TURBINE BLADE ASSEMBLY

• US 20200131916A1
• Filed: 10/31/2018
• Published: 04/30/2020
• Est. Priority Date: 10/31/2018
• Status: Abandoned Application

• Current Assignee: Raytheon Technologies Corporation d/b/a RTX (Rtx Corporation)
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Eastwood, Jonathan Jeffery, Bergman, Russell J., Caprario, Joseph T., Virkler, Scott D.
• Application Number:
  US16/177,331
• Publication Number:
  US 20200131916A1
• Time in Patent Office: Days
• CPC Class Codes: F01D 5/3007 of axial insertion type
• CPC Class Codes: F01D 5/323 Locking of axial insertion type blades by means of a key or the like parallel to the axis of the rotor
• CPC Class Codes: F01D 5/326 Locking of axial insertion type blades by other means
• CPC Class Codes: F05D 2260/30 Retaining components in desired mutual position

Abstract

A turbine blade assembly for a gas turbine engine is provided. The turbine blade assembly having: a rotor; a plurality of turbine blades mounted to the rotor, each of the plurality of turbine blades having a forward tab that contacts a front face of the rotor and an aft tab; a plurality of tabs that extend from an aft surface of the rotor; and a retaining ring that engages each aft tab of the plurality of turbine blades and each of the plurality of tabs that extend from the aft surface of the rotor.

Claims

1. A turbine blade assembly for a gas turbine engine, comprising:

a rotor;

a plurality of turbine blades mounted to the rotor, each of the plurality of turbine blades having a forward tab that extends radially downward from the turbine blade and contacts a front face of the rotor when a root portion of the turbine blade is slid into a complementary slot of the rotor, wherein each of the plurality of turbine blades has an aft tab that extends from the root portion to define a gap between an aft surface of the root portion and the aft tab;

a plurality of tabs that extend from an aft surface of the rotor; and

a retaining ring that engages each aft tab of the plurality of turbine blades and each of the plurality of tabs that extend from the aft surface of the rotor, wherein an upper peripheral edge of the retaining ring contacts a surface of the plurality of tabs that extend from the aft surface of the rotor when the retaining ring is inserted between the aft tab of the plurality of turbine blades and the rotor, and wherein the upper peripheral edge of the retaining ring does not contact an inner diameter surface of the plurality of aft tabs that are located radially above the retaining ring when the retaining ring is inserted between the aft tabs of the plurality of turbine blades and the rotor.

2. The turbine blade assembly as in claim 1, wherein the plurality of tabs are integrally formed with the rotor.

3. The turbine blade assembly as in claim 1, wherein the retaining ring further comprises an anti-rotation feature that is located between a pair of aft tabs of adjacent turbine blades secured to the rotor.

4. The turbine blade assembly as in claim 3, wherein the anti-rotation feature protrudes from an aft surface of the retaining ring.

5. The turbine blade assembly as in claim 4, wherein the retaining ring is provided with a split.

6. The turbine blade assembly as in claim 5, wherein the split is located radially below one of plurality of tabs that extends from the aft surface of the rotor.

7. The turbine blade assembly as in claim 1, wherein the retaining ring is provided with a split.

8. The turbine blade assembly as in claim 7, wherein the split is located radially below one of plurality of tabs that extends from the aft surface of the rotor.

9. The turbine blade assembly as in claim 1, wherein the outside diameter of the plurality of turbine blades when secured to the rotor is up to 15.5 inches.

10. The turbine blade assembly as in claim 1, wherein the outside diameter of the rotor is up to 11 inches.

11. A gas turbine engine having at least one blade assembly, the at least one blade assembly comprising:

a rotor;

a plurality of turbine blades mounted to the rotor, each of the plurality of turbine blades having a forward tab that extends radially downward from the turbine blade and contacts a front face of the rotor when a root portion of the turbine blade is slid into a complementary slot of the rotor, wherein each of the plurality of turbine blades has an aft tab that extends from the root portion to define a gap between an aft surface of the root portion and the aft tab;

a plurality of tabs that extend from an aft surface of the rotor; and

a retaining ring that engages each aft tab of the plurality of turbine blades and each of the plurality of tabs that extend from the aft surface of the rotor, wherein an upper peripheral edge of the retaining ring contacts a surface of the plurality of tabs that extend from the aft surface of the rotor when the retaining ring is inserted between the aft tab of the plurality of turbine blades and the rotor, and wherein the upper peripheral edge of the retaining ring does not contact an inner diameter surface of the plurality of aft tabs that are located radially above the retaining ring when the retaining ring is inserted between the aft tab of the plurality of turbine blades and the rotor.

12. The gas turbine engine as in claim 11, wherein the plurality of tabs are integrally formed with the rotor.

13. The gas turbine engine as in claim 11, wherein the retaining ring further comprises an anti-rotation feature that is located between a pair of aft tabs of adjacent turbine blades secured to the rotor.

14. The gas turbine engine as in claim 13, wherein the anti-rotation feature protrudes from an aft surface of the retaining ring.

15. The gas turbine engine as in claim 14, wherein the retaining ring is provided with a split and wherein the split is located radially below one of plurality of tabs that extends from the aft surface of the rotor.

16. The gas turbine engine as in claim 11, wherein a distal end of the forward tab extends radially past the root portion of the turbine blade it is secured to.

17. The gas turbine engine as in claim 16, wherein the aft tab extends axially from the aft surface of the root portion to define the gap.

18. The gas turbine engine as in claim 11, wherein the aft tab extends axially from the aft surface of the root portion to define the gap.

19. A method of securing a turbine blade to a rotor of a gas turbine engine, comprising:

sliding a root portion of a plurality of turbine blades into a corresponding slot of the rotor, each of the plurality of turbine blades having a forward tab that extends radially downward from the turbine blade and an aft tab that extends from the root portion to define a gap between a surface of the root portion and the aft tab;

contacting a front face of the rotor with the forward tab;

engaging a plurality of tabs that extend from an aft surface of the rotor and the aft tab of the plurality of turbine blades with a retaining ring, wherein an upper peripheral edge of the retaining ring contacts a surface of the plurality of tabs that extend from the aft surface of the rotor when the retaining ring is inserted between the aft tab of the plurality of turbine blades and the rotor, and wherein the upper peripheral edge of the retaining ring does not contact an inner diameter surface of the plurality of aft tabs that are located radially above the retaining ring when the retaining ring is inserted between the aft tab of the plurality of turbine blades and the rotor.

20. The method as in claim 19, further comprising locating an anti-rotation feature between a pair of aft tabs of adjacent turbine blades secured to the rotor.