoils and between the inner and outer shrouds, resulting in some of the heat of the gases being transferred to the vane segments. As turbine engine performance has been increased with increasing combustion gas temperature, there has been a continuing need to improve cooling to the various portions of vane segments in order to avoid or minimize deterioration of the material forming the vane segments.
SUMMARY OF THE INVENTIONIn accordance with a first aspect of the present invention, a component is provided in a gas turbine engine. The component comprises an airfoil and a shroud. The airfoil is adapted to extend radially through an annular hot gas path extending in a generally axial direction through the turbine engine. The airfoil includes a pressure side and a suction side, an upstream leading edge and a downstream trailing edge. The shroud has an outer surface supporting an end of the airfoil and defines a portion of the annular gas path through the gas turbine engine. The shroud includes an upstream edge, a downstream edge, and opposing axial edges extending between the upstream edge and the downstream edge. Each of the axial edges includes a generally axially extending seal slot adapted to receive a seal member extending between the shroud and an adjacent shroud. A cooling air channel extends generally axially substantially parallel to at least one of the axial edges between the upstream edge and the downstream edge. A cooling air supply passage extends from a cooling air chamber at an inner surface of the shroud to the cooling air channel. At least one cooling air exit passage extends from the cooling air channel to the one of the axial edges. The cooling air channel is located radially between the outer surface of the shroud and the seal slot at the one of the axial edges for effecting convective cooling of a corner defined at an intersection of the outer surface and the one of the axial edges.
In accordance with a second aspect of the present invention, a vane is provided in a gas turbine engine. The vane comprises an airfoil and a shroud. The airfoil is adapted to extend radially through an annular hot gas path extending in a generally axial direction through the turbine engine. The airfoil includes a pressure side and a suction side, an upstream leading edge and a downstream trailing edge. The shroud has an outer surface supporting an end of the airfoil and defines a portion of the annular gas path through the gas turbine engine. The shroud includes an upstream edge, a downstream edge, and opposing axial edges extending between the upstream edge and the downstream edge. Each of the axial edges includes a generally axially extending seal slot adapted to receive a seal member extending between the shroud and an adjacent shroud. A cooling air channel extends generally axially substantially parallel to at least one of the axial edges between the upstream edge and the downstream edge. A cooling air supply passage extends from a cooling air chamber at inner surface of the shroud to the cooling air channel. At least one cooling air exit passage extends from the cooling air channel to the one of the axial edges, the cooling air exit passage comprising a purge passage having an exit opening providing a volume of air for purging hot gas from the seal slot and the seal. The exit opening is located at a downstream end of the cooling air channel adjacent the downstream edge of the shroud. The cooling air channel is located radially between the outer surface of the shroud and the seal slot at the one of the axial edges for effecting convective cooling of a corner defined at an intersection of the outer surface and the one of the axial edges.
In accordance with a third aspect of the present invention, a vane is provided in a gas turbine engine. The vane comprises an airfoil and a shroud. The airfoil is adapted to extend radially through an annular hot gas path extending in a generally axial direction through the turbine engine. The airfoil includes a pressure side and a suction side, an upstream leading edge and a downstream trailing edge. The shroud has an outer surface supporting an end of the airfoil and defines a portion of the annular gas path through the gas turbine engine. The shroud includes an upstream edge, a downstream edge, and opposing axial edges extending between the upstream and downstream edges. Each of the axial edges includes a generally axially extending seal slot adapted to receive a seal member extending between the shroud and an adjacent shroud. A cooling air channel extends generally axially substantially parallel to at least one of the axial edges between the upstream and the downstream edges. A cooling air supply passage extends from a cooling air chamber at an inner surface of the shroud to the cooling air channel. A plurality of impingement passages extend from the cooling air channel to the one of the axial edges, the impingement passages having exit openings providing a flow of cooling air impinging on an axial edge of the adjacent shroud, the exit openings being located at an upstream end of the cooling air channel and adjacent the upstream edge of the shroud. The cooling air channel is located radially between the outer surface of the shroud and the seal slot at the one of the axial edges for effecting convective cooling of a corner defined at an intersection of the outer surface and the one of the axial edges.
