Patent Publication Number: US-2011058957-A1

Title: Blade for a gas turbine

Description:
CROSS REFERENCE TO RELATED APPLICATION 
     This application is a continuation of International Application No. PCT/EP2009/053116 filed Mar. 17, 2009, which claims priority to Swiss Patent Application No. 00470/08, filed Mar. 31, 2008, the entire contents of all of which are incorporated by reference as if fully set forth. 
    
    
     FIELD OF INVENTION 
     The present invention relates to the field of gas turbine technology, in particular to a blade for a gas turbine. 
     BACKGROUND 
     Gas turbines with sequential combustion are known and have been proved to be successful in industrial operation. 
     Such a gas turbine, which has been known among experts as GT24/26, is disclosed for example in an article by Joos, F. et al., “Field Experience of the Sequential Combustion System for the ABB GT24/GT26 Gas Turbine Family”, IGTI/ASME 98-GT-220, 1998 Stockholm.  FIG. 1  there shows the basic construction of such a gas turbine, wherein  FIG. 1  there is reproduced in the present application as  FIG. 1 . Furthermore, such a gas turbine is disclosed in EP-B1-0 620 362. 
       FIG. 1  shows a gas turbine  10  with sequential combustion, in which a compressor  11 , a first combustion chamber  14 , a high-pressure turbine (HPT)  15 , a second combustion chamber  17  and a low-pressure turbine (LPT)  18  are arranged along an axis  19 . The compressor  11  and the two turbines  15 ,  18  are part of a rotor which rotates around the axis  19 . The compressor  11  draws in air and compresses it. The compressed air flows into a plenum and from there into premix burners where this air is mixed with at least one fuel, at least with fuel which is introduced via the fuel feed line  12 . Such premix burners are disclosed in principle in EP-A1-0 321 809 or EP-A2-0 704 657. 
     The compressed air flows into the premix burners, where the mixing with at least one fuel takes place, as explained above. This fuel/air mixture then flows into the first combustion chamber  14 , into which this mixture is combusted, forming a stable flame front. The hot gas which is thus made available is partially expanded in the adjoining high-pressure turbine  15 , performing work, and then flows into the second combustion chamber  17  where a further feed  16  of fuel takes place. As a result of the high temperatures which the hot gas, which is partially expanded in the high-pressure turbine  15 , always has, a combustion, which is based on self-ignition, takes place in the second combustion chamber  17 . The hot gas which is reheated in the second combustion chamber  17  is then expanded in a multistage low-pressure turbine  18 . 
     The low-pressure turbine  18  comprises a plurality of rows, arranged in series in the flow direction, of rotor blades and stator blades, which are arranged in alternating sequence. For example, the stator blades of the third stator blade row in the flow direction are provided with the designation  20 ′ in  FIG. 1 . 
     The stator blades in their interior are provided with a cooling passage which is guided back and forth mostly in a serpentine manner between the ends of the blade airfoil and through which flows a cooling medium, mostly cooling air. This also applies to all the thermally highly loaded rotor blades. 
     For producing such a blade, a casting process, in which a casting core is used for forming the cooling passage, is predominantly used. For production engineering reasons, the casting core projects from the blade at one or both ends and after completion of the casting process correspondingly leaves behind one or more core outlet openings which later have to be sealed off. A method for sealing off such openings is described for example in printed publication U.S. Pat. No. 6,837,417B2. With this method, the opening in the blade is sealed off by a sintered cap which neither on the inner side nor on the outer side aligns with the respective wall surface in a flush manner. This leads to uneven, stepped surfaces which impede the flow of the medium which is used for cooling and so impair the effectiveness of the cooling, even partially cancelling it out. 
