Abstract:
In an axial flow gas turbine ( 30 ), a substantial reduction of the consumption of cooling air can be achieved by providing, within a turbine stage (TS), structure ( 39 - 44 ) to reuse the cooling air that has already been used to cool, especially the airfoils of, the vanes ( 33 ) of the turbine stage (TS), for cooling the stator heat shields ( 38 ) of that turbine stage (TS) downstream of the vanes ( 33 ).

Description:
This application claims priority under 35 U.S.C. §119 to Russian Federation application no. 2010148728, filed 29 Nov. 2010, the entirety of which is incorporated by reference herein. 
     BACKGROUND 
     1. Field of Endeavor 
     The present invention relates to gas turbines, and in particular to axial flow gas turbines. 
     2. Brief Description of the Related Art 
     The invention relates to an axial flow gas turbine, an example of which is shown in  FIG. 5 . The gas turbine  10  of  FIG. 5  operates according to the principle of sequential combustion. It includes a compressor  1 , a first combustion chamber  4  with a plurality of burners  3  and a first fuel supply  2 , a high-pressure turbine  5 , a second combustion chamber  7  with the second fuel supply  6 , and a low-pressure turbine  8  with alternating rows of vanes  13  or  33  and blades  16  or  36 , which are arranged in a plurality of turbine stages arranged along the machine axis  9 . 
     The gas turbine  10  according to  FIG. 5  includes a stator and a rotor. The stator includes a housing with the vanes  13 ,  33  mounted therein; these vanes  13 ,  33  are necessary to form profiled channels where hot gas developed in the combustion chamber  7  flows through. Gas flowing in the required direction hits against the blades  16 ,  36  installed in shaft slits of a rotor shaft and causes the turbine rotor to rotate. To protect the stator housing against the hot gas flowing above the blades  16 ,  36 , stator heat shields installed between adjacent vane rows are used. High temperature turbine stages require cooling air to be supplied into vanes, stator heat shields and blades. 
     A section of a typical cooled gas turbine stage TS of a gas turbine  10  is shown in FIG.  1 . Within a turbine stage TS of the gas turbine  10 , a row of vanes  13  is mounted on a vane carrier  11 . Downstream of the vanes  13  a row of rotating blades  16  is provided, each of which has an outer platform  17  at its tip. Opposite to the tips of the blades  16 , stator heat shields  18  are mounted on the vane carrier  11 . Each of the vanes  13  has an outer platform  14 . The vanes  13  and blades  16  with their respective outer platforms  14  and  17  border a hot gas path  12 , through which the hot gases from the combustion chamber flow. 
     To ensure operation of such a high temperature gas turbine  10  with long-term life span, all parts forming its flow path  12  should be cooled effectively. Therefore, cooling air  23  is directed through respective cooling bores  21  and  22  from a plenum  20  to the stator heat shields  18  and vanes  13  and hot outer platforms  17  of the blades  16 . However, the known turbine design of  FIG. 1  requires sufficient additional amount of cooling air  23  to be supplied into a cavity  19  on the back of the stator heat shields  18  to cool those stator heat shields and the outer blade platform  17 , and this feature can be considered as a shortcoming of this design. Another drawback is the traditional way of stator heat shield fixation, where a gap exists between a vane  13  and the stator heat shield  18  (see the encircled zone A in  FIG. 1 ), and a portion of cooling air leaks from the cavity  19  through that gap into the turbine flow path  12  (see arrows in the zone A). 
     SUMMARY 
     One of numerous aspects of the present invention includes a gas turbine with a turbine stage cooling scheme, which can avoid drawbacks of the known cooling configuration and substantially reduce the consumption of cooling air within the turbine stage. 
