TURBINE BLADE FOR A STATIONARY GAS TURBINE

Abstract

A turbine blade having a blade airfoil. A first cooling path for a first coolant stream and a second cooling path for a second coolant stream are formed within the blade airfoil. The first cooling path includes a first coolant passage, which is designed for cyclone cooling of the leading edge, and a second coolant passage, which adjoins the first coolant passage and extends below the blade tip from the leading edge toward the trailing edge. The second cooling path includes a serpentine coolant passage for cooling a central region of the blade airfoil and a first trailing-edge coolant passage for partially cooling a trailing-edge region.

Claims

1. A turbine blade for a gas turbine which is flowed through in particular axially, in particular for one of the high-pressure turbine stages thereof, comprising: a blade root and a blade airfoil comprising a pressure-side side wall and a suction-side side wall, which side walls extend extends along a spanwise direction from a root-side end to a blade tip and along a chordwise direction, which is oriented transversely to the spanwise direction, from a leading edge to a trailing edge, wherein, in the interior of the blade airfoil, a first cooling path for a first coolant stream and a second cooling path, substantially separated from the first cooling path, for a second coolant stream are formed, wherein the first cooling path comprises a first coolant passage, which is configured for cyclone cooling of the leading edge, and a second coolant passage, which adjoins the first coolant passage and extends below the blade tip from the leading edge in the direction of the trailing edge, wherein the second cooling path comprises a serpentine coolant passage for cooling of a middle region of the blade airfoil, which middle region is arranged behind the leading-edge region in the chordwise direction, and a first trailing-edge coolant passage for at least partial cooling of a trailing-edge region of the blade airfoil, which trailing-edge region is arranged behind the middle region in the chordwise direction and extends as far as the trailing edge, wherein the first trailing-edge coolant passage is connected in terms of flow to a multiplicity of first exit holes arranged in the trailing edge, wherein the first coolant passage and/or the serpentine coolant passage are/is free of exit holes, and wherein the first cooling path comprises a third coolant passage, which adjoins the second coolant passage and extends mainly radially inwardly, and a second trailing-edge coolant passage, which adjoins the third coolant passage and is configured for cooling of a blade-tip-side region of the trailing-edge region and is connected in terms of flow to a multiplicity of second exit holes arranged in the leading edge.

2. The turbine blade as claimed in claim 1, wherein one or more exit holes for coolant are arranged in the blade tip and are connected in terms of flow to the second coolant passage.

3. The turbine blade as claimed in claim 1, wherein the first cooling path comprises a supply passage for the first coolant passage, which, in a manner arranged directly adjacent to the first coolant passage and extending at least over a major part of the span width of the blade airfoil, is connected in terms of flow to the first coolant passage via a multiplicity of passage openings, wherein the passage openings have means for imparting swirl to the coolant flowing in the first coolant passage.

4. The turbine blade as claimed in claim 3, wherein a density, ascertainable in the spanwise direction, of passage openings is greatest at the root-side end, and preferably decreases in a stepped manner or continuously toward the blade tip.

5. The turbine blade as claimed in claim 1, wherein a multiplicity of pedestals arranged in a pattern is provided in each trailing-edge coolant passage.

6. The turbine blade as claimed in claim 1, wherein provision is made of two cooling-channel arms, which widen the second coolant passage and, with increasing extent in the chordwise direction, expand radially inward and open out in the third coolant passage.

7. The turbine blade as claimed in claim 6, wherein a separating wall is arranged between the second coolant passage and the serpentine coolant passage and connects the two side walls to one another and extends in the chordwise direction, wherein, with progressively closer proximity to the trailing edge, the separating wall forms a displacement wedge which narrows preferably to a point and which, in conjunction with the inner surfaces of the two side walls, laterally delimits the two cooling-channel arms.

8. The turbine blade as claimed in claim 1, wherein a rear separating rib is provided between the third coolant passage and the second trailing-edge coolant passage and extends in the spanwise direction.

9. The turbine blade as claimed in claim 1, wherein the trailing edge has a normalized height of 100%, beginning at its root-side end at 0% and ending at the blade tip at 100%, and wherein the two trailing-edge coolant passages are separated from one another by a separating rib which extends mainly in the chordwise direction and which is arranged at a height of between 45% and 75% of the normalized height.

10. The turbine blade as claimed in claim 1, wherein the serpentine coolant passage comprises at least two channel sections, extending in the spanwise direction, and at least two reversal sections, wherein the reversal section situated further downstream in the coolant stream is connected in terms of flow directly to the first trailing-edge coolant passage.