While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:
In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, specific preferred embodiments in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.
Referring to
Referring to
As noted above, the vane segments 12 are suspended in a circumferential row 10 about the turbine rotor, such that the airfoils 16 are spaced apart and define flow passages 30 therebetween for channeling the hot working gas HG through the turbine section 14 during engine operation. Each flow passage 30 forms a portion of an annular path for the hot working gas HG and is bounded by the pressure sidewall 22 of one airfoil 16 and the suction sidewall 24 of an adjacent airfoil 16. The flow passage 30 is also defined between the inner shroud 18 and the outer shroud 20 and extends in a flow direction from an upstream edge 32 to a downstream edge 34 of the inner shroud 18 (
Each vane segment 12 includes a first generally axially extending mating edge 40 extending between the upstream edge 32 and the downstream edge 34 of the inner shroud 18, and an opposing second generally axially extending mating edge 42 extending generally parallel to the first mating edge 40 between the upstream edge 32 and the downstream edge 34 of the inner shroud 18. Each of the mating edges 40, 42 includes a generally axially extending seal slot 44 adapted to receive an axially extending seal member 46 (
Referring to
Referring now to
As shown in
The inner surface 52 of the outer wall 51 further defines a radially outer boundary for a mid-chord cooling air chamber 60 (
The mid-chord cooling air chamber 60 receives cooling air, i.e., compressor discharge air, for cooling a mid-chord portion 18A and a trailing edge portion 18B of the inner shroud 18, see
The portion of the mid-chord cooling air chamber 60 associated with the trailing edge portion 18B of the inner shroud 18 is associated with a cover plate 66 (
Referring to
As shown in
Each cooling air exit passage 69 is in fluid communication with the cooling air channel 48 and includes an exit opening 70 located between the outer surface 50 of the inner shroud 18 and the seal slot 44 (see
During operation of the engine, the row 10 of vane segments 12, which, as noted above, is a first row 10 of vane segments 12 according to this embodiment, is exposed to the high temperature hot working gas HG entering the turbine section 14 from the combustor assembly. A region of the vane segment 12 along the inner shroud 18 adjacent to and downstream from the trailing edge 28 is believed to be particularly susceptible to exposure to high temperature gases that may lead to elevated temperatures of the surfaces of the inner shroud 18. In particular, the mating edges 40, 42 may be exposed to elevated temperatures in the region 71, which may generally be defined as extending from a location at or near an axial location of the trailing edge 28, as identified by line 73 in
Cooling air, e.g., compressor discharge air, enters the leading edge cooling air chamber 53 from the first cavity 55 through the impingement holes 57A formed in the impingement plate 57. The cooling air entering the chamber 53 through the impingement holes 57A contacts and provides impingement cooling to the inner surface 52 of the outer wall 51.
A first portion of the cooling air in the leading edge cooling air chamber 53 passes into the cooling air channels 48 through the respective primary cooling air supply passages 58. Another portion of the cooling air in the chamber 53 passes through the film cooling holes 59 in the wall 51 and provides film cooling for the outer surface 50 of the wall 51 as discussed above.
The cooling air flowing through the cooling air channels 48 provides convective cooling for the inner shroud 18 adjacent to the mating edges 40, 42. Since the cooling air channels 48 are located close to the mating edges 40, 42 and to the outer surface 50 of the wall 51, the corners 56 of the inner shroud 18 are convectively cooled by the cooling air flowing through the cooling air channels 48.
As a result of the cooling air in the cooling air channels 48 convectively cooling the inner shroud 18 adjacent to the mating edges 40, 42 by removing heat from the inner shroud 18, the cooling air flowing through the cooling air channels 48 is heated as it flows downstream toward the downstream portions 48B of the cooling air channels 48.