     Another aspect includes an axial flow gas turbine that comprises a rotor with alternating rows of air-cooled blades and air-cooled rotor heat shields, and a stator with alternating rows of air-cooled vanes and air-cooled stator heat shields mounted on a vane carrier, whereby the stator coaxially surrounds the rotor to define a hot gas path in between, such that the rows of blades and stator heat shields, and the rows of vanes and rotor heat shields are correlated with each other, respectively, and a row of vanes and the next row of blades in the downstream direction define a turbine stage. Within a turbine stage, means are provided to reuse the cooling air that has already been used to cool, especially the airfoils of, the vanes of the turbine stage, for cooling the stator heat shields of that turbine stage downstream of the vanes. 
     According to an embodiment, the means for reusing comprises first means for collecting the used cooling air when exiting the vanes, and second means for directing the collected used cooling air onto the stator heat shields of said turbine stage downstream of the vanes, for cooling. 
     Preferably, the means for reusing further comprises third means for directing the collected used cooling air onto outer platforms of the blades of said turbine stage downstream of the vanes, for cooling. 
     According to another embodiment, the vanes of the turbine stage each comprise an outer platform, and the means for reusing are integrated into the vanes just above the outer platforms. 
     According to another embodiment, the collecting means comprises a first cavity for each of the vanes located at the exit of the vane cooling air on the upper side of the outer platform, the directing means comprises a second cavity extending in the circumferential direction and being connected to said first cavity, whereby a plurality of first, axially oriented holes, which are equally distributed along the circumferential direction, direct used cooling air from the second cavity onto the outside of the adjacent stator heat shields of the turbine stage, for cooling. 
     According to another embodiment, a plurality of second axially oriented holes, which are equally distributed along the circumferential direction, direct used cooling air from the second cavity onto the outside of the outer platforms of the adjacent blades of the turbine stage, for cooling. 
     Preferably, the outer platforms of the blades of the turbine stage each comprise a circumferentially oriented forward tooth, the vanes of the turbine stage overlap said forward tooth with a circumferentially extending downstream projection at the rear wall of their outer platform, and each downstream projection is provided with a honeycomb just opposite to the forward tooth. 
     According to another embodiment, the first cavity is established by a rib in the form of a frame on the upper side of the outer platform, which frame is covered by a sealing screen. 
     According to another embodiment, the second cavity is established by a recess in the rear wall of the outer platform, which recess is covered by a sealing screen. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The present invention is now to be explained more closely by means of different embodiments and with reference to the attached drawings. 
         FIG. 1  shows cooling details of a turbine stage of a gas turbine according to the prior art; 
         FIG. 2  shows cooling details of a turbine stage of a gas turbine according to an embodiment of the invention; 
         FIG. 3  shows in a perspective view the configuration of the outer platform of the vane of  FIG. 2  in accordance with an embodiment of the invention, whereby all of the screens are removed; 
         FIG. 4  shows in a perspective view the configuration of the outer platform of the vane of  FIG. 3  with all of the screens put in place; and 
         FIG. 5  shows a well-known basic design of a gas turbine with sequential combustion, which may be used as a starting point for implementing embodiments of the invention. 
     