11. The turbine blade as claimed in claim 10, wherein the two channel sections, by a displacement body and by the two side walls, are, in a cross-sectional view of the blade airfoil, each of substantially C-shaped form with a suction-side channel arm, a pressure-side channel arm and a connecting arm connecting the two channel arms and are arranged in relation to one another in such a way that they almost completely surround the displacement body.

12. The turbine blade as claimed in claim 11, wherein the displacement body, in a cross-sectional view, reaches around a cavity and is supported via webs against the two side walls.

13. The turbine blade as claimed in claim 11, wherein the serpentine coolant passage is delimited by at least one, preferably by two, support ribs, which connect the pressure-side side wall to the suction-side side wall and extend from the root-side end toward the blade tip and at which provision is made, preferably on the support rib or on the inner surfaces, delimiting the connecting arms, of the displacement body, of elements, preferably turbulators, which reduce a transverse flow of coolant from the suction-side channel arm into the pressure-side channel arm through the connecting arm.

14. The turbine blade as claimed in claim 12, wherein the cavity cannot be flowed through by coolant and in particular has no exit opening for coolant.

15. The turbine blade as claimed in claim 12, wherein the turbine blade is cast, and wherein an opening which is present in the blade root after the casting of the turbine blade and which is connected directly to the cavity is closed off by a separately produced cover plate.

16. The turbine blade as claimed in claim 1, which wherein the turbine blade is cast.

17. The turbine blade as claimed in claim 15, wherein an opening which is present in the blade root after the casting of the turbine blade and which is connected directly to the first trailing-edge coolant passage is closed off by a separately produced cover plate.

18. The turbine blade as claimed in claim 1, wherein, for each cooling path, provision is made of one or more inlets which are connected in terms of flow directly to the first coolant passage or the supply passage or to the serpentine coolant passage or one of the channel sections thereof.

19. The turbine blade as claimed in claim 1, comprising: a blade-airfoil aspect ratio HSP/SL of a trailing-edge span width to a chord length to be measured at the root-side end that is 3.0 or less.

20. A first or second turbine stage of a stationary gas turbine, comprising: a turbine blade as claimed in claim 1, and a turbine-entry temperature, occurring during nominal operation under ISO conditions, of at least 1300° C. and/or having a compressor pressure ratio, occurring during nominal operation under ISO conditions, of 19:1 or greater.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0034] In the figures:

[0035] FIG. 1 shows a side view of a turbine rotor blade according to a first exemplary embodiment,

[0036] FIG. 2 shows the cooling schemes for the turbine rotor blade as per FIG. 1,

[0037] FIG. 3 shows a longitudinal section through the turbine rotor blade according to the first exemplary embodiment,

[0038] FIG. 4 shows a cross section through the turbine rotor blade as per FIG. 3 along the section line A-A,

[0039] FIGS. 5-7 show longitudinal sections through the turbine rotor blade as per FIG. 3 along the section lines B-B, C-C and D-D, respectively,

[0040] FIG. 8 shows a cross section through the turbine rotor blade as per FIG. 1 along the section line E-E, and

[0041] FIG. 9 shows a stationary gas turbine in a schematic illustration.

DETAILED DESCRIPTION OF INVENTION

[0042] In the figures, all technical features denoted by identical reference signs have the same technical effect.

[0043] The invention will be discussed below on the basis of a turbine blade 10 which is in the form of a turbine rotor blade. The invention may however also involve a turbine guide vane.

[0044] FIG. 1 shows, in a side view, a turbine blade 10 as a first exemplary embodiment of the invention. The turbine blade 10, which is preferably produced in a precision casting process, comprises a blade root 12, which is shown only partially. The blade root 12 may, in a known manner, be of dovetail-shaped or fir-tree-shaped form. Said blade root is adjoined by a platform 13, from which a blade airfoil 18 extends in a spanwise direction R from a root-side end 20 to a blade tip 22. If the turbine rotor blade 10 is installed in a gas turbine which is flowed through axially, the spanwise direction and the radial direction of the gas turbine coincide. In a chordwise direction S, which is oriented transversely to the spanwise direction R, the blade airfoil 18 extends from a leading edge 24 to a trailing edge 26. In the trailing edge 26, exit holes 46, 56 are distributed along the spanwise direction. An aspect ratio HSP/SL of a trailing-edge span width HSP to a chord length SL to be measured at the root-side end is 1.9 according to this exemplary embodiment and preferably lies in the range between 1.5 and 3.