     SUMMARY 
     The present disclosure is directed to a blade for a gas turbine. The blade is produced in accordance with a casting process and includes a blade airfoil which extends in a radial direction between a blade tip and a shroud and in an interior of which extends a cooling passage, which bypasses the shroud and blade tip. A cooling medium flows through the cooling passage for cooling the blade. In end-face ends of the blade there are core outlet openings which arise from the use of a casting core and which connect the cooling passage to an outside space and are sealed off by a sealing element. The sealing element is formed and inserted into the core outlet openings so that it aligns with a wall surface of the cooling passage in a flush manner. 
     The present disclosure is also directed to a gas turbine including a blade which is produced in accordance with a casting process and includes a blade airfoil which extends in a radial direction between a blade tip and a shroud and in an interior of which extends a cooling passage, which bypasses the shroud and blade tip. A cooling medium flows through the cooling passage for cooling the blade. In end-face ends of the blade there are core outlet openings which arise from the use of a casting core and which connect the cooling passage to an outside space and are sealed off by a sealing element. The sealing element is formed and inserted into the core outlet openings so that it aligns with a wall surface of the cooling passage in a flush manner, with the blade being arranged in a turbine of the gas turbine. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS  
       The invention shall subsequently be explained in more detail based on exemplary embodiments in conjunction with the drawing. All elements which are not essential for the direct understanding of the invention have been omitted. Like elements are provided with the same designations in the different figures. The flow direction of the media is indicated by arrows. In the drawings: 
         FIG. 1  shows the principle construction of a gas turbine with sequential combustion according to the prior art, 
         FIG. 2  shows a stator blade in a perspective side view, 
         FIG. 3  shows the shroud, with a first core outlet opening, in plan view from the top, 
         FIG. 4  shows the section through the sealed-off core outlet opening in the plane IV-IV of  FIG. 3  according to an exemplary embodiment of the invention, 
         FIG. 5  shows the inner platform, with a second core outlet opening, in plan view from the bottom, and 
         FIG. 6  shows the section through the sealed-off core outlet opening in the plane VI-VI of  FIG. 5  according to another exemplary embodiment of the invention. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS  
     Introduction to the Embodiments 
     It is an object of the invention to create a blade of the type referred to in the introduction which avoids the disadvantages of known blades and which provides an optimized, undisturbed flow of the cooling medium in the blade. 
     The object is achieved by the entirety of the features of the invention. In the invention, the sealing elements are formed and inserted into the core outlet openings so that they align with the wall surface of the cooling passage in a flush manner. As a result of this, negative influencing of the flow of the cooling medium by means of the sealing elements is reliably avoided. 
     In one development, the sealing elements are formed as prefabricated sealing plugs. These can be inserted into the core outlet openings in a simple manner and fixed quickly and reliably there. This takes place preferably by the sealing elements, or sealing plugs, being hard-soldered into the core outlet openings. 
     The sealing element or the sealing plug can be positioned especially simply if abutting surfaces, upon which lie the sealing elements or sealing plugs, are formed in the core outlet openings. 
     According to another development, the sealing elements or sealing plugs are inserted into the core outlet openings so that they align with the outer surfaces of the platforms in a flush manner. As a result of this, fluidic advantages also ensue in the outside space of the blade. 
     The blade according to the invention is advantageously used in a gas turbine. 
     The gas turbine in this case can be a gas turbine with sequential combustion, having a first combustion chamber with a downstream high-pressure turbine, and a second combustion chamber with a downstream low-pressure turbine, wherein the blades are arranged both in the low-pressure turbine and in the high-pressure turbine. In particular, the low-pressure turbine in such a gas turbine has a plurality of rows of stator blades and rotor blades in series in the flow direction. 
     DETAILED DESCRIPTION  
     In  FIG. 2 , in a perspective side view, is a stator blade which can be used for example in the low-pressure turbine of a gas turbine with sequential combustion according to  FIG. 1 , and which is suitable for realization of the invention. The use of the subject according to the invention, however, is limited neither to said gas turbine type nor to a special stator blade or rotor blade nor to a specific blade row. The stator blade  20  which is taken as a basis here comprises a blade airfoil  22  which is curved in space and extends in the longitudinal direction (in the radial direction of the gas turbine) between a blade tip  23  and a shroud  21 , and in the direction of the hot gas flow  30  extends from a leading edge  27  to a trailing edge  28 . Between the two edges  27  and  28 , the blade airfoil  22  is delimited on the outside by a pressure side  31  (facing the viewer in  FIG. 2 ) and an (opposite) suction side. 