    
    
     DETAILED DESCRIPTION OF EXEMPLARY EMBODIMENTS 
       FIG. 2  presents an exemplary embodiment of a high temperature turbine stage, where cooling air is partly saved due to utilization of air used up in the vanes of the turbine stage. The gas turbine  30  of  FIG. 2  includes a turbine stage TS with a row of vanes  33  followed by a row of blades  36 . The blades  36  are mounted on a rotor, not shown in the Figure. The vanes  33  are mounted on a vane carrier  31 , which surrounds the rotor to define a hot gas path  32 . Also mounted on the vane carrier  31  are stator heat shields  38 , in opposition to outer platforms  37  at the tips of the blades  36 . The outer platforms  37  are provided on their outer side with several teeth, each extending in the circumferential direction. One of these teeth, the forward tooth, has the reference numeral  50 . 
     Air used up in the vane  33  passes from the vane airfoil through the outer platform  34  into a small cavity  39  partitioned off from the basic (outer) platform  34  with a rib  40  (see  FIGS. 2 and 3 ). The air then flows from the cavity  39  into a neighbouring cavity  41 , which extends along the circumferential direction, and is distributed into two parallel rows of first and second holes  42  and  43  equally spaced in the circumferential direction (see  FIGS. 2 and 3 ). First holes  42  direct jets of used cooling air onto the other side of rotor heat shields  38 . Second holes  43  direct jets of used cooling air  1  to the forward teeth  50  of the outer blade platforms  37 . The cavities  39  and  41  are closed with a common sealing screen  44  ( FIG. 4 ). Another (perforated) screen  45  is situated above the remaining largest part of the outer platform  34 , and air for cooling the platform surface and for passing into the interior of the vane airfoil passes through holes of this screen. 
     The efficient utilization of used-up air described above makes it possible to avoid an additional supply of fresh cooling air to the stator heat shields  38  and blade shrouds or outer platforms  37 . 
     Another innovation of the design according to  FIG. 2  is the provision of a projection  47  on the rear wall of the outer vane platform  34  (see  FIGS. 2-4 ). This projection  47  is equipped on its lower side with a honeycomb  51 . The forward tooth  50  of the outer blade platform  37  is situated under the projection  47 , and this tooth  50  prevents additional leakages of used-up air from the cavity  46  between outer platform  37  and stator heat shield  38  into the turbine flow path  32 . 
     When the proposed shape of the outer vane platform  34  according to  FIG. 2  is compared with that of outer vane platform  14  presented in  FIG. 1 , it is clear that leakage minimization is also a result of the absence of an additional gap (see zone A marked in  FIG. 1 ). Thus, used-up air passes without losses through the first holes  42  into the cavity  46  between a stator heat shield  38  and an outer blade platform  37 . This air substantially improves the thermal state of the outer blade platforms  37  and makes it possible to avoid additional air supply for cooling the stator heat shields  38 . 
     Used-up air passes also into a cavity  52  between the vane carrier  31  and stator heat shields  38  through gaps in part joints. Used-up air passing through the second holes  43  serves to protect the forward teeth  50  of the outer blade platforms  37 . 
     With the foregoing, the following advantages can be achieved: 
     1. Air used up in a vane is then utilized to cool other parts. 
     2. There is no need to introduce additional air for cooling the stator heat shields. 
     3. The proposed shape of the outer vane platform with an additional projection  47  on its rear wall makes it possible to avoid additional cooling air leakages through the slit marked by zone A in  FIG. 1 . 
     4. Utilized air fills the cavity  52  (see  FIG. 2 ) and protects the vane carrier  31  against overheating. 
     Thus, a combination of the vane with projection  47  at its outer platform  34  and a separate collector (cavity  39 ) for utilized air, as well as a combination of a non-cooled stator heat shield  38  and a three-pronged outer blade platform  37  with the cavity  46  formed in between, enables a modern high-performance turbine to be created. 
     LIST OF REFERENCE NUMERALS 
     
         
         
           
               1  compressor 
               2 , 6  fuel supply 
               3  burner 
               4 , 7  combustion chamber 
               5  high-pressure turbine 
               8  low-pressure turbine 
               9  axis 
               10 , 30  gas turbine 
               11 , 31  vane carrier 
               12 , 32  hot gas path 
               13 , 33  vane 
               14 , 34  outer platform (vane) 
               15 , 35  cavity 
               16 , 36  blade 
               17 , 37  outer platform (blade) 
               18 , 38  stator heat shield 
               19  cavity 
               20  plenum 
               21 , 22  cooling bore 
               23  cooling air 
               39 , 41 , 46 , 52  cavity 
               40  rib 
               42  hole 
               43  hole 
               44  sealing screen 
               45  screen 
               47  projection 
               48 , 49  hook 
               50  forward tooth (blade outer platform) 
               51  honeycomb 
             TS turbine stage 
           
         
       
    
     While the invention has been described in detail with reference to exemplary embodiments thereof, it will be apparent to one skilled in the art that various changes can be made, and equivalents employed, without departing from the scope of the invention. The foregoing description of the preferred embodiments of the invention has been presented for purposes of illustration and description. It is not intended to be exhaustive or to limit the invention to the precise form disclosed, and modifications and variations are possible in light of the above teachings or may be acquired from practice of the invention. The embodiments were chosen and described in order to explain the principles of the invention and its practical application to enable one skilled in the art to utilize the invention in various embodiments as are suited to the particular use contemplated. It is intended that the scope of the invention be defined by the claims appended hereto, and their equivalents. The entirety of each of the aforementioned documents is incorporated by reference herein.