[0045] Exit openings 28 open out at a lateral surface of the platform 13 too. The exit holes 46, 56 and the exit openings 28 are connected in terms of flow to an inner cooling system of the turbine rotor blade 10.

[0046] The cooling system of the turbine rotor blade 10 and in particular of the blade airfoil 18 is represented schematically in FIG. 2 by cooling schemes. A first coolant stream M1 and a second coolant stream M2 can be fed separately to the turbine rotor blade 10. The first coolant stream M1 flows through a first cooling path 30, which is made up of multiple coolant passages 31, 32, 33, 34, 36a, 36b, 38, 40, 44. A supply passage 31 follows downstream of an inlet (not illustrated in FIG. 2) for the coolant stream M1, and is connected in terms of flow to a first coolant passage 32 via a multiplicity of passage openings 33. The first coolant passage 32 serves for cyclone cooling of the leading edge 24 of the blade airfoil 18 and of the directly adjoining leading-edge region 39. In the region of the blade tip 22, the first coolant passage 32 transitions into a second coolant passage 34, which, for the purpose of cooling the blade tip 22, extends from the leading edge 24 in the direction of the trailing edge 26 over a relatively large chord length of the blade tip 22. Third exit holes 67 may be arranged in the blade tip for the purpose of cooling rubbing edges, which are mentioned later. Furthermore, the second coolant passage 34 comprises two cooling-channel arms 36a, 36, which begin only in the second half of the second cooling passage 34 and, just like the downstream end of the second coolant passage 34, are connected to a third coolant passage 38. The latter is connected in terms of flow to a second trailing-edge coolant passage 44 via a reversal section 40. The coolant stream M1 flowing through the first cooling path 30 can then exit the turbine rotor blade 10 at the trailing edge 26 thereof via a multiplicity of second exit holes 46. Arranged parallel to the first cooling path 30 and in a manner separated preferably completely in terms of flow therefrom is a second coolant path 50, which, downstream of an inlet (not illustrated in more detail in FIG. 2), has a serpentine coolant passage 52. According to this exemplary embodiment, the serpentine coolant passage 52 comprises, for cooling of a middle region 48 (FIG. 1), two channel sections 55a, 55b which extend in the spanwise direction and which are connected to one another via a reversal section 57a arranged therebetween. A second reversal section 57b adjoins the downstream end of the second channel section 55b, and connects the second channel section 55b to a first trailing-edge coolant passage 54 in terms of flow. The second coolant stream M2 flowing through the second cooling path 50 can then exit the turbine rotor blade 10 at the trailing edge 26 thereof via a multiplicity of first exit holes 46. Both trailing-edge coolant passages 44, 54 serve for cooling a trailing-edge region 59 (FIG. 1).

[0047] FIG. 3 shows, in the form of a longitudinal section, an inner structure of the turbine rotor blade 10 as per FIG. 1, which is formed in a manner corresponding to the cooling schemes in FIG. 2. To this end, the turbine rotor blade 10 comprises a series of differently arranged walls and ribs that separate the individual cooling paths and coolant passages from one another. In the blade root 12, provision is made of two inlets 80 for the two coolant streams M1 and M2 or for the two cooling paths 30, 50. Arranged between the two inlets 80 is a front support rib 66v, which connects the side walls 14, 16 to one another and, for a first section, separates the first cooling path 30 from the second cooling path 50. Moreover, a front separating rib 49v separates the supply passage 31 from the first coolant passage 32, wherein a multiplicity of passage openings 33 (detail of FIG. 4) are arranged in the front separating rib 49v. In FIG. 3, however, of these, merely the mouths of the passage openings are illustrated. As emerges from FIG. 3, a higher density of passage openings 33 is provided in the region close to the platform than in the region close to the tip. The position and the orientation of the passage openings 33 in the front separating rib 49v is selected in such a way that a relatively intensely swirled coolant flow can be formed in the first coolant passage 32. A swirled coolant flow is to be understood as meaning one which can be formed in a cyclone-like manner, or analogously to a spiral line or a helix, from the root-side end 20 to the blade tip 22. Consequently, said passage openings are arranged in the front separating rib 49v eccentrically and in particular in a manner aligned with the inner walls of the suction-side wall 16 (or pressure-side wall), possible even with an inclination toward the blade tip 22 in order to at least partially compensate for the weakening of the swirl during the flow through the first coolant passage 32.

