Patent Publication Number: US-8979482-B2

Title: Gas turbine of the axial flow type

Description:
This application claims priority under 35 U.S.C. §119 to Russian Federation application no. No. 2010148727, filed 29 Nov. 2010, the entirety of which is incorporated by reference herein. 
     BACKGROUND 
     1. Field of Endeavor 
     The present invention relates to the technology of gas turbines, and more specifically to a gas turbine of the axial flow type. 
     More specifically, the invention relates to designing a stage of an axial flow turbine for a gas turbine unit. Generally the turbine stator includes a vane carrier with slots where a row of vanes and a row of stator heat shields are installed one after another. The same stage includes a rotor having a rotating shaft with slots where a row of rotor heat shields and a row of blades are installed one after another. 
     2. Brief Description of the Related Art 
     This disclosure relates to a gas turbine of the axial flow type, an example of which is shown in  FIG. 1 . The gas turbine  10  of  FIG. 1  operates according to the principle of sequential combustion. It includes a compressor  11 , a first combustion chamber  14  with a plurality of burners  13  and a first fuel supply  12 , a high-pressure turbine  15 , a second combustion chamber  17  with a second fuel supply  16 , and a low-pressure turbine  18  with alternating rows of blades  20  and vanes  21 , which are arranged in a plurality of turbine stages arranged along the machine axis  22 . 
     The gas turbine  10  according to  FIG. 1  has a stator and a rotor. The stator includes a vane carrier  19  with the vanes  21  mounted therein; these vanes  21  are necessary to form profiled channels where hot gas developed in the combustion chamber  17  flows through. Gas flowing through the hot gas path  29  in the required direction hits against the blades  20  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  20 , 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 air-cooled gas turbine stage TS of a gas turbine  10  is shown in  FIG. 2 . Within a turbine stage TS of the gas turbine  10 , a row of vanes  21  is mounted on the vane carrier  19 . Downstream of the vanes  21  a row of rotating blades  20  is provided each of which has at its tip an outer platform  24  with teeth ( 52  in  FIG. 3(B) ) arranged on the upper side. Opposite to the tips (and teeth  52 ) of the blades  20 , stator heat shields  26  are mounted on the vane carrier  19 . Each of the vanes  21  has an outer vane platform  25 . The vanes  21  and blades  20  with their respective outer platforms  25  and  24  border a hot gas path  29 , 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  29  should be cooled effectively. Cooling of turbine parts is realized using air fed from the compressor  11  of the gas turbine unit. To cool the vanes  21 , compressed air is supplied from a plenum  23  through the holes  27  into the cavity  28  located between the vane carrier  19  and outer vane platforms  25 . Then the cooling air passes through the vane airfoil and flows out of the airfoil into the turbine flow path  29  (see horizontal arrows at the trailing edge of the airfoil in  FIG. 2 ). The blades  20  are cooled using air which passes through the blade shank and airfoil in vertical (radial) direction, and is discharged into the turbine flow path  29  through a blade airfoil slit and through an opening between the teeth  52  of the outer blade platform  24 . Cooling of the stator heat shields  26  is not specified in the design presented in  FIG. 2  because the stator heat shields  26  are considered to be protected against a detrimental effect of the main hot gas flow by the outer blade platform  24 . 
     Disadvantages of the above described design can be considered to include, firstly, the fact that cooling air passing through the blade airfoil does not provide cooling efficient enough for the outer blade platform  24  and thus its long-term life span. The opposite stator heat shield  26  is also protected insufficiently against the hot gas from the hot gas path  29 . 
     Secondly, a disadvantage of this design is the existence of a slit within the zone A in  FIG. 2 , since cooling air leakage occurs at the joint between the vane  21  and the subsequent stator heat shield  26 , resulting in a loss of cooling air, which enters into the turbine flow path  29 . 
     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 combines a reduction in cooling air mass flow and leakage with an improved cooling and effective thermal protection of critical parts within the turbine stages of the turbine. 
     Another aspect includes a rotor with alternating rows of air-cooled blades and rotor heat shields, and a stator with alternating rows of air-cooled vanes and 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 opposite to each other, respectively, and a row of vanes and the next row of blades in the downstream direction define a turbine stage, and whereby the blades are provided with outer blade platforms at their tips. Means are provided within a turbine stage to direct cooling air that has already been used to cool, especially the airfoils of, the vanes of the turbine stage, into a first cavity located between the outer blade platforms and the opposed stator heat shields for protecting the stator heat shields against the hot gas and for cooling the outer blade platforms. 
     According to an exemplary embodiment, the outer blade platforms are provided on their outer side with parallel teeth extending in the circumferential direction, and said first cavity is bordered by said parallel teeth. 
     According to another embodiment, the vanes each comprise an outer vane platform, the directing means comprises a second cavity for collecting the cooling air, which exits the vane airfoil, and the directing means further comprises means for discharging the collected cooling air radially into said first cavity. 
     Preferably, the discharging means comprises a projection at the rear wall of the outer vane platform, which overlaps the first teeth in the flow direction of the adjacent outer blade platforms, and a screen, which covers the projection such that a channel for the cooling air is established between the projection and the screen, which ends in a radial slot just above the first cavity. 
     According to another embodiment, the second cavity and the discharging means are connected by a plurality of holes, which pass the rear wall of the outer vane platform and are equally spaced in the circumferential direction. 
     According to another embodiment, the second cavity is separated from the rest of the outer vane platform by a shoulder, and the second cavity is closed 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 a well-known basic design of a gas turbine with sequential combustion, which may be used with embodiments in accordance with the invention; 
         FIG. 2  shows cooling details of a turbine stage of a gas turbine according to the prior art; 
         FIG. 3  shows cooling details of a turbine stage of a gas turbine according to an embodiment of the invention; 
         FIG. 4  shows, in a perspective view, the configuration of the outer platform of the vane of  FIG. 3  in accordance with an embodiment of the invention, whereby all of the screens are removed; and 
         FIG. 5  shows in a perspective view the configuration of the outer platform of the vane of  FIG. 3  with all screens put in place. 
     
