EP2476863A1 - Turbine blade for a gas turbine - Google Patents

Abstract

The blade (12) has platforms (27, 30) for radially limiting an annular channel (22), and a shovel blade (10) attached to the platforms. A cooling channel system (21) comprising channel sections (24-26) is arranged in the shovel blade, where the sections are connected with respective sets of outlet openings (36, 38). A deflection area (43) and a flow throttle (44) i.e. cover plate, are arranged between the sections. The flow throttle is arranged in the channel system such that the throttle lies opposite to a passage and connected with a pressure-sided blade wall and a suction-sided blade wall.

Description

The invention relates to a turbine blade, in particular a guide vane for a gas turbine, which is designed according to the preamble of claim 1.

Guide vanes of gas turbines are known from the extensive state of the art in many ways. For example, the WO 01/36790 A1 a generic turbine blade with an airfoil, which has a suction-side and a pressure-side blade wall, which extend from a common leading edge to a trailing edge. In the blade of the known turbine blade, a cooling channel system with meandering successive straight channel sections is provided which extend from the so-called blade root to the so-called blade head. Between two adjacent channel sections a deflection region is located, in which the coolant flow is deflected in its direction: with respect to the installation position of the guide vane in the gas turbine, the flow is deflected either from radially outward flowing radially inwardly flowing, or vice versa. According to the WO 01/36790 A1 For example, the vane includes an adjustable throttle during manufacture that is set upon completion of manufacture. This comprises a centrally arranged in the deflection screwed plug, with the aid of the flow-through channel cross-section can be changed.

The disadvantage, however, is that the known method for producing the turbine blade is comparatively complicated, since each turbine blade to be produced has to be checked in a measuring stand. In this case, the coolant flow emerging at the rear edge is detected and the penetration depth of the plug is changed so that the desired amount of coolant is consumed. Accordingly, the plug is firmly connected to the turbine blade.

Furthermore, it is known that in the transitional region of the airfoil to the flow channel, radially outwardly and / or internally delimiting platform, hollow-groove-like fillets are provided which represent mass accumulations. However, these mass accumulations are comparatively difficult to cool, so that at this point more wear, such as cracks, can occur.

The object of the invention is therefore to provide a turbine blade in which the disadvantages described above are eliminated.

The object underlying the invention is achieved with a turbine blade according to the features of claim 1. Further advantageous features are listed in the subclaims and can be combined with each other in an advantageous manner.

According to the invention, the turbine blade comprises an airfoil with a cooling channel system with meandering successive rectilinear duct sections, of which one of these duct sections are directly connected to first outlets opening in the front edge and another of the duct sections are connected to second outlets opening into or at the rear edge and between the two concerned Channel portions a deflection region and a flow restrictor are provided, wherein the flow restrictor is located in the cooling channel system that it faces the outer transition of the blade to the platform and is connected to the pressure-side blade wall and the suction-side blade wall.

The application of the invention is provided in turbine blades with a meander channel whose front edge side cooling passage portion directly connected to the first outlet openings is located in the leading edge of the airfoil and at the same time has a trailing edge side cooling channel portion which is also directly connected to at or in the trailing edge second outlet opening. In this case, the first outlet openings serve as a so-called "shower head", with which the leading edge can be protected from the inflowing hot gas with the discharge of coolant to form a film cooling. At the same time, the comparatively thin trailing edge in front of the hot gas is protected by convective cooling taking place in the interior. As known, the pressure in the hot gas front edge side is greater than the trailing edge side, offers a corresponding gradation of pressure in the coolant, which is required on the one hand to feed the first outlet openings and the other to feed the second outlet openings. By this intended gradation of the coolant pressure can be saved particularly efficient coolant, which can then be used to increase efficiency in the gas turbine combustion efficiency.

In addition, the invention is based on the finding that the flow restrictor does not necessarily have to be located centrally in the deflection region. Rather, it is possible to move them along the cooling channel system, without thereby an increased or decreased cooling air consumption occurs. At the same time, it was recognized that the desired accuracy can be achieved with the help of currently available casting methods, including flow throttles implemented in the casting. In addition, it was recognized that a mitgegossene flow restrictor, which connects the pressure-side blade wall with the suction-side blade wall, can be considered simultaneously as a local cooling element, which can eliminate the hitherto insufficient cooling in connection with the mass accumulations in the transition between the blade and the platform. The arrangement of the flow restrictor in the transition from the deflection region and channel section - ie where the outer contour of the turbine blade from the airfoil into the platform transitions like a gutter - leads to a lower thermal Burden of mass accumulation, which reduces or even prevents wear and tear at this point during operation.

Overall, the invention leads to a comparatively easy to manufacture turbine blade, which has no undesirably high consumption of coolant and at the same time has a prolonged life compared to not inventively designed turbine blades.

Preferably, the invention is used on guide vanes of gas turbines. Nevertheless, the invention is also applicable to blades.

