Gas turbine blade and method for producing a blade

• US 9,062,555 B2
• Filed: 06/20/2012
• Issued: 06/23/2015
• Est. Priority Date: 06/29/2011
• Status: Active Grant

• Current Assignee: Ansaldo Energia Switzerland AG
• Original Assignee: Alstom Technology Ltd. (Alstom SA)
• Inventors: Schnieder, Martin, Krckels, Jrg
• Primary Examiner(s): EDGAR, RICHARD A
• Application Number:
  US13/528,013
• Publication Number:
  US 20130004332A1
• Time in Patent Office: 1,098 Days
• Field of Search: 416/97.R, 416/95, 416/96.R, 416/232, 416/233
• US Class Current: 1/1
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F05D 2260/22141 using fins or ribs
• CPC Class Codes: Y10T 29/49341 with cooling passage

Abstract

A blade (10) for a gas turbine has a blade airfoil (11), the blade wall (18) of which encloses an interior space (17). For cooling the blade wall (18), the blade wall (18) includes a cooling arrangement (19) which has a radial passage (20) extending in the longitudinal direction of the blade and from which a multiplicity of cooling passages (21, 22), extending in the blade wall (18), branch in the transverse direction, and from which a multiplicity of film-cooling holes (23) are led to the outside in the transverse direction. Particularly efficient cooling is made possible by the distribution of the film-cooling holes (23) along the radial passage (20) being selected independently of the distribution of the cooling passages (21, 22) along the radial passage (20).

Claims

1. A blade for a gas turbine, comprising:

a blade airfoil having a blade wall which encloses an interior space;

wherein said blade wall comprises a cooling arrangement configured and arranged to cool the blade wall, the cooling arrangement including a radial passage extending in a longitudinal direction of the blade, a plurality of cooling passages extending in the blade wall from the radial passage and which branch out in a transverse direction, and a plurality of film-cooling holes extending transversely from the radial passage to outside the blade airfoil;

an opening in the blade wall through which the radial passage is accessible from the outside at one end; and

a sealing element in the opening and sealing off the radial passage.

2. The blade as claimed in claim 1, wherein the radial passage is offset towards the inside of the blade airfoil from the middle of the blade wall.

3. The blade as claimed in claim 2, wherein the plurality of film-cooling holes forms a fan-like arrangement.

4. The blade as claimed in claim 1, comprising:

a platform into which the blade airfoil merges at a lower end; and

wherein the radial passage is accessible from the outside at a transition between the blade airfoil and the platform.

5. The blade as claimed in claim 1, comprising:

a platform into which the blade airfoil merges at a lower end, forming a fillet; and

cooling passages in the region of the fillet configured and arranged to cool the fillet.

6. The blade as claimed in claim 1, comprising:

turbulence elements in the plurality of cooling passages configured and arranged to improve cooling.

7. The blade as claimed in claim 6, wherein the turbulence elements comprise ribs or pins.

8. The blade as claimed in claim 1, comprising:

impingement cooling holes which lead from the interior space to the plurality of cooling passages.

9. The blade as claimed in claim 1, wherein the plurality of cooling passages extend only from the radial passage on one side.

10. The blade as claimed in claim 1, wherein the plurality of cooling passages extend from the radial passage on both sides.

11. A method for producing a blade for a gas turbine, the blade comprising a blade airfoil having a blade wall which encloses an interior space, wherein said blade wall comprises a cooling arrangement configured and arranged to cool the blade wall, the cooling arrangement including a radial passage extending in a longitudinal direction of the blade, a plurality of cooling passages extending in the blade wall from the radial passage and which branch out in a transverse direction, and a plurality of film-cooling holes extending transversely from the radial passage to outside the blade airfoil;

wherein the distribution of the plurality of film-cooling holes along the radial passage is selected independently of the distribution of the plurality of cooling passages along the radial passage, an opening in the blade wall through which the radial passage is accessible from the outside at one end, and a sealing element in the opening and sealing off the radial passage, the method comprising;

providing the blade with the radial passage which is open on one side;

inserting a strip-like insert into the open radial passage;

forming film-cooling holes in the blade from the outside, wherein the wall of the radial passage opposite the film-cooling holes is protected by the insert during said forming; and

removing the insert from the radial passage.

12. The method as claimed in claim 11, comprising:

sealing off the radial passage with the sealing element after removing the insert.

13. The method as claimed in claim 12, further comprising:

hard-soldering the sealing element.

14. The method as claimed in claim 11, wherein:

forming film-cooling holes comprises laser drilling; and

inserting a strip-like insert comprises inserting a PTFE strip.