US20090252615A1

US20090252615A1 - Cooled Turbine Rotor Blade

Info

Publication number: US20090252615A1
Application number: US12/310,690
Authority: US
Inventors: Heinz-Jürgen Gross
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2006-09-04
Filing date: 2007-06-27
Publication date: 2009-10-08
Also published as: RU2009112405A; EP2059655B1; CN101512106A; WO2008028702A1; ES2340338T3; ATE458126T1; EP2059655A1; DE502007002880D1; EP1895096A1; JP2010502872A; RU2410546C2; PL2059655T3

Abstract

A cooled turbine rotor blade for a gas turbine which is traversed axially by flow and is equipped with an attachment area and an airfoil profile is provided. Meandering cooling channels with interposed deflecting regions are provided in the interior of the airfoil profile. In the deflecting regions, it is possible to prevent dead water regions, which are generated in the prior art, by virtue of at least one of the ribs running so as to curve towards the leading edge or towards the trailing edge in the region of the airfoil tip. At the same time, an opening is provided in the curvature of the rib. Through this opening a part of the coolant flows in the deflecting region and can pass over into the adjacent cooling duct.

Description

This application is the US National Stage of International Application No. PCT/EP2007/056425, filed Jun. 27, 2007 and claims the benefit thereof. The International Application claims the benefits of European application No. 06018490.0 EP filed Sep. 4, 2006, both of the applications are incorporated by reference herein in their entirety.

FIELD OF INVENTION

The invention relates to a cooled turbine rotor blade.

BACKGROUND OF INVENTION

By way of example, a turbine rotor blade of this generic type and having an airfoil profile is known from EP 0 735 240 A1. A plurality of mutually adjacent cooling channels are provided in order to cool the airfoil profile, are arranged in a meandering shape, and a coolant can flow through them sequentially. The cooling channels in this case each run parallel to the leading edge. Respectively adjacent cooling channels are separated from one another by ribs, with the ribs ending in a direction-reversal area in which the adjacent cooling channels merge into one another. In order to avoid regions with lower flow rates and in consequence inadequate cooling in these direction-changing areas, in which the cooling air changes its direction, for example, from a flow directed outward to a flow directed inward, direction-changing blades (FIG. 12) are provided at these points. Despite the direction-changing blades, it is, however, still possible for local overheating to occur in the direction-changing area, and this then reduces the life of the turbine blade.
Furthermore, a turbine blade is known from U.S. Pat. No. 5,246,340, which has a plurality of mutually parallel cooling channels in the interior. In this case, the cooling channels are in each case separated by a rib. An opening which connects two adjacent cooling channels is provided in one of the ribs in the area of the blade tip, through which opening a lateral flow can pass for impingement cooling of the blade airfoil tip.
Furthermore, GB 2 106 996 discloses a turbine blade having an impingement cooling insert in the form of a laminate.