According to a first advantageous embodiment of the invention, the flow restrictor is configured in the form of a diaphragm extending transversely to the coolant flow direction. This can on the one hand a good throttle effect can be achieved, which avoids excessive coolant consumption. On the other hand, during operation, the throttle element is thus flowed transversely from the coolant, whereby upstream of the throttle element to set stagnation points to form an impingement cooling. As is known, the impingement cooling is particularly efficient, so that the heat energy occurring in the blade walls can be transferred to the coolant in a particularly efficient manner via the throttle element connected to this monolithically connected throttle element, and can thus be dissipated.

An alternative embodiment provides that the flow restrictor comprises a plurality of cylindrically shaped pedestals arranged along a row, wherein the row extends transversely to the flow direction. As a result, on the one hand, only a moderate throttling of the coolant can take place, if desired, and nevertheless the additional cooling effect can be maintained.

Particularly preferred is the embodiment in which a discharge opening for coolant is provided in the deflection region. If necessary, this removal opening makes it possible to gain access to the flow restrictor in order to be able to effect a post-processing of the flow cross section of the flow restrictor, if required in a few individual cases. The removal opening may also be provided for the removal of coolant. This can be used, on the one hand, to block gaps and / or on the other hand to cool a so-called U-ring, on which all the vanes of a ring are coupled radially inward in the turbine.

Figur 1

einen Querschnitt durch ein Schaufelblatt einer Turbinenschaufel,

Figur 2

den Längsschnitt durch eine Turbinenschaufel gemäß der Schnittlinie II aus Figur 1 und

Figur 3, 4

jeweils einen Ausschnitt aus dem Längsschnitt einer alternativen Turbinenschaufel.

The further explanation of the invention will be made with reference to the embodiments illustrated in the drawings. In detail show:

FIG. 1

a cross section through an airfoil of a turbine blade,

FIG. 2

the longitudinal section through a turbine blade according to the section line II FIG. 1 and

FIG. 3, 4

each a section of the longitudinal section of an alternative turbine blade.

FIG. 1 shows the cross section through an airfoil 10 of a turbine blade 12. The turbine blade 12 is formed as a guide vane for an axially flow-through gas turbine. The airfoil 10 is aerodynamically curved in a known manner and therefore has one of a working fluid - for example, hot gas - anströmbare leading edge 14. The working medium flowing around the blade 10 leaves the turbine blade 12 at the trailing edge 16. Between the leading edge 14 and the trailing edge 16 there extend a suction-side blade wall 18 and a pressure-side blade wall 20. The blade 10 is hollow in the interior and has a cooling channel system 21. In the illustrated embodiment, the cooling channel system 21 comprises three successive rectilinear channel sections 24, 25, 26. The channel sections 24, 25, 26 extend from one in the installed state ( FIG. 2 ) radially outwardly arranged platform 30 to a radially installed in the installed state platform 27. The two platforms 27, 30 are part of the turbine blade 12 produced in a casting process and belong to the boundary of the annular channel 22 of a gas turbine, not shown, in which the working fluid flows during operation. Between two - viewed in the flow direction of the coolant - consecutive channel sections 24, 25 and 25, 26 each have a deflection region 43 is provided, in which the coolant is deflected. This results in the meandering shape of the cooling channel system 21.

The front edge side cooling channel section 24 is divided by a rib 28 along its longitudinal extent. The rib 28 has a plurality of uniformly distributed passage openings 31 - known as cross-over holes - on.

The longitudinal section II-II through the turbine blade 12 shows FIG. 2 ,

In the front edge 14, a plurality of first outlet openings 36 and in the trailing edge 16 a plurality of second outlet openings 38 are located. The first outlet openings 36 are directly fluidically connected to the channel portion 24, the second outlet openings 38 are connected to the flow channel 22 via trailing edge channels 40 directly fluidically. Immediately means that no further straight channel sections are interposed, which are flowed through by coolant.

The turbine blade 12 has a supply opening 32 for a coolant 34, preferably cooling air, provided on the outer platform 30. The channel sections 24, 25, 26 are in sequential flow communication with the supply port 32, so that the coolant 34 flowing into the turbine blade 12 through the supply port 32 first enters the cooling passage section 24. Portions of the coolant entering the cooling channel section 24 can exit through the first outlet opening 36 arranged in the front edge 14 and mix with the working fluid around the airfoil 10. The first outlet openings 36 are arranged in a grid shape. The remaining part of the coolant 34 flows through the channel section 24 as far as the first deflection region 43, where the coolant 34 is for the most part deflected and then enters the cooling channel section 25. From there, the coolant 34 flows to the radially outer end of the cooling passage section 25 and is reversed again in its flow direction by a further deflection region 43. After the coolant 34 has flowed into the cooling channel section 26, it can branch into the trailing edge channels 40 and exit at the trailing edge side at the second outlet openings 38. In the deflection regions 43, the flow direction of the coolant 34 is usually changed by approximately 180 °.