     The stator blade  20  is fastened on the turbine casing by hook-like fastening elements  24  and  25  which are formed on the upper side of the shroud  21 , while blade tip  23  butts against the rotor with sealing effect. 
     In the interior of the blade airfoil  22 , provision is made for a cooling passage ( 39  in  FIGS. 4 ,  6 ), which extends back and forth in a serpentine manner between the platforms  21 ,  23 , for cooling the blade  20 , as is shown for example in printed publication WO-A1-2006029983. For producing such a cooling passage by a casting technique there is a requirement for a core which in the present example leaves behind in the platforms  21  and  23  the core outlet openings  40  in the shroud ( FIG. 3 ,  4 ) or  41  in the blade tip ( FIGS. 5 ,  6 ). 
     The core outlet openings  40 ,  41  are formed and sealed off by corresponding sealing plugs  32  or  36  according to  FIG. 4  and  FIG. 6  so that the outer surfaces of the sealing plugs  32 ,  26  align with the wall surfaces of the surroundings in a flush manner at least where the wall surfaces are impinged upon by the flowing cooling medium. This is particularly the case in the cooling passage  39  through which the cooling medium is guided in the interior of the blade. 
     In the case of the round core outlet opening  40 , which is provided in the shroud  21 , an annular abutment surface  33  is created in the core outlet opening by a diameter step, the sealing plug  32  being seated on this by a corresponding shoulder ( FIG. 4 ). The sealing plug  32  in this case is dimensioned and formed so that after its insertion into the core outlet opening  40  both the outer surface of the shroud  21  and the surface  35  of the inner wall of the cooling passage  39  are continuous. The sealing plug  32  is fixed in the core outlet opening  40  preferably by means of a hard-soldered connection  34 . 
     A similar procedure is applied in the case of the four-sided core outlet opening  41  in the blade tip  23 . In the core outlet opening  41 , provision is made on opposite sides, at a specified depth, for abutting surfaces  37  on which is seated the sealing plug  36  which is inserted into the core outlet opening  41  and adapted in the edge contour ( FIG. 6 ). Also in this case, the sealing plug  36  is fixed in the core outlet opening  41  by means of hard-soldered connections  38  and aligns with the surrounding surface in a flush manner. 
     By means of the invention, which in principle can be used in all cooled blades of turbines, the disturbing influence of the sealing elements upon the flow of the cooling medium is minimized. As a result, the walls of the blade are optimally cooled, which leads to an extension of the blade service life. A preferred use of the blade according to the invention is to be encountered in large stationary gas turbines, for example in gas turbines with sequential combustion, which have been known among experts under the designation GT24/26. In the case of the last-named gas turbines, the preferred use of such a blade can be in the low-pressure turbine. Such a blade can also be used in other gas turbine types. 
     LIST OF DESIGNATIONS 
     
         
           10  Gas turbine 
           11  Compressor 
           12 ,  16  Fuel feed line 
           13  EV burner 
           14 ,  17  Combustion chamber 
           15  High-pressure turbine 
           18  Low-pressure turbine 
           19  Axis 
           20 ,  20 ′ Blade 
           21  Shroud 
           22  Blade airfoil 
           23  Blade tip 
           24 ,  25  Fastening element (hook-like) 
           27  Leading edge 
           28  Trailing edge 
           29  Throttling element 
           30  Hot gas flow 
           31  Pressure side 
           32 ,  36  Sealing plug 
           33 ,  37  Abutment surface 
           34 ,  38  Hard-soldered connection 
           35  Surface (cooling passage) 
           39  Cooling passage 
           40 ,  41  Core outlet opening