SUMMARY OF INVENTION

The object of the present invention is to provide a turbine rotor blade whose life is further improved.
The object relating to the provision of a turbine rotor blade of this generic type is achieved by designing this turbine rotor blade according to the characterizing part of claim 1. It is proposed that at least one of the ribs—seen from the attachment area to the tip area—has an essentially constant rib thickness and is curved toward the leading edge or trailing edge forming a cooling-channel corner area, which has an acute angle in longitudinal section, in the area of the airfoil tip, and that at least one opening is provided, which is arranged in the curvature, connects two adjacent cooling channels and through which a part of the coolant flow of the cooling channel which is adjacent to the corner area can flow into the acute-angled corner area of the cooling channel.
The curved rib results in the direction of the cooling air flowing through the cooling channels being changed in a considerably more aerodynamic manner. The direction change is an integral component of the rib, as a result of which the regions with a relatively low flow rate or no flow rate (dead-water regions) in the direction-changing area can be avoided. The flow rate is in consequence kept approximately constant in that cooling channel toward which the rib is curved. However, the curvature of the rib results in an acute-angled corner area in the adjacent cooling channel, in which dead-water regions could now once again occur. In order now to avoid the dead-water regions in the corner area in the adjacent cooling channel, at least one opening is also provided, which is arranged in the curvature and connects the two adjacent cooling channels, and through which a part of the cooling flow can pass over or flow over from one cooling channel into the other cooling channel at an early stage.
Furthermore, the opening which is arranged in the curved rib can be provided in a particularly simple form. The casting apparatus which is used for casting the turbine rotor blades comprises, in order to produce the cavities through which a coolant can flow, a casting core which has core elements arranged in a meandering shape. In order to support these adjacent core elements, which are arranged in a meandering shape, with respect to one another, a core support can be provided between two adjacent core elements and, after removal of the casting core from the cast, integral turbine blade, leaves behind it the opening within the curved rib. This results in a stabilized casting core which improves the accuracy of the production method.
Further advantageous refinements of the invention are specified in the dependent claims.
In one particularly advantageous refinement, the terminating wall, which is likewise frequently subject to local overheating and is also referred to as a crown base, can also be impingement-cooled on the basis of the coolant jet passing through the opening, in such a way that this likewise makes it possible to cool the terminating wall particularly efficiently. To do this, the opening just has to be inclined such that its longitudinal extent is directed at the terminating wall.
The rib which is adjacent to the trailing edge is preferably curved in the area of the airfoil tip. In this case, the rib—seen from the attachment area to the tip area—is curved toward the leading edge thus making it possible to provide an essentially constant flow cross-sectional area in a part of the direction-changing area between two adjacent coolant channels. This reduces the pressure losses in the coolant. In order to provide a particularly lightweight turbine rotor blade, the rib has an essentially constant rib thickness along its curvature.
In one advantageous development of the invention, the inner face of the terminating wall is equipped with turbulators, thus making it possible to improve the cooling of the terminating wall or of the crown base in a simple manner. Depending on the configuration of the turbine rotor blades, it is possible for a coolant to flow sequentially or else in parallel through the adjacent cooling channels. If the flow passes in parallel through the coolant channels, care must be taken to ensure that there is an adequate pressure gradient between them, in order to obtain a coolant flow which passes through the opening.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be explained with reference to a drawing. The single FIGURE in this case shows a longitudinal section through a turbine rotor blade according to the invention with cooling channels arranged in a meandering shape.

DETAILED DESCRIPTION OF INVENTION

The FIGURE shows a longitudinal section through a turbine rotor blade 10 which is produced by a casting method. The turbine rotor blade 10, which is therefore integral, has an attachment area 12, with a firtree-shaped cross section, with a platform 14 and an airfoil profile 16 arranged thereon. The airfoil profile 16, which has an aerodynamically profiled cross section, is formed by a suction-side blade wall and a pressure-side blade wall, which each extend from a leading edge 18 to a trailing edge 20 and in this case surround a cavity, which is arranged in the interior of the airfoil profile 16 and in which a plurality of cooling channels 22 a, 22 b, 22 c, 22 d are provided. The cooling channels 22 are adjacent to one another and each run approximately parallel to the leading edge 18. The mutually adjacent cooling channels 22 are each separated from one another in places by a rib 24 a, 24 b, 24 c which connects the pressure-side blade wall to the suction-side blade wall. The cooling channels 22 are bounded by a terminating wall 28 in the area of the airfoil tip 27 which is opposite the attachment area 12. The terminating wall 28 is also referred to as a crown base.
The turbine rotor blade 10 which is illustrated in the FIGURE has a cooling channel 22 a on the leading-edge side to which, on the attachment side, a coolant 29, for example cooling air or cooling vapor, can be supplied. The cooling air that is supplied cools the area of the leading edge 18 of the airfoil profile 16 using conventional cooling methods, for example convection cooling, impingement cooling and/or film cooling.
The coolant 29, which can be supplied to the root end of the cooling channel 22 b, flows along the channel 22 b to the airfoil tip 27, and its direction is then changed in a direction-changing area 30 in order to reverse its flow direction, specifically toward the attachment area 12. For this purpose, the rib 24 c which is adjacent to the trailing edge 20 is curved in the area of the airfoil tip 27, with a constant rib thickness D. The curvature 32 is such that the rib 24 c—seen from the attachment area 12 to the tip area 26—is curved toward the leading edge 18. This results in a part of the direction-changing area 30 having a cooling channel width B which is approximately constant in comparison to the cooling channel 22 c. This makes it possible to change the direction, in a particularly aerodynamic manner, of the coolant 29 which flows through the cooling channels 22 b, 22 c sequentially.
An acute-angled corner area 34 is formed by the curvature 32 of the rib 24 c, which is adjacent to the trailing edge 20, in the cooling channel 22 d in the area of the airfoil tip 27. An opening 40 is provided in the rib 24 c in the area of the curvature 32, through which opening 40 the coolant 29 which is flowing in the direction-changing area 30 can partially flow out therefrom and can flow into the corner area 34 by virtue of the pressure ratio there. If required, a plurality of openings 40 may also be provided in order to influence the flow more specifically in the corner areas 34. The corner area 34 can therefore be adequately cooled. Areas with reduced coolant flow rates and in consequence with inadequate cooling are therefore reliably avoided at this point.
The coolant jet passing through the opening 40 impinges on the inner face 42 of the terminating wall 28 and in this case provides impingement cooling for the airfoil tip 27. In order to further improve the cooling effect of the impingement cooling jet, turbulators 44 can also be provided on the inner face 42 of the terminating wall 28, further enlarging the surface area to be cooled. In addition, the coolant 29 which flows along the inner face 42 of the terminating wall 28 can further increase the heat transfer coefficient on the cooling-air side by virtue of the stimulation of turbulence, thus making it possible to achieve even better cooling of the crown base.
It is also feasible according to the invention for the rib 24 a to merge, curved in the direction of the trailing edge 20, into the terminating wall 28 in the tip area 26 of the turbine rotor blade 10, with one or more openings likewise being provided in the curvature.
Overall, the invention specifies a turbine rotor blade 10 for an axial-flow gas turbine, in particular a stationary gas turbine, which is equipped with an attachment area 12, an airfoil profile 16 and a plurality of cooling channels 22 which are arranged in a meandering shape in the interior of the airfoil profile 16. In order to avoid areas with reduced flow rates of coolant 29 in the direction-changing area 30 or at the channel end, the invention proposes that at least one of the ribs 24 run in a curved form toward the leading edge 18 or toward the trailing edge 20 in the area of the airfoil tip 27, with the rib thickness D remaining constant, and that at least one opening 40 be provided in the curvature 32 of the rib 24, through which opening 40 a portion of the coolant 29 which is flowing in the direction-changing area 30 can pass into the adjacent cooling channel 22 d.