Although the turbine blade 12 after FIGS. 1 and 2 With only one supply port 32 for coolant 34, it is desirable to provide coolant at different pressures at the first exit ports 36 and at the second exit ports 38. For this reason, a flow restrictor 44 is located in the transition between the deflection region 43 and a rectilinear cooling passage section 26. The flow restrictor 44 according to FIG. 2 comprises two rib-like elements 46 which each extend from the pressure-side blade wall 20 to the suction-side blade wall 18 and are simultaneously produced therewith in the casting process. As a result, the flow restrictor 44 is monolithically connected to the blade walls 18, 20 in the form of an aperture extending transversely to the flow direction of the coolant 34.

The flow restrictor 44 causes a lower coolant supply pressure prevails in the cooling passage section 26 than in the cooling passage section 24, so that the second outlet openings 38 are acted upon by a lower coolant supply pressure than the first outlet openings 36. The latter are with the still unthrottled coolant 34 supplied via the feeder 32. As a result of this configuration, a virtually identical pressure ratio prevails at the respective outlet openings 36, 38 in relation to the respective local pressure in the working medium. The rib-shaped elements 46 of the flow restrictor 44 are - with respect to the outer contour of the turbine blade 12 - located so that their imaginary extension extends through a cusp-like transition 48, which connects the airfoil 10 with the platform 30. Thus, the rib-shaped elements 46 serve on the one hand as a throttle element and on the other as cooling elements, which now better cool the previously insufficiently cooled transitions 48.

As a result, thermal stresses in the region of the transition 48, between the blade 10 and platform 30, which have hitherto occurred due to the mass accumulations, can be reduced.

FIG. 3 shows an alternative turbine blade 12 with only two rectilinear cooling duct sections 24, 25, between which a deflection region 43 is arranged. To FIG. 1 and FIG. 2 identical features are in FIG. 3 and also in FIG. 4 provided with identical reference numerals.

In the transition from the deflection region 43 to the cooling channel portion 25 - viewed in the flow direction of the coolant 34 - a flow restrictor 44 is provided. This was made simultaneously with the blade walls 18, 20 during casting of the turbine blade 12 so that it merges into the blade walls 18, 20. In the FIG. 3 illustrated embodiment of the flow restrictor 44 is modeled on a diaphragm with a central and two channel wall-side passages for the coolant 34. The longitudinal extent of the flow restrictor 44 is thus aligned transversely to the coolant flow.

An alternative form of the flow restrictor 44 shows FIG. 4 im too FIG. 3 identical longitudinal section through a turbine blade 12. Instead of in FIG. 3 from two rib-shaped Elements 46 formed flow restrictor 44 includes the flow restrictor 44 according to the in FIG. 4 shown solution, arranged along a row cylindrical base 50 at the level of the fillet-like rounding between the blade 10 and platform.

The embodiments illustrated in the figures are merely illustrative of the invention and are not limiting for this. For example, it is also possible, the flow restrictor - as viewed in flow direction of the coolant - settle in the transition from rectilinear cooling passage section to deflection.

Overall, the invention specifies a turbine blade 12, which has a suction-side blade wall 18 and a pressure-side blade wall 20, which extend from a common leading edge 14 to a trailing edge 16, wherein in the blade leaf 10 a cooling channel system 21 with meandering successive channel sections 24, 25, 26 is provided, of which one of the channel sections 24 are connected to opening into the front edge 14 first outlet openings 36 and another of the channel sections 26 with opening into or at the trailing edge 16 second outlet openings 38. On the one hand to avoid an undesirably high consumption of coolant and on the other hand to increase the life of the turbine blade, it is proposed that a deflection region and a flow restrictor 44 are provided between the two respective channel sections, the flow restrictor 44 being arranged in the transition from the deflection region 43 to channel section 26 and connected to the pressure side vane wall 20 and the suction side vane wall 18.

Claims (4)

Turbine blade (12),
with at least one platform (27, 30) provided for the radial delimitation of an annular channel (22) and an airfoil (10) adjoining it via a hollow throat-like transition (48), which has a suction-side blade wall (18, 20) and a pressure-side blade wall (18 , 20) extending from a common leading edge (14) to a trailing edge (16),
wherein in the airfoil (10) a cooling channel system (21) with meandering successive straight channel sections (24, 25, 26) is provided, of which one of the channel sections (24, 25, 26) with in the front edge (14) opening first outlet openings (36 , 38) and another of the channel sections (24, 25, 26) with directly in or at the trailing edge (16) opening second outlet openings (36, 38) are directly connected and
wherein a deflection region (43) and a flow throttle (44) are provided between the two respective channel sections (24, 25, 26),
characterized,
in that the flow throttle (44) is located in the cooling channel system (21) so as to face the transition (48) and is connected to the pressure-side blade wall (18, 20) and the suction-side blade wall (18, 20). Turbine blade (12) according to claim 1,
in which the flow throttle (44) is designed in the form of a diaphragm extending transversely to the flow direction. Turbine blade (12) according to claim 1,
wherein the flow restrictor (44) comprises a plurality of pedestals (50) arranged along a row, the row extending transversely to the flow direction. Turbine blade (12) according to claim 1, 2 or 3, in which a removal opening (45) is provided in the deflection region (43).

Images