[0048] The outer end of the first coolant passage 32 is adjoined by the second coolant passage 34 for cooling of a base 37 of the blade tip 22, wherein the second coolant passage 34 is separated from the serpentine coolant passage 52 by a separating wall 60. That end of the second coolant passage 34 which is close to the trailing edge is adjoined by the third coolant passage 38, which extends from the blade tip 22 in the direction of the root-side end 22, although only to approximately half the height of the blade airfoil 18, wherein the height of the blade airfoil 18 is to be measured at the trailing edge 26. Said third coolant passage is adjoined by a further reversal section 40, by means of which the first coolant stream M1 can be fed to the second trailing-edge coolant passage 44. The third coolant passage 38 is mostly separated from the second trailing-edge coolant passage 54 by a correspondingly formed rear separating rib 49h.

[0049] In the second trailing-edge coolant passage 44, pedestals 53 which can flowed around by the coolant M1 are arranged one behind the other in multiple rows. In the exemplary embodiment shown, the pedestals have more of a racetrack-shaped form, and relatively narrow passages, so as to bring about the greatest possible pressure loss. The first cooling path 30 ends in second exit holes 46 which are provided in the trailing edge 26 and through which at least a major part of the coolant stream M1 fed through the associated inlet 80 can be released from the turbine rotor blade 10.

[0050] The second cooling path 50 for guiding the second coolant stream M2 and comprises substantially the serpentine coolant passage 52 and the first trailing-edge coolant passage 44. The former can be subdivided into four sections which follow one after the other, of which the first one is referred to as first channel section 55a. There follow in an adjoining manner in succession a first reversal section 57a, a second channel section 55b and a second reversal section 57b. The latter connects the serpentine coolant passage 52 to the second trailing-edge coolant passage 54, which, analogously to the first trailing-edge coolant passage 44, is formed with racetrack-shaped pedestals 53 arranged in multiple rows.

[0051] The two channel sections 55a, 55b of the serpentine coolant passage 52 extend along the spanwise direction R over a major part of the blade airfoil 18. The first channel section 55a as well as the second channel section 55b are, as additionally illustrated in FIG. 4, substantially U-shaped with in each case one channel arm 55as, 55bs arranged on the suction side, one channel arm 55ad, 55bd arranged on the pressure side, and one connecting arm 55av, 55bv connecting the respective channel arms. Accordingly, the first channel section 55a is surrounded by the pressure-side side wall 14, by the front support rib 66v, by the suction-side side wall 16, and by a displacement body 70 (in cross section in FIG. 4) that is arranged in the interior. The second channel section 55b is surrounded by the pressure-side side wall 14, by a rear support rib 66h, by the suction-side side wall 16, and by the displacement body 70 arranged in the interior. The displacement body 70 itself reaches around a cavity 72 and is supported via webs 71 against the pressure-side side wall 14 and the suction-side side wall 16. The webs 71 extend approximately over the entire height of the blade airfoil 18 and serve for monolithic fastening of the displacement body 70 in the turbine rotor blade 10, on the one hand, and for separating the two channel sections 55, 57, on the other hand. Referring to FIG. 2, it can be seen that the displacement body 72 is supported, at its radially outer end, at the trailing-edge side. This measure improves the mechanical integrity of the turbine rotor blade 10 and in particular its resistance to vibration.

[0052] The two trailing-edge coolant passages 44, 54 are separated from one another at least substantially, if not completely, by a separating rib 64 which extends mainly in the chordwise direction S. According to the exemplary embodiment, the separating rib 64 ends at a height of 55% of a normalized blade-airfoil height of the trailing edge 26. Preferably, the separating rib 64 is arranged at a height of between 45% and 75% of the normalized height.

[0053] FIGS. 5 to 7 shows sections through the tip of the turbine rotor blade 10 according to the three section lines B-B, C-C and D-D from FIG. 3. Rubbing edges 78 are provided on the outer end of the blade tip 22, both on the suction side and on the pressure side. Moreover, it can be seen that the displacement body 70, at its radially outer end, is not closed off but is open toward the first reversal section 57a. Although an inflow of the second coolant stream M2 would thus be possible, since an opening 74a at the blade root 12 required for the creation of the cavity 72 or of the displacement body 70 is closed off by a cover plate 76a (FIG. 1) attached after the casting, the cavity 72 lacks exit openings. Therefore, said cavity cannot be flowed through, but rather is in the form of a dead-water space. Consequently, it is expedient for the inner configuration thereof, possibly still during the design phase, to be varied by means of the provision of further structures, such as ribs, struts or the like, if modal adaptation is required. The particular advantage would be that the natural frequency of the turbine blade alone would be adapted, without other properties, such as aerodynamics or heat exchange, being influenced.