The second portion of cooling air, which is provided into the cooling air channels 48 from the mid-chord cooling air chamber 60 via the replenishing cooling air supply passages 68, is added to and cools the first portion of cooling air. Hence, the cooling air that is available for convective cooling in the downstream portions 48B of the cooling air channels 48, i.e., for cooling the mating edges 40, 42 and the corners 56, is cool enough to sufficiently cool the inner shroud 18 adjacent to the downstream edge 34 thereof. That is, the cooling air exiting the cooling air exit passages 69 through the exit openings 70 is cool enough, and comprises an adequate volume, to sufficiently cool the seal member 46 and the mating edges 40, 42 and corners 56 of the inner shroud 18, and to provide a cool air barrier for the hot working gases HG. Further, since the exit openings 70 of the respective adjacent inner shrouds 18 are axially offset from one another, the cooling air exiting the respective exit openings 70 provides a substantially even distribution of cooling air between the mating edges 40, 42 of the respective inner shrouds 18. It is noted that any cooling air in the channels 48 that does not exit through the cooling air exit passages 69 may exit the inner shroud 18 through the outlets 49 of the channels 48.
Referring now to
The vane segment 110 includes a first generally axially extending mating edge 116 extending between an upstream edge 118 and a downstream edge 120 of the inner shroud 114, and an opposing second generally axially extending mating edge 122 extending generally parallel to the first mating edge 116 between the upstream edge 118 and the downstream edge 120 of the inner shroud 114. Each of the mating edges 116, 122 includes a generally axially extending seal slot 124 adapted to receive an axially extending seal member 126 (
As shown in
Referring now to
As shown in
A first portion of cooling air is provided into the cooling air channels 130 via respective primary cooling air supply passages 152 that extend from the leading edge cooling air chamber 150 to the upstream portions 130A of the respective cooling air channels 130, see
Referring to
As shown in
Each cooling air exit passage 158 is in fluid communication with a respective cooling air channel 130 and includes an exit opening 160 located between the outer surface 132 of the inner shroud 114 and the seal slot 124 (see
During operation of the engine, the row of vane segments 110, which, as noted above, is a second row of vane segments 110 according to this embodiment, is exposed to high temperature hot working gas entering the turbine section from the combustor assembly as described above with reference to
Cooling air, e.g., compressor discharge air, enters the mid-chord cooling air chamber 142 from the internal cooling passageways 144 that extend through the airfoils 112. The cooling air entering the mid-chord cooling air chamber 142 provides convective cooling to the inner shroud 114 around the chamber 142.
As noted above, cooling air is provided from the mid-chord chamber cooling air chamber 142 into the leading edge cooling air chamber 150 through the passageways 148. The cooling air in the leading edge cooling air chamber 150 provides convective cooling to the inner shroud 114 adjacent to the upstream edge 118 thereof. A first portion of the cooling air in the leading edge cooling air chamber 150 passes into the cooling air channels 130 through the respective primary cooling air supply passages 152.
The cooling air flowing through the cooling air channels 130 provides convective cooling for the inner shroud 114 adjacent to the mating edges 116, 122. Since the cooling air channels 130 are located close to the mating edges 116, 122 and to the outer surface 132 of the wall 134, the corners 138 of the inner shroud 114 are convectively cooled by the cooling air flowing through the cooling air channels 130. As a result of the cooling air in the channels 130 convectively cooling the inner shroud 114 adjacent to the mating edges 116, 122, the cooling air flowing through the cooling air channels 130 is heated as it flows downstream toward the downstream portions 130B of the cooling air channels 130. Some of the first portion of cooling air exits the channels 130 through the cooling air exit passages 158 near the upstream portions 130A of the channels 130.
