Gas turbine rotor with an internally cooled gas turbine blade and connecting configuration including an insert strip bridging adjacent blade platforms

• US 6,561,764 B1
• Filed: 09/19/2001
• Issued: 05/13/2003
• Est. Priority Date: 03/19/1999
• Status: Expired due to Fees

• Current Assignee: Siemens AG
• Original Assignee: Siemens AG
• Inventors: Tiemann, Peter
• Primary Examiner(s): VERDIER, CHRISTOPHER M
• Application Number:
  US09/936,990
• Time in Patent Office: 601 Days
• Field of Search: 416/193.A, 416/190, 416/191, 416/220.R, 416/221, 416/248, 416/96.R, 416/97.R, 415/135, 415/138, 415/139, 277/630, 277/634, 277/643, 277/650
• US Class Current: 416/96.00R
• CPC Class Codes: F01D 11/006 Sealing the gap between rotor blades or blades and rotor
• CPC Class Codes: F01D 5/3015 with side plates

Abstract

A gas turbine rotor includes an internally cooled gas turbine blade which has a blade root and a blade platform with recesses into which insert strips are inserted. The gas turbine rotor is arranged in such a way that both sealing against the ingress of hot gas and the emergence of cooling air and the securing of the gas turbine blade can be effected at a low outlay and at the same time are highly reliable. The recess reaches as far as the disk-side base of the blade platform and the insert strip to have a form fit to the disk, which protects against axial displacement in a direction of insertion of the gas turbine blade.

Claims

1. A gas turbine rotor, comprising:

an internally cooled gas turbine blade, the blade including a blade root and a blade platform; and

a disk including transverse disk grooves, into one of which the blade root is inserted, the blade platform being arranged outside an outer circumference of the disk and including, on an end wall portion of one longitudinal side, a recess into which is inserted an insert strip which engages into a corresponding second recess of a second blade platform of a second gas turbine blade and bridges and seals off an interspace between the two blade platforms, wherein the recess reaches as far as a disk-side base of the blade platform, and wherein the insert strip includes a form fit to the disk which protects against axial displacement in a direction of insertion of the gas turbine blade, and wherein the recess runs obliquely in the direction of a longitudinal rotor axis of the gas turbine rotor and at its disk-side end, issues into a chamfer of the disk head, said chamfer being oblique corresponding to the recess.

2. The gas turbine rotor as claimed in claim 1, wherein a recess is included on each of the end wall portions of one longitudinal side of the blade platform, and in which its insert strips bear on two mutually opposite disk end faces and protect against axial displacements of the gas turbine blade relative to the disk.

3. The gas turbine rotor as claimed in claim 1, wherein, with the gas turbine blade inserted into the transverse disk groove, the insert strip can be pushed into the recess on the disk side and is secured counter to centrifugal forces by form fit on at least one of the disk and the blade platform.

4. The gas turbine rotor as claimed in claim 1, wherein between the disk-side end of the recess and the outer circumference of the disk, there is a gap allowing relative movements between the disk and the blade platform.

5. The gas turbine rotor as claimed in claim 1, wherein the recess is designed as a groove.

6. The gas turbine rotor as claimed in claim 1, wherein the recess is wider at its disk-side end than in its remaining part.

7. The gas turbine rotor as claimed in claim 1, wherein the insert strip is fixed in its push-in position on the disk.

8. The gas turbine rotor as claimed in claim 1, wherein a screw part serves for fixing the insert strip, the screw part engaging into a recess of the insert strip being supported on the insert strip under the action of centrifugal force.

9. The gas turbine rotor as claimed in claim 1, wherein the insert strip engages at its disk-side end into a securing recess of the disk end face.

10. The gas turbine rotor as claimed in claim 1, wherein a securing strip leads through two insert strips below the blade platform and is bent round at its ends.

11. The gas turbine rotor as claimed in claim 1, wherein the insert strip is designed in the form of at least one of a wire and of a metal sheet.

12. The gas turbine rotor as claimed in claim 1, wherein the insert strip engages with a predetermined engagement depth into the recess of the blade platform, and wherein the engagement depth of the insert strip into the recess is greater than the interspace between two adjacent blade platforms.

13. The gas turbine rotor as claimed in claim 1, wherein, with the gas turbine blade inserted into the transverse disk groove, the insert strip can be pushed into the recess on the disk side and is secured counter to centrifugal forces by form fit on at least one of the disk and the blade platform.

14. A gas turbine rotor, comprising:

an internally cooled gas turbine blade, the blade including a blade root and a blade platform; and

a disk including transverse disk grooves, into one of which the blade root is inserted, the blade platform being arranged outside an outer circumference of the disk and including, on an end wall portion of one longitudinal side, a recess into which is inserted an insert strip which engages into a corresponding second recess of a second blade platform of a second gas turbine blade and bridges and seals off an interspace between the two blade platforms, wherein the recess reaches as far as a disk-side base of the blade platform, and wherein the insert strip includes a form fit to the disk which protects against axial displacement in a direction of insertion of the gas turbine blade, wherein a recess is included on each of the end wall portions of one longitudinal side of the blade platform, and in which its insert strips bear on two mutually opposite disk end faces and protect against axial displacements of the gas turbine blade relative to the disk, and wherein the recess runs obliquely in the direction of a longitudinal rotor axis of the gas turbine rotor and at its disk-side end, issues into a chamfer of the disk head, said chamfer being oblique corresponding to the recess.

15. The gas turbine rotor as claimed in claim 14, wherein, with the gas turbine blade inserted into the transverse disk groove, the insert strip can be pushed into the recess on the disk side and is secured counter to centrifugal forces by form fit on at least one of the disk and the blade platform.

16. The gas turbine rotor as claimed in claim 14, wherein between the disk-side end of the recess and the outer circumference of the disk, there is a gap allowing relative movements between the disk and the blade platform.

17. The gas turbine rotor as claimed in claim 14, wherein the recess is designed as a groove.

18. The gas turbine rotor as claimed in claim 14, wherein the recess is wider at its disk-side end than in its remaining part.