BLADE FOR A GAS TURBINE

Abstract

A blade is provided for a gas turbine, especially for the low-pressure turbine of a gas turbine with sequential combustion, and is produced in accordance with a casting process and has a blade airfoil which extends in the radial direction between an inner platform and an outer platform, and in the interior of which extends a cooling passage, bypassing the platforms, and through which flows a cooling medium, especially cooling air, for cooling the blade. In the outer and/or inner platform there are core outlet openings which arise from the use of a casting core and which connect the cooling passage to the outside space and are sealed off by means of a sealing element. Optimum cooling is ensured by the sealing elements being formed and inserted into the core outlet openings so that they align with the wall surface of the cooling passage in a flush manner.

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

Claims

1. A blade (20) for a gas turbine (10), said blade (20) being produced in accordance with a casting process and comprises a blade airfoil (22) which extends in a radial direction between a blade tip (23) and a shroud (21) and in an interior of which extends a cooling passage (39), which bypasses the shroud and blade tip (21, 23), and through which flows a cooling medium for cooling the blade (20), wherein in end-face ends (21 or 23) of the blade (20) there are core outlet openings (40 or 41) which arise from the use of a casting core and which connect the cooling passage (39) to an outside space and are sealed off by a sealing element (32 or 36), the sealing element (32, 36) is formed and inserted into the core outlet openings (40, 41) so that it aligns with a wall surface of the cooling passage (39) in a flush manner.

2. The blade as claimed in claim 1, wherein the sealing elements are formed as prefabricated sealing plugs (32, 36).

3. The blade as claimed in claim 1, wherein the sealing elements (32, 36) are hard-soldered into the core outlet openings (40, 41).

4. The blade as claimed in claim 3, wherein abutment surfaces (33, 37), upon which lie the sealing elements (32, 36), are formed in the core outlet openings (40, 41).

5. The blade as claimed in claim 1, wherein the sealing elements (32, 36) are inserted into the core outlet openings (40, 41) so that they align with outer surfaces of the shroud and blade tip (21, 23) in a flush manner.

6. The blade as claimed in claim 2, wherein the sealing plugs (32, 36) are hard-soldered into the core outlet openings (40, 41).

7. The blade as claimed in claim 6, wherein abutment surfaces (33, 37), upon which lie the sealing plugs (32, 36), are formed in the core outlet openings (40, 41).

8. The blade as claimed in claim 2, wherein the sealing plugs (32, 36) are inserted into the core outlet openings (40, 41) so that they align with outer surfaces of the shroud and blade tip (21, 23) in a flush manner.

9. A gas turbine comprising a blade (20), said blade (20) being produced in accordance with a casting process and has a blade airfoil (22) which extends in a radial direction between a blade tip (23) and a shroud (21) and in an interior of which extends a cooling passage (39), which bypasses the shroud and blade tip (21, 23), and through which flows a cooling medium for cooling the blade (20), wherein in end-face ends (21 or 23) of the blade (20) there are core outlet openings (40 or 41) which arise from the use of a casting core and which connect the cooling passage (39) to an outside space and are sealed off by a sealing element (32 or 36), the sealing element (32, 36) is formed and inserted into the core outlet openings (40, 41) so that they align with a wall surface of the cooling passage (39) in a flush manner, wherein the blade (20) is arranged in a turbine (15, 18) of the gas turbine (10).

10. The gas turbine as claimed in claim 9, wherein the gas turbine (10) is a gas turbine with sequential combustion, having a first combustion chamber (14) with a downstream high-pressure turbine (15), and a second combustion chamber (17) with a downstream low-pressure turbine (18), and the blade (20) is a stator blade which is arranged in the low-pressure turbine (18).

11. The gas turbine as claimed in claim 10, wherein the low-pressure turbine has a plurality of rows of stator blades arranged in series in the flow direction, and the stator blade (20) is arranged in a middle stator blade row.

12. The gas turbine as claimed in claim 9, wherein the blade (20) is a rotor blade.