The second portion of cooling air, which is provided to the cooling air channels 130 directly from the mid-chord cooling air chamber 142 via the replenishing cooling air supply passages 156, is added to and cools the remaining portion of the first portion of cooling air. Hence, the cooling air that is available for cooling within the cooling air channels 130 is cool enough to sufficiently cool the inner shroud 114, i.e., the corners 138 and the mating edges 116, 122, and also to cool the seal members 126. Additionally, since the exit openings 160 of the adjacent inner shrouds 114 are axially offset from one another, the cooling air exiting the respective exit openings 160 provides a substantially even distribution of cooling air between the mating edges 116, 122 of the respective inner shrouds 114 for providing impingement cooling for the opposed mating edges 116, 122. It is noted that any cooling air in the channels 130 that does not exit through the cooling air exit passages 158 may exit the inner shroud 114 through the outlets 131 of the channels 130.
While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.
Claims
1. A component in a gas turbine engine, said component comprising:
- an airfoil adapted to extend radially through an annular hot gas path extending in a generally axial direction through said turbine engine, said airfoil including a pressure side and a suction side, an upstream leading edge and a downstream trailing edge;
- a shroud having an outer surface supporting an end of said airfoil, said shroud defining a portion of said annular gas path through said gas turbine engine and including an upstream edge and a downstream edge, and opposing axial edges extending between said upstream edge and said downstream edge;
- each of said axial edges including a generally axially extending seal slot adapted to receive a seal member extending between said shroud and an adjacent shroud;
- a cooling air channel extending generally axially substantially parallel to at least one of said axial edges between said upstream edge and said downstream edge;
- a cooling air supply passage extending from a cooling air chamber at an inner surface of said shroud to said cooling air channel;
- at least one cooling air exit passage extending from said cooling air channel to said one of said axial edges; and
- said cooling air channel being located radially between said outer surface of said shroud and said seal slot at said one of said axial edges for effecting convective cooling of a corner defined at an intersection of said outer surface and said one of said axial edges.
2. The component of claim 1, wherein said cooling air channel is located on a radial plane passing through said seal slot at said one of said axial edges.
3. The component of claim 1, wherein said cooling air exit passage includes an exit opening located between said outer surface of said shroud and said seal slot.
4. The component of claim 1, wherein said cooling air exit passage comprises a purge passage having an exit opening providing a volume of air for purging hot gas from said seal slot and said seal, said exit opening being located at a downstream end of said cooling air channel adjacent said downstream edge of said shroud.
5. The component of claim 4, further comprising a replenishing cooling air supply passage extending from a mid-chord cooling air chamber at said inner surface of said shroud to said cooling air channel, said replenishing cooling air supply passage located between said cooling air supply passage and said shroud downstream edge.
6. The component of claim 5, wherein said cooling air supply passage supplies a first portion of cooling air from a leading edge cooling air chamber to a location adjacent an upstream end of said cooling air channel and said replenishing cooling air supply passage supplies a second portion of cooling air from said mid-chord cooling air chamber to a location proximate to said purge passage.
7. The component of claim 5, wherein said component comprises a vane in a first row of vanes within said gas turbine engine.
8. The component of claim 1, wherein said cooling air exit passage comprises a plurality of impingement passages having exit openings providing a flow of cooling air impinging on an axial edge of said adjacent shroud, said exit openings being located at an upstream end of said cooling air channel adjacent said upstream edge of said shroud.
9. The component of claim 8, wherein said upstream end of said cooling air channel receives cooling air from a leading edge cooling air chamber at said inner surface of said shroud.
10. The component of claim 9, wherein said cooling air supply passage comprises a replenishing cooling air supply passage that supplies cooling air from a mid-shroud impingement cavity located at an axial midpoint between said upstream edge and said downstream edge at said inner surface of said shroud, said replenishing cooling air supply passage providing replenishing cooling air to said cooling air channel.
11. The component of claim 9, wherein said component comprises a vane in a second row of vanes within said gas turbine engine.
