Patent Document

The present invention relates generally to cooling turbine engine components and more specifically, to cooling stator shrouds, or other stator components having a similar geometry, and associated seals within the hot gas path of a gas turbine, downstream of the turbine combustor(s). 
     BACKGROUND OF THE INVENTION 
     In general, gas turbines combust a mixture of compressed air and fuel to produce hot combustion gases. The combustion gases may flow through one or more turbine sections to generate power to drive, for example, an electrical generator and/or a compressor. Within the gas turbine sections, the combustion gases typically flow through one or more stages of nozzles and blades (or buckets). The turbine nozzles may include circumferential rings of stationary vanes that direct the combustion gases to the rotating blades or buckets attached to the turbine rotor. As the combustion gases flow past the buckets, the combustion gases drive the buckets to rotate the rotor, which, in turn, drives the generator or other device. The hot combustion gases are contained using seals between circumferentially-adjacent arcuate segments of stationary shrouds surrounding the nozzle vanes and/or buckets; between the platforms of circumferentially-adjacent rotating buckets or bucket segments on a rotor wheel; and seals between axially adjacent nozzle and bucket shrouds of the same or successive turbine stages. 
     The seals are designed to prevent or minimize ingestion of higher-pressure compressor discharge or extraction flows into the lower-pressure hot gas path. Nevertheless, leakage about the seals is inevitable and results in reduced compressor performance which contributes to an overall reduction in the efficiency of the turbine. 
     At the same time, the hot gas path components, including the shroud segments and seals must be cooled to withstand the extremely high combustion gas temperatures. Conventional cooling schemes usually involve some combination of internal cooling features and associated cooling technique (for example, impingment, serpentine, pin-fin bank, near-wall cooling) where the cooling air is eventually exhausted through film-cooling holes that enable additional cooling of the surface of the component. In some instances, however, it is not desirable to exhaust all or part of the internal cooling flow in this manner. 
     While various techniques have been employed to cool the shrouds and seals between adjacent shroud and other similar stator component segments, it remains desirable to provide enhanced cooling for the shrouds and seals, and to use the heated or spent cooling air for at least one other purpose, for example, to purge the segment gap, i.e., diluting the hot combustion gases below (i.e., radially inward of) the seal, thus cooling the seal while also preventing or minimizing compressor extraction flows from leaking into the hot gas path. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one exemplary but non limiting embodiment, there is provided a segment for a ring-shaped rotary machine stator component comprising a segment body having an end face formed with a circumferentially-facing seal slot adapted to receive a seal extending between the segment body and a corresponding seal slot in an adjacent segment body; a channel provided in the segment body in proximity to the seal slot, supplied with cooling air; and a passage extending from the channel into the seal slot. 
     In another exemplary aspect, there is provided an annular turbine component comprising: plural arcuate segments arranged to form a complete annular ring, each segment having end faces provided with seal slots; a seal extending between seal slots of adjacent segments sealing radially oriented gaps between the segments; a channel provided in each segment in proximity to at least one of said seal slots, and adapted to be supplied with cooling air; and a passage extending from said channel and opening into said at least one seal slot on a radially-outer, high-pressure side of the seal. 
     In still another aspect, there is provided a gas turbine stator comprising first and second axially adjacent, annular shrouds having opposed end faces provided with respective seal slots; wherein a circumferential, axially-extending gap is formed between the opposed end faces; a circumferential seal seated in the respective seal slots to thereby seal the axially-extending gap, the seal, in use, separating relatively higher and lower pressure areas on radially-outer and radially-inner sides thereof, said radially-inner side exposed to a hot gas path; and one or more cooling channels provided within each of the first and second axially-adjacent, annular shrouds adapted to be supplied with cooling air, the one or more cooling channels arranged to introduce cooling air into a respective one of the seal slots or axially-extending gaps in the relatively lower pressure area on the radially-inner side of the seal. 
     The invention will now be described in greater detail in connection with the drawings identified below. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a partial sectional view of a gas turbine engine along an axis of rotation of the engine; 
         FIG. 2  is an enlarged detail of the encircled area indicated by reference numeral  36  in  FIG. 1 ; 
         FIG. 3  is a partial front view of a gas turbine shroud segment in accordance with an exemplary but nonlimiting embodiment; and 
         FIG. 4  is a partial side circumferential view of a gas turbine shroud segment in accordance with the embodiment of  FIG. 3 . 
         FIG. 5  is a partial front view of a gas turbine shroud segment in accordance with another exemplary but nonlimiting embodiment. 