[0054] FIGS. 5 to 7 furthermore show how, with progressively closer proximity to the trailing edge 26, the separating wall 60 forms a displacement wedge 62 which narrows to a point and which, in conjunction with the inner surfaces of the two side walls 14, 16, laterally delimits each of the two cooling-channel arms 36a and 36b. With the aid of the displacement wedge 62 that narrows to a point, the truncation of the displacement body 70 can be compensated such that guidance of the coolant stream M2 close to the side walls in the truncated region, and thus sufficient cooling thereof, is still efficiently possible. If the truncation of the displacement body is not absolutely necessary, the size of the displacement wedge can be reduced. Possibly, it can even be dispensed with completely.

[0055] FIG. 8 shows, in a view directed toward the blade tip 22, that is to say toward the outside, a cross section of the downstream half of the blade tip 22 according to section line E-E from FIG. 3.

[0056] According to an exemplary embodiment that is not shown in any further detail, instead of or in addition to the supply passage 31, provision may be made of a blade-root-side channel section which is able to provide an extension of the first coolant passage 32 as far as the bottom side of the blade root 12. In said blade-root-side channel section, provision may be made of correspondingly suitable swirl generators, for example spiral ribs, which swirl the coolant stream M1 in a cyclone-like manner during the flow through the blade-root-side channel section. In this case, the first coolant passage 32 would be separated by the front support rib 66v from the connecting channel 55av such that passage openings 33 arranged in the front support rib 66v could promote replenishment or boosting of the swirl momentum. In this respect, it may possibly even be expedient for the two coolant streams M1 and M2 to be guided through the turbine blade 10 not entirely separated from one another but so as to permit an exchange to a very small extent, in that, at a very small number of locations, individual holes with preferably small diameters connect to one another the two cooling paths, which are otherwise separated in terms of flow.

[0057] FIG. 9 shows, merely schematically, a gas turbine 100 with a compressor 110, a combustion chamber 120 and a turbine unit 130. According to this embodiment, a generator 150 for generating electricity is coupled to a rotor 140. The compressor 110 is designed in such a way that, during operation under ISO standard conditions, it can produce a pressure ratio of compressed ambient air VL to sucked-in ambient air L of 19:1 or greater. The compressed air VL is then mixed with a fuel F, and combusted to form a hot gas HG, in the combustion chamber 120. Combustion chamber 120 and turbine unit 130 are designed in such a way that the hot gas HG flowing at the exit of the combustion chamber 120 and at the entry of the turbine unit 130 has a temperature of at least 1300° C. under ISO standard conditions, wherein the rotor blades and guide vanes of the first turbine stage or the second turbine stage are designed in the manner described herein. The hot gas HG, which is expanded in the turbine unit 130, exits the latter as flue gas RG.

[0058] Altogether, the invention proposes a turbine blade 10 having a blade root 12 and having a blade airfoil 18 that extends along a spanwise direction R from a root-side end 20 to a blade tip 22 and along a chordwise direction S, which is oriented transversely to the spanwise direction R, from a leading edge 24 to a trailing edge 26, wherein, in the interior of the blade airfoil 18, a first cooling path 30 for a first coolant stream M1 and a second cooling path 50 for a second coolant stream M2 are formed, wherein the first cooling path 30 comprises a first coolant passage 32, which is configured for cyclone cooling of the leading edge 24, and a second coolant passage 34, which adjoins the first coolant passage 32 and extends below the blade tip 22 from the leading edge 24 in the direction of the trailing edge 26, wherein the second cooling path 50 comprises a serpentine coolant passage 52 for cooling of a middle region 48 of the blade airfoil 18, which middle region is arranged behind the leading-edge region 39 in the chordwise direction, and a first trailing-edge coolant passage 54 for at least partial cooling of a trailing-edge region 59 of the blade airfoil 18, which trailing-edge region is arranged behind the middle region 48 in the chordwise direction and extends as far as the trailing edge, wherein the first trailing-edge coolant passage 54 is connected in terms of flow to a multiplicity of first exit holes 56 arranged in the trailing edge 26. In order to provide a turbine blade with a further reduced coolant consumption, it is proposed that the first coolant passage 32 and/or the serpentine coolant passage 52 are/is configured for locally closed cooling and the first cooling path 30 comprises a third coolant passage 38, which adjoins the second coolant passage 34 and extends mainly radially inwardly, and a second trailing-edge coolant passage 44, which adjoins the third coolant passage 38 and is configured for cooling of a blade-tip-side region of the trailing-edge region 59 and is connected in terms of flow to a multiplicity of second exit holes 46 arranged in the trailing edge 26.