12. A vane in a gas turbine engine, said vane comprising:
- an airfoil adapted to extend radially through an annular hot gas path extending in a generally axial direction through said turbine engine, said airfoil including a pressure side and a suction side, an upstream leading edge and a downstream trailing edge;
- a shroud having an outer surface supporting an end of said airfoil, said shroud defining a portion of said annular gas path through said gas turbine engine and including an upstream edge and a downstream edge, and opposing axial edges extending between said upstream edge and said downstream edge;
- each of said axial edges including a generally axially extending seal slot adapted to receive a seal member extending between said shroud and an adjacent shroud;
- a cooling air channel extending generally axially substantially parallel to at least one of said axial edges between said upstream edge and said downstream edge;
- a cooling air supply passage extending from a cooling air chamber at an inner surface of said shroud to said cooling air channel;
- at least one cooling air exit passage extending from said cooling air channel to said one of said axial edges, said cooling air exit passage comprising a purge passage having an exit opening providing a volume of air for purging hot gas from said seal slot and said seal, said exit opening being located at a downstream end of said cooling air channel adjacent said downstream edge of said shroud; and
- said cooling air channel being located radially between said outer surface of said shroud and said seal slot at said one of said axial edges for effecting convective cooling of a corner defined at an intersection of said outer surface and said one of said axial edges.
13. The vane of claim 12, further comprising a replenishing cooling air supply passage extending from a mid-chord cooling air chamber at said inner surface of said shroud to said cooling air channel, said replenishing cooling air supply passage located between said cooling air supply passage and said shroud downstream edge.
14. The vane of claim 13, wherein said cooling air supply passage supplies a first portion of cooling air from a leading edge cooling air chamber to a location adjacent an upstream end of said cooling air channel and said replenishing cooling air supply passage supplies a second portion of cooling air from said mid-chord cooling air chamber to a location proximate to said purge passage.
15. The vane of claim 12, wherein said cooling air channel is without openings for discharge of air along a length of said cooling air channel from said cooling air supply passage to within close proximity of said replenishing cooling air supply passage.
16. The vane of claim 12, wherein said cooling air channel is located on a radial plane passing through said seal slot at said one of said axial edges.
17. A vane in a gas turbine engine, said vane comprising:
- an airfoil adapted to extend radially through an annular hot gas path extending in a generally axial direction through said turbine engine, said airfoil including a pressure side and a suction side, an upstream leading edge and a downstream trailing edge;
- a shroud having an outer surface supporting an end of said airfoil, said shroud defining a portion of said annular gas path through said gas turbine engine and including an upstream edge and a downstream edge, and opposing axial edges extending between said upstream and downstream edges;
- each of said axial edges including a generally axially extending seal slot adapted to receive a seal member extending between said shroud and an adjacent shroud;
- a cooling air channel extending generally axially substantially parallel to at least one of said axial edges between said upstream and said downstream edges;
- a cooling air supply passage extending from a cooling air chamber at an inner surface of said shroud to said cooling air channel;
- a plurality of impingement passages extending from said cooling air channel to said one of said axial edges, said impingement passages having exit openings providing a flow of cooling air impinging on an axial edge of said adjacent shroud, said exit openings being located at an upstream end of said cooling air channel and adjacent said upstream edge of said shroud; and
- said cooling air channel being located radially between said outer surface of said shroud and said seal slot at said one of said axial edges for effecting convective cooling of a corner defined at an intersection of said outer surface and said one of said axial edges.
18. The vane of claim 17, wherein said upstream end of said cooling air channel receives cooling air from a leading edge cooling air chamber at said inner surface of said shroud.
19. The vane of claim 18, wherein said cooling air supply passage comprises a replenishing cooling air supply passage that supplies cooling air from a mid-shroud impingement cavity located at an axial midpoint between said upstream edge and said downstream edge at said inner surface of said shroud, said replenishing cooling air supply passage providing replenishing cooling air to said cooling air channel.
20. The vane of claim 17, wherein said cooling air channel is located on a radial plane passing through said seal slot at said one of said axial edges.
Type: Application
Filed: Jan 6, 2011
Publication Date: Jul 12, 2012
Inventors: Gm Salam Azad (Oviedo, FL), Ching-Pang Lee (Cincinnati, OH), Zhihong Gao (Orlando, FL)
Application Number: 12/985,571
International Classification: F01D 5/18 (20060101);