         FIG. 6  is a partial side view of a gas turbine shroud segment in accordance with the embodiment of  FIG. 5 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  is a cross-sectional side view of a conventional gas turbine engine  10  taken along a longitudinal axis  12 , i.e., the axis of rotation of the turbine rotor. With reference also to the enlarged detail in  FIG. 2 , it will be appreciated that air enters the gas turbine engine  10  through the air intake section  14  of a compressor  16 . The compressed air exiting the compressor  16  is directed to the combustors  18  (one shown) to mix with fuel which combusts to generate hot combustion gases. Multiple combustors  18  may be annularly disposed within the turbine combustor section  20 , and each combustor  18  may include a transition piece  22  that directs the hot combustion gases from the combustor  18  to the gas turbine section  24 . In other words, each transition piece  22  defines a hot gas path from its respective combustor  18  to the turbine section  24 . 
     The illustrated, exemplary gas turbine section  24  includes three separate stages  26 . Each stage  26  includes a set or row of buckets  28  coupled to a respective rotor wheel  30  that is rotatably attached to the turbine rotor or shaft represented by the axis of rotation  12 . Between each wheel  30  is a set of nozzles  40  incorporating a circumferential row of stationary vanes or blades  42 . The nozzle vanes  42  are supported between segmented, inner and outer stator shrouds or side walls  44 ,  46 , each segment incorporating one or more vanes, while the buckets  28  are surrounded by stationary, stator shroud segments  48 . The nozzle and bucket shrouds serve to contain the hot combustion gases and allow a motive force to be efficiently applied to the buckets  28 . The hot combustion gases exit the gas turbine section  24  through the exhaust section  34 . 
     Applications for the present invention relate to seals extending across radially-oriented gaps between circumferentially-adjacent nozzle vane and/or bucket shroud segments; between circumferentially-adjacent buckets; and between axially-adjacent shrouds (nozzle and bucket) in the same or adjacent stage. 
     It will be understood, of course, that although the turbine section  24  is illustrated as a three-stage turbine, the cooling and sealing arrangements described herein may be employed in turbines with any number of stages and shafts, e.g., a single stage turbine, a dual turbine that includes a low-pressure turbine section and a high-pressure turbine section, or in a multi-stage turbine section with three or more stages. Furthermore, the cooling and sealing arrangements described herein may be utilized in gas turbines, steam turbines, hydroturbines, etc. 
     Typically, discharge air from the compressor  16  (also known as compressor extraction flow) ( FIG. 1 ), which may act as a cooling fluid, may be directed through the stationary vanes  42 , the inner and outer band segments  44  and  46 , and/or the shroud segments  48  to provide the required cooling of these components. 
     In the exemplary but nonlimiting embodiment described herein, the discharge air from the compressor  16  is also used as a cooling fluid to mitigate or control the buildup of thermal energy on the hot side of the shroud segments  48  facing the buckets  28 . 
     In some embodiments, other cooling fluids may be used in addition to or in lieu of the compressor discharge air, such as steam, recirculated exhaust gas, or fuel. 
       FIGS. 3 and 4  are partial end views of a stator shroud segment  50  (i.e., one arcuate segment of the annular shroud  48 ) in accordance with a first exemplary but nonlimiting embodiment. It will be understood that the shroud segment  50  as viewed in  FIG. 3  includes a radially-inner surface  52  that faces or lies radially adjacent a row of buckets  28  on a turbine wheel as described in connection with  FIG. 2 . A circumferential interface surface  54  (or end face) lies opposite an adjacent shroud segment  56  (shown in phantom), with a radially-extending gap  58  therebetween. A seal slot  60  formed in the interface surface or end face  54  is aligned with a similar slot  62  in the adjacent interface surface  64 , the pair of slots adapted to receive a seal  66  that inhibits radially-inward leakage of higher-pressure compressor extraction flows into the hot combustion gases flowing along the hot gas path  67  ( FIG. 4 ). It will be understood that a similar seal/seal slot arrangement is provided on the opposite interface surface such that the seals extend between adjacent slots of adjacent segments about the entire annular shroud. 
     In the illustrated embodiment, surface  52  (or hot-gas-facing side) may be coated with a known thermal barrier coating (TBC)  68  to provide some protection for the surface  54  which is directly exposed to the hot combustion gases. 
