Blade damper

• US 5,226,784 A
• Filed: 09/25/1992
• Issued: 07/13/1993
• Est. Priority Date: 02/11/1991
• Status: Expired due to Term

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Park, Sang Y., Wines, Daniel E., Mueller, Peter W., Glynn, Christopher C., Stevenson, Joseph T., Panovsky, Josef
• Primary Examiner(s): KWON, JOHN
• Application Number:
  US07/951,929
• Time in Patent Office: 291 Days
• Field of Search: 416/248, 416/500
• US Class Current: 416/248
• CPC Class Codes: F01D 11/006 Sealing the gap between rotor blades or blades and rotor
• CPC Class Codes: F01D 5/22 Blade-to-blade connections, e.g. for damping vibrations
• CPC Class Codes: Y10S 416/50 Vibration damping features

Abstract

The present invention provides a phase independent damper and damper assembly capable of damping rotor blade vibrations in the axial, circumferential, and radial directions. One embodiment, particularly useful in an aircraft gas turbine engine compressor rotor, provides a damper and damper assembly including a generally axially extending blade damper operable to fictionally engage a generally axially extending circumferentially facing surface of a disk and axially extending angled surface under the platform of the blade. The invention provides an additional advantage of allowing simple and easy modification of existing engines to incorporate the damper of the present invention.

Claims

1. A gas turbine engine rotor blade damper for loosely mounting in a recess of a disk having an axis of rotation, said recess forming a chamber between the disk and blade platform of a blade extending radially outward from the disk, said damper comprising:

a generally axially extending body including at least one radially extending portion having a right angle triangular cross-section wherein at least one leg of said triangular cross-section defines a friction damping surface;

whereby during engine operation centrifugal force seats said damper against said disk and said blade platform.

2. A gas turbine engine rotor blade damper as claimed in claim 1, wherein said body includes a second radially extending portion that depends from said first portion and provides a means for adjusting the mass of the damper.

3. A gas turbine engine rotor blade to ground damper assembly comprising:

a blade having a blade root disposed in a blade slot, said slot formed in a disk between adjacent posts,

a generally axially extending chamber formed by a recess in one of said posts between said disk and said blade, said chamber operable to loosely contain a generally axially extending damper,

said damper having generally axially extending surfaces conforming to generally axially extending surfaces on said blade and disk, and

said damper operable to engage said disk and blade under the centrifugal force due to its mass.

4. A gas turbine engine rotor blade damper assembly comprising:

a generally axially extending chamber between a disk and a blade formed by a generally axially extending recess in a inter-slot post on the disk and a generally axially extending radially angled and radially inward facing of a rotor blade platform, said chamber operable to loosely contain a generally axially extending damper including at least one radially extending portion having a right angle triangular cross-section;

whereby during engine operation the mass of said damper centrifugally forces said damper into engagement with the disk and the blade platform.

5. A gas turbine engine rotor blade to ground damper assembly comprising:

a blade having a blade root disposed in a blade slot, said slot formed in a root between adjacent posts;

an axially extending blade vibration dampening means to dampen rotor blade vibrations independent of the blade vibration'"'"'s phase wherein said blade vibration dampening means includes a damper having planar frictional surfaces, said damper loosely disposed in a chamber formed by a recess in one of said posts between said blade and said rotor;

whereby during engine operation said damper is operable due to centrifugal force.

6. A gas turbine engine rotor blade to ground damper assembly as claimed in claim 5 wherein said blade vibration dampening means comprises a generally axially extending and circumferentially facing planar rotor surface on said rotor, a generally axially extending and generally circumferentially and radially inward facing planar blade surface of said blade, wherein said rotor surface generally faces said blade surface forming at least a portion of said chamber therebetween.

7. A gas turbine engine rotor blade to ground damper assembly as claimed in claim 6 wherein said damper has at least one radially extending portion having a triangular cross-section including a radially outward facing and generally circumferentially opposite facing planar damper surface facing and and conforming to and said planar blade surface.

8. A gas turbine engine rotor blade to ground damper assembly as claimed in claim 7 wherein said blade surface is disposed on the underside of a platform of said blade.