    
    
     DETAILED DESCRIPTION OF EXEMPLARY EMBODIMENTS 
       FIG. 3  shows cooling details of a turbine stage of a gas turbine  30  according to an exemplary embodiment and demonstrates the proposed design of the turbine stages TS, where cooling air is saved due to utilization of air used up in the vanes  31 . A novelty of this includes not only cooling air savings, but also effective protection of the outer blade platform  34  against hot gas from the hot gas path  39 , due to a continuous sheet of cooling air discharged vertically from the slit ( 50  in  FIG. 3(B) ) into a cavity  41  between parallel teeth  52  on the upper side of the outer blade platforms  34  of the blades  32  with an a turbine stage TS. The slit  50  is formed by a screen  43  covering a projection  44  at the rear wall of the outer vane platform  35  (see  FIG. 3 , zone B, and  FIG. 3(B) ). 
     In general, cooling air from the plenum  33  flows into cavity  38  through the cooling air hole  37 , passes a perforated screen  49  and enters the cooling channels in the interior of the vane airfoil. The cooling air used up in the vane  31  for cooling passes from the airfoil into a cavity  46  partitioned off from the basic outer vane platform  35  by a shoulder  48  (see also  FIG. 4 ). Then, this air is distributed from the cavity  46  into a row of holes  45  equally spaced in the circumferential direction. The cavity  46  is closed with sealing screen  47  (see also  FIG. 5 ). As already mentioned above, perforated screen  49  (see  FIG. 5 ) is situated above the remaining largest portion of the outer vane platform  35 , and air is supplied through the holes in this screen to cool the platform surface and to enter the internal vane airfoil cavity (not shown in the figures). 
     Another new feature of the design is also the provision of the projection  44  on the rear wall of the vane outer platform  35  equipped with a honeycomb  51  on the underneath (see  FIGS. 3-5 ). The forward one of the teeth  52  of the outer blade platform  34 , which prevents additional leakages of used-up air from the cavity  41  into the turbine flow path  39 , is situated directly under the projection  44 . Due to the presence of this projection, an additional gap (see  FIG. 2 , zone A) making way for cooling air leakages, is avoided. 
     Thus, efficient utilization of used-up cooling air makes it possible to avoid supply of additional cooling air to the stator heat shields  36  and to blade shrouds or outer blade platforms  34  because used-up air closes the cavity  41  effectively. 
     In summary, the proposed cooling scheme can have the following advantages: 
     1. Air used up in a vane  31  is utilized to cool parts, especially outer blade platforms  34 . 
     2. There is no need in additional air for cooling the stator heat shields  36 . 
     3. A projection  44 , which is covered by a screen  43 , generates a continuous air sheet of cooling air, which, in combination with the forward tooth  52  of the outer blade platform  34 , closes the cavity  41  located between the teeth  52  on the outer side of the outer blade platforms  34 . 
     4. The shape of the projection  44  on the outer vane platform  35  makes it possible to avoid additional cooling air leakages within the jointing zone (see A in  FIG. 2 ) between the vanes  31  and the stator heat shields  36 . 
     5. Used-up air penetrates through gaps between adjacent stator heat shields  36  into a backside cavity  42  (see  FIG. 3 ) and prevents stator parts from being overheated. 
     Thus, a combination of vanes  31  with the projection  44  and a separate collector  46  to  48  for utilized air, as well as combination of non-cooled stator heat shields  36  and two-pronged outer blade platforms  34  with a cavity  41  formed between the outer teeth  52  of these outer blade platforms  34 , enables a modern high-performance turbine to be designed. 
     LIST OF REFERENCE NUMERALS 
     
         
         
           
               10 , 30  gas turbine 
               11  compressor 
               12 , 16  fuel supply 
               13  burner 
               14 , 17  combustion chamber 
               15  high-pressure turbine 
               18  low-pressure turbine 
               19 , 40  vane carrier (stator) 
               20 , 32  blade 
               21 , 31  vane 
               22  machine axis 
               23 , 33  plenum 
               24 , 34  outer blade platform 
               25 , 35  outer vane platform 
               26 , 36  stator heat shield 
               27 , 37  hole 
               28 , 38  cavity 
               29 , 39  hot gas path 
               41 , 42 , 46  cavity 
               43 , 47 , 49  screen 
               44  projection 
               45  hole 
               48  shoulder 
               50  slit 
               51  honeycomb 
               52  tooth (outer blade platform) 
             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.