     A channel  70  is formed in the surface  52 , extending in an axial direction (parallel to the hot gas path) in the exemplary embodiment. The channel  70  could also extend in a circumferential direction and could also have a wavy, zig-zag or other suitable shape. The channel  70 , which may be of any desired length, is supplied with cooling air, e.g., compressor extraction air, by means of a passage  72  extending angularly from a radially-outer surface  74  of the shroud segment  50  and opening into the channel  70  at one end thereof. Thus, the passage  72  maybe regarded as an inlet passage. In an exemplary embodiment shown in  FIG. 3 , an outlet passage  76  is formed in the shroud segment, extending radially outward from an opposite end of the channel  70 , and into the seal slot  60 . In this way, cooling air passing through the channel  70  absorbs heat, and thus cools the surface  52  (and TBC  68 ), and the heated cooling air is then exhausted to the seal slot  60  where it cools the underside or low-pressure side of the seal, and then enters and purges the part of the gap  58  which lies radially inward of the seal  66 , i.e., the spent cooling air mixes with and dilutes the hot gas in the segment gap that would otherwise make the seal and segment end faces too hot. The flow of air into that part of the gap radial inward of the seal  66  also inhibits leakage of higher-pressure compressor air into the hot gas path. It will be understood that different seal configurations will dictate the exact flow of the heated cooling air upon reaching the seal slot  60 . It will also be understood that a similar cooling arrangement is provided in the adjacent shroud segment  56 . 
     In another exemplary shown in  FIGS. 5 and 6 , the shroud segment  150  includes a radially inner surface  152 , a circumferential interface surface  154  that faces an adjacent shroud segment (similar to shroud segment  56 ) with a radially-extending gap  158  therebetween. Seal slot  160  is similar to seal slot  60  and cooperates with an adjacent seal slot (similar to slot  62 ). The radially-inner surface  152  may also be coated with a TBC  168 . As in the previously-described embodiment, an inlet passage  172  extends from a radially-outer surface  174  of the shroud segment and opens into a channel  170 . In this embodiment, however, the outlet passage  176  from the channel  170  opens on the end face or surface  154  radially inwardly of the seal slot  160 , so as to purge that portion of the gap  158  radially inward of the seal. By having the outlet from passage  176  sufficiently distanced (in the radially outward direction) from the hot gas path, the purge air will be more effective in diluting hot gas in the gap. If the outlet from passage  176  is too close to the hot gas path, the purge air would be immediately sucked into the hot gas path, and additional flow would be required to purge the gap. 
     In both embodiments, the air otherwise needed to purge the gaps between shroud segments is reduced by the configurations disclosed herein where spent cooling air is exhausted into the gaps radially inward of the seals. 
     It will also be understood that the TBC coating  68  or  168  may be applied over a plate or other substrate covering the radially-inward side of the channel  70 ,  170 , or the coating itself may close the open side of the microchannel. 
     With respect to channels  70 ,  170 , various dimensional relationships and geometries are possible. For example, in accordance with certain embodiments, the channels  70  and  170  may be provided as microchannels having widths and depths between approximately 50 microns and 4 mm in any suitable combination. While illustrated as square or rectangular in cross-section, the microchannels may be any suitable shape that may be formed using grooving, etching, or similar forming techniques. For example, the microchannels may have circular, semi-circular, curved, triangular or rhomboidal cross-sections in addition to or in lieu of the square or rectangular cross-sections illustrated. In addition, width and depth of the channel(s) may also vary uniformly or differentially throughout its length. Therefore, the disclosed microchannels may have straight or curved geometries consistent with such cross-sections. 
     It will be understood that the cooling/sealing arrangement as described above in connection with the bucket shroud  48  is applicable as well to the segments of the inner and outer nozzle shrouds  44 ,  46 . In addition, the cooling/sealing arrangemnts are also applicable to seals located axially between the nozzle shrouds and the bucket shrouds, for example, between nozzle shroud  46  and bucket shroud  48 . In the case of axially-adjacent shrouds, seal  66  (configured as a circumferential seal) could be considered as sealing an axial gap  58  between a nozzle shroud  50  and an axially-adjacent bucket shroud  56 , recognizing that the opposed edge faces  54 ,  64  may not be as shown in  FIG. 3 . 
     It will also be appreciated that the invention is applicable to any turbine stage although it is believed that stages  1  and  2  would likely benefit from the described arrangements. 
     While various embodiments are described herein, it will be appreciated from the specification that various combinations of elements, variations or improvements therein may be made by those skilled in the art, and are within the scope of the invention.

Technology Category: 2