Claims

1.-6. (canceled)

7. A cooled turbine rotor blade for a stationary axial-flow gas turbine, comprising:

an airfoil profile formed by a suction-side blade and a pressure-side blade wall;

an airfoil tip;

an attachment area from which the airfoil profile extends as far as the airfoil tip;

a plurality of cooling channels which lie adjacent to one another in an interior of the airfoil profile;

a plurality of ribs;

a terminating wall bounding the plurality of cooling channels at the airfoil tip end; and

an opening arranged in a curvature of one of the plurality of ribs connecting two adjacent cooling channels;

wherein the airfoil profile has a leading edge and a trailing edge,

wherein at least one of the plurality of ribs connects the pressure-side blade wall to the suction-side blade wall and extends from the attachment area to the airfoil tip,

wherein at least one of the plurality of ribs has an essentially constant rib thickness and is curved toward the leading edge or the trailing edge forming a cooling-channel corner area, which has an acute angle in a longitudinal section, in the area of the airfoil tip,

wherein the plurality of cooling channels are at least partially separated from one another by in each case one of the plurality of ribs, and

wherein a part of a coolant flow of one of the plurality of the cooling channels, the cooling channel that is adjacent to a corner area, can flow through the opening into an acute-angled corner area of the adjacent cooling channel.

8. The cooled turbine rotor blade as claimed in claim 7, wherein the rib adjacent to the trailing edge is curved in the area of the airfoil tip.

9. The cooled turbine rotor blade as claimed in claim 7, wherein the opening is arranged such that the terminating wall is impingement-cooled.

10. The cooled turbine rotor blade as claimed in claim 7, wherein an inner face of the terminating wall is equipped with turbulators.

11. The cooled turbine rotor blade as claimed in claim 7, wherein the coolant flows sequentially through the plurality of cooling channels.

12. The cooled turbine rotor blade as claimed in claim 7, wherein the coolant flows in parallel through the plurality of cooling channels.

13. The cooled turbine rotor blade as claimed in claim 7, wherein the cooled turbine rotor blade is cast.

14. The cooled turbine rotor blade as claimed in claim 7, wherein a second rib which is the second one adjacent to the leading edge, curves in the direction of the trailing edge into the terminating wall in the tip area of the cooled turbine rotor blade with an opening provided in a curvature.