TURBINE BLADE OR A TURBINE VANE FOR A GAS TURBINE

Abstract

A turbine blade or vane for a gas turbine has successively along a radial direction of the gas turbine, a root for attaching the turbine blade or vane to a carrier, a platform, an aerodynamically shaped hollow airfoil with a suction side wall and a pressure side wall extending with respect to the direction of a hot gas flow from a common leading edge to common a trailing edge and extending transversely thereof from the platform to an airfoil tip. The airfoil has at least one cooling cavity extending in a cooling fluid flow direction from a platform level to the airfoil tip, the cooling cavity in fluid connection with a number of cooling fluid outlets distributed along the trailing edge through an array of impingement cooling features located therebetween. The array extends into a region which is located radially outside the airfoil within the platform having impingement cooling features.

Claims

1. A turbine blade or turbine vane for a gas turbine, comprising successively along a radial direction of said gas turbine, a root for attaching the turbine blade or turbine vane to a carrier, a platform, an aerodynamically shaped hollow airfoil comprising a suction side wall and a pressure side wall extending with respect to the direction of a hot gas flow from a common leading edge to common a trailing edge and extending transversely thereof from said platform to an airfoil tip, wherein the airfoil comprises at least one cooling cavity extending in accordance to a cooling fluid flow direction from a platform level to said airfoil tip, said at least one cooling cavity being in fluid connection with a number of cooling fluid outlets distributed along the trailing edge through an array of impingement cooling features located there between, wherein said array extends into a region which is located radially outside the airfoil within the platform comprising also impingement cooling features.

2. A turbine blade or turbine vane according to claim 1, wherein the impingement cooling features are formed as cross-over-hole, wherein said array comprises at least one row of cross-over-holes, at least one of said rows comprises at least one cross-over-hole completely located within the platform.

3. A turbine blade or turbine vane according to claim 1, wherein the impingement cooling features are formed as pin fins, wherein said array comprises at least one row of pin fins, the pin fins have, as seen in longitudinal section of the turbine blade or turbine vane, a rectangular shape.

4. A turbine blade or turbine vane according to claim 1, wherein said cooling cavity is also bordered from an airfoil stiffening rib ending radially inwardly at a rib end at a turnaround section for said cooling fluid, said rib end located radially inward of said platform level.

5. A turbine blade or turbine vane according to claim 4, wherein the rib and the array end underneath a platform hot gas surface on the same level.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0019] Embodiments of the invention are now described, by way of example only, with reference to the accompanying drawings of which:

[0020] FIG. 1 shows a longitudinal cross through a turbine blade and

[0021] FIG. 2 shows a longitudinal cross section through a turbine vane.

DETAILED DESCRIPTION OF INVENTION

[0022] The illustration in the drawings is in schematic form. It is noted that in different figures, similar or identical elements may be provided with the same reference signs.

[0023] FIG. 1 shows a longitudinal cross section through a turbine blade 10 according to the invention and FIG. 2 shows also a longitudinal section through a turbine vane 20 according to the invention.

[0024] The turbine blade 10 and turbine vane 20 each comprise a root 12 for attaching the respective part to a carrier. With respect to the turbine blade 10 the carrier could be designed as a rotor disk while with respect to the turbine vane 20 the carrier could be designed as a turbine vane carrier. Rotor disks and turbine vane carriers are well known in the prior art. Turbine vanes 20 can also be fixed at their inner diameter via u-rings.

[0025] Both the turbine blade 10 and turbine vane 20 comprises further successively along a radial direction of said gas turbine a platform 14 and an aerodynamically shaped hollow airfoil 15 comprising a suction side wall and a pressure side wall extending with respect to the direction of a hot gas flow 16 from a common leading edge 18 to a common trailing edge 22 and extending transversely thereof from said platform 14 to an airfoil tip 24. For turbine vanes 20 said airfoil tip is also known as vane head. Further each the turbine blade 10 and the turbine vane 20 comprises cooling fluid entries 26 through which during operation of the gas turbine cooling fluid 28 could be fed into the interior. Each entry 26 is in fluid connection with a cooling cavity 30 through one or more cooling passages 32. Each of said cooling passages a cooling cavity 30 extends substantially between the platform 14 and the airfoil tip 24. In view of the cooling fluid direction an array 34 of impingement cooling features 29 follows the cooling cavity 30. Further downstream of the array 34 of impingement cooling features 29 a number of cooling fluid outlets 38 are arranged in the trailing edge 22 of the airfoil 15.

[0026] As displayed in FIG. 1 the array of impingement cooling feature 29 could comprise three rows of cross-over-holes 31 followed by the cooling fluid outlets 38 while the array 34 of impingement cooling features 29 of the turbine vane 20 comprises only two rows pin fins 36. Each pin fin 36 connects the suction side wall with the pressure side wall for enabling heat transfer from said wall into the cooling fluid stream surrounding the pin fins 36. Within each row of pin fins 36 subchannels 35 are provided for passing the cooling fluid towards the cooling fluid outlets 38.

[0027] The individual cooling passages 32 and cooling cavity 30 are separated by a set of airfoil stiffening rips 40. As displayed in the drawings the individual cooling passages and cooling cavities mergers into each other in turnaround sections 42.

[0028] Each platform 14 has a first surface 33 facing the hot gas path 13. As shown by the dashed line said first surface 33 determines radially a platform level 17.

[0029] Said platform level 17 defines the separating plane between the airfoil 15 and the platform 14. According to the invention the array 34 of cross-over-holes 31 or pin fins appears on both sides of said platform level 17 hence extending radially significantly into a platform region 37 that is located radially outside the airfoil 15 within the platform 14.

[0030] In operation cooling fluid 28 is fed through the entries 26 to the turbine blade 10 or turbine vane 20 and flows through their cooling passages 32 into the cooling cavity 30 from which it distributes into the individual subchannels located between the pin fins of the first row of pin fins 36. Downstream thereof the cooling fluid impinges onto the pin fins of the subsequent rows located of respective subchannels cascadely.

[0031] Hence also in the platform region 37 said cooling occurs. This reduces the temperature of the airfoil walls and especially the fillet between airfoil 15 and platform 14, also upstream with regard to the hot gas flow direction of the trailing edge 22 without technical disadvantages that film cooling holes would generate if applied there. Finally the heated cooling fluid leaves the airfoil 15 at the trailing edge through the outlets 38.

[0032] Of course the idea of the array extending into the platform is also applicable for turbine vanes 20 at their inner diameter platform. Even pin fins were explained on the basis of the turbine vane 20 and cross-over-holes 31 were explained on the basis of the turbine blade 10, it is understood that pins fins could be applied in turbine blades and cross-over-holes 31 could be applied in turbine vanes, both alone or in combination the corresponding impingement cooling feature 29.

[0033] As displayed in FIGS. 1 and 2 the airfoil stiffening rip 40 which separate the cooling passage 32 from the cooling cavity 30 ends with its rip end 46 on the same radial level as the array 34 ends. This provides a reliable cooling fluid supply for this section of the array 34, which is outside of the airfoil 15.