MASS OFFSET FOR DAMPING PERFORMANCE

• US 20150125304A1
• Filed: 03/11/2014
• Published: 05/07/2015
• Est. Priority Date: 03/13/2013
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Rapp, Brandon M., Hough, Matthew Andrew
• Granted Patent: US 10,641,109 B2
• Time in Patent Office: Days
• US Class Current: 416/190
• CPC Class Codes: F01D 11/008 by spacer elements between the blades, e.g. independent interblade platforms
• CPC Class Codes: F01D 5/22 Blade-to-blade connections, e.g. for damping vibrations
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A gas turbine engine rotor assembly includes a plurality of blades spaced apart from each other for rotation about an axis. Each of the blades includes a platform having an inner surface and an outer surface. Adjacent platforms are separated from each other by a gap. A damper is positioned at an offset position within the gap to contact the inner surfaces of the adjacent platforms.

Claims

1. A gas turbine engine rotor assembly comprising:

a plurality of blades spaced apart from each other for rotation about an axis, each of the blades including a platform having an inner surface and an outer surface, and wherein adjacent platforms are separated from each other by a gap; and

a damper positioned at an offset position within the gap to contact the inner surfaces of the adjacent platforms.

2. The gas turbine engine rotor assembly according to claim 1, wherein adjacent blades define a blade-to-blade centerline, and wherein the damper is offset relative to the blade-to-blade centerline.

3. The gas turbine engine rotor assembly according to claim 2, wherein the damper is defined by a center of mass, and wherein the center of mass is offset from the blade-to-blade centerline.

4. The gas turbine engine rotor assembly according to claim 1, wherein the adjacent platforms are associated with adjacent blades, and wherein the damper provides a load split of approximately 60% for one of the adjacent blades and approximately 40% for the other of the adjacent blades.

5. The gas turbine engine rotor assembly according to claim 1, wherein the damper is positioned closer to one of the adjacent platforms than the other platform.

6. The gas turbine engine according to claim 1, including a rotor disk that supports the plurality of blades, and wherein each blade includes a root that fits into a corresponding slot in the rotor disk.

7. A gas turbine engine rotor assembly comprising:

a rotor disk defining an axis of rotation;

a plurality of blades supported by the rotor disk and spaced apart from each other for rotation about the axis of rotation, the plurality of blades including at least a first blade having a first platform and a second blade having a second platform separated from the first platform by a gap, and wherein each of the first and second platforms includes an inner surface and an outer surface; and

a damper positioned at an offset position within the gap such that the damper provides an unequal load split between the first and second blades.

8. The gas turbine engine rotor assembly according to claim 7, wherein the damper provides a load split of approximately 60% for the first blade and approximately 40% for the second blade.

9. The gas turbine engine rotor assembly according to claim 7, wherein the first and second blades define a blade-to-blade centerline, and wherein the damper is offset relative to the blade-to-blade centerline.

10. The gas turbine engine rotor assembly according to claim 9, wherein the damper is defined by a center of mass, and wherein the center of mass is offset from the blade-to-blade centerline.

11. The gas turbine engine rotor assembly according to claim 10, wherein the damper provides a load split of approximately 60% for the first blade and approximately 40% for the second blade.

12. The gas turbine engine rotor assembly according to claim 9, wherein the damper is positioned closer to one of the adjacent platforms than the other platform.

13. The gas turbine engine rotor assembly according to claim 7, wherein the gap is defined by a circumferential width and a radial depth, and wherein the damper is defined by a center of mass that is positioned circumferentially closer to one of the first and second platforms than the other of the first and second platforms.

14. A gas turbine engine comprising:

a compressor section; and

a turbine section downstream of the compressor section, the turbine section including at least one turbine stage having a rotor disk defining an axis of rotation;

a plurality of blades supported by the rotor disk and spaced apart from each other for rotation about the axis of rotation, the plurality of blades including at least a first blade having a first platform and a second blade having a second platform separated from the first platform by a gap, and wherein each of the first and second platforms includes an inner surface and an outer surface; and

a damper positioned at an offset position within the gap to contact the inner surface of the first platform and the inner surface of the second platform.

15. The gas turbine engine according to claim 14 wherein the damper provides an unequal load split between the first and second blades.

16. The gas turbine engine rotor assembly according to claim 15, wherein the damper provides a load split of approximately 60% for the first blade and approximately 40% for the second blade.

17. The gas turbine engine according to claim 15, wherein the first and second blades define a blade-to-blade centerline, and wherein the damper is offset relative to the blade-to-blade centerline.

18. The gas turbine engine according to claim 17, wherein the damper is defined by a center of mass, and wherein the center of mass is offset from the blade-to-blade centerline.

19. The gas turbine engine according to claim 18, wherein the damper provides a load split of approximately 60% for the first blade and approximately 40% for the second blade.

20. The gas turbine engine according to claim 14, further comprise a fan section, wherein the turbine section drives the fan section via a geared architecture.