Abstract:
In one example, an arcuate segment for a ring-shaped, rotary machine component such as a stator nozzle or bucket shroud, includes 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 to seal a radially-extending gap between the adjacent segment bodies. A cooling channel is provided in the segment body in proximity to the seal slot, and is adapted to be supplied with cooling air. A passage extends from the cooling channel into the seal slot, at a location where the cooling air can be supplied to the higher pressure area on the radially-outer side of the seal.

Description:
[0001]    The present invention relates generally to cooling turbine engine components and, more specifically, to reducing secondary cooling flows in the area of seals between shrouds and shroud segments that are used to prevent ingress of high-pressure compressor air into the hot combustion gas path. 
       BACKGROUND OF THE INVENTION 
       [0002]    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 a load, such as 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. The combustion gases drive the buckets to rotate the rotor, thereby driving the load. 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. 
         [0003]    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. 
         [0004]    At the same time, the hot gas path components, including the shroud segments, buckets 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, or exhausted into the hot gas path. In some instances, however, it is not desirable to exhaust all or part of the internal cooling flow in this manner. 
         [0005]    While various techniques have been employed to cool the shrouds, buckets and associated seals, it remains desirable to provide enhanced cooling for the shrouds, buckets and/or seals, and to use the heated or spent cooling air for at least one other purpose, for example, to further reduce the leakage of compressor air into the hot gas path. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0006]    In one exemplary aspect, the invention provides an arcuate segment for a ring-shaped, rotary machine 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 to seal a radially-extending gap between the adjacent segment bodies, and wherein, in use, the seal separates relatively higher and lower pressure areas in the radially-extending gap, on radially outer and radially inner sides respectively, of the seal; a cooling channel provided in the segment body in proximity to the seal slot, adapted to be supplied with cooling air; and a passage extending from the cooling channel into the seal slot at a location enabling supply of cooling air to the higher pressure area on the radially-outer side of the seal. 
         [0007]    In another aspect, the invention provides 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 the seal slots, and adapted to be supplied with cooling air; and a passage extending from the channel and opening into the at least one seal slot on a radially-outer, high-pressure side of the seal. 
         [0008]    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; 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 in the relatively higher pressure area on the radially-outer side of the seal. 
         [0009]    The invention will now be described in greater detail in connection with the drawings identified below. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0010]      FIG. 1  is a partial sectional view of a gas turbine engine, taken along an axis of rotation of the turbine rotor; 
           [0011]      FIG. 2  is an enlarged detail of the encircled area indicated by reference numeral  36  in  FIG. 1 ; 
           [0012]      FIG. 3  is a simplified section view of turbine stator and rotor shrouds and exemplary locations where seals are used between adjacent segments; 
           [0013]      FIG. 4  is a partial section of a stator shroud segment illustrating a shroud internal cooling circuit adjacent a shroud segment seal cavity and seal in accordance with a first exemplary but nonlimiting embodiment; and 
           [0014]      FIG. 5  is a partial section illustrating an internal cooling circuit adjacent a seal cavity and seal between adjacent rotor components in accordance with another exemplary but nonlimiting embodiment 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0015]      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 respective 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 . 
         [0016]    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 . 
         [0017]    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. 
         [0018]    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. 
         [0019]    Typically, discharge air from the compressor  16  ( FIG. 1 ), which may act as a cooling fluid, is supplied to internal cooling circuits in the stationary vanes  42 , the inner and outer band segments  44  and  46 , and/or the bucket shroud segments  48  ( FIG. 2 ) to provide the required cooling of these components. In  FIG. 3 , a seal cavity  50  is shown that is adapted to receive a seal extending between axially-adjacent shrouds  54 ,  56 . Seal cavities  52 ,  53  are adapted to receive seals between circumferentially-adjacent segments of the shrouds  54 ,  56 , respectively. These are just two of several locations where cooling flow circuits and seal cavities as described herein may be used to not only perform a cooling function, but also serve an additional function by using the heated or spent cooling air to replace the leakage flow in the higher-pressure areas of the radially-oriented gaps between segments and/or axially-oriented gaps between axially-adjacent shrouds. 
         [0020]    Turning to  FIG. 4 , a nozzle vane shroud segment  58  is formed or provided with a seal slot  60  along an edge face  62 . An adjacent segment  64  has a similar seal slot  66  provided in the opposing edge face  68 . A seal  70  bridging the gap  72  between the edge faces  62 ,  68 , is seated in the respective opposed slots  60 ,  66  and is intended to block the flow of higher-pressure compressor air radially inwardly into the hot gas path.  FIG. 4  also shows a cooling passage or cavity  74  which is located to cool the radially-inner surface  76  of the shroud segment  58 . Note that surface  76 , which is exposed to hot combustion gases, may be coated with a thermal barrier coating (TBC). 
         [0021]    Rather than simply exhausting the spent cooling air into the hot gas path, a further passage  78  is provided to connect the cooling passage or cavity  74  to a plenum or cavity  80  which permits introduction of the spent cooling air at a location upstream of the seal  70  (on the high-pressure side of seal, i.e., above the seal as viewed in  FIG. 4 ). It will be understood that similar plenums or cavities are provided at spaced locations along the length of the seal slot. By periodically relieving the side surface of the seal slot to form the plenums  80 , the seal is retained in place (and prevented from blocking the passage(s)  78 ) while permitting the flow of spent cooling air into the segment gap  72  radially outward of the seal  70 , thus serving to replace the higher-pressure compressor air that would otherwise leak past the seal  70 . While the spent cooling air will eventually leak around the seal  70  and mix with the hot combustion gases, the continued exhausting of spent cooling air into the gap  72  via the passages  78  and plenums  80  outwardly of the seal  70 , reduces secondary compressor flows and thus results in higher turbine efficiency. 
         [0022]    It will be understood that a substantially similar arrangement may be provided between circumferentially-adjacent bucket shroud segments (for example, bucket shroud  48  in  FIG. 2  and bucket shroud  56  in  FIG. 3 ); and between axially adjacent nozzle and bucket shrouds (e.g., between shrouds  46  and  48  in  FIG. 2  or between shrouds  54  and  56  in  FIG. 3 ). In the case of axially-adjacent shrouds, seal  70  (configured as a circumferential seal) could be considered as sealing an axial gap  72  between a nozzle shroud  58  and an axially-adjacent bucket shroud  64 , recognizing that the opposed edge faces  62 ,  68  may not be as shown in  FIG. 3 . 
         [0023]    This same concept can be applied to the seals extending along edges of circumferentially-adjacent buckets in an annular row of buckets mounted on a turbine rotor wheel.  FIG. 5  illustrates an exemplary but nonlimiting embodiment where a damper pin/seal  82  extends in a generally axial direction along the opposite side edges  84 ,  86  of circumferentially-adjacent bucket platforms  88 ,  90 , respectively, seated partially within opposed damper pin slots  92 ,  94 . Secondary compressor flow used to cool the bucket platform  88 , may be exhausted from cooling cavity  96  via passage  98  into a plurality of cavities or plenums  100  which introduce the spent cooling air to the high-pressure side of the seal  82  (the lower side as viewed in  FIG. 5 ). A similar arrangement could be provided in the adjacent bucket platform  90 . The manner in which the spent cooling flow is used to replace compressor leakage flow is substantially as described above in connection with the stationary stator nozzle and bucket shrouds. 
         [0024]    It will be understood that other seal slot configurations may employed with the same result. For example, rather than discrete plenums periodically spaced along the length of the seal slot, the seal slot itself could be formed with an offset step or shoulder along the base of the slot which would provide the required space for receiving the spent cooling air on the high pressure side of the seal while also preventing lateral shifting of the seal within its respective opposed slots. 
         [0025]    In all cases, by re-using spent cooling air as described herein, it is possible to reduce the overall amount of compressor extraction flow required for cooling and/or purging and thereby increase compressor and turbine efficiency. 
         [0026]    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.