Turbine circumferential dovetail leakage reduction

Patent metadata

Field	Value
Publication number	US-11486261-B2
Application number	US-202016881895-A
Country	US
Kind code	B2
Filing date	May 22, 2020
Priority date	Mar 31, 2020
Publication date	Nov 1, 2022
Grant date	Nov 1, 2022

Abstract

Methods, apparatus, systems and articles of manufacture are disclosed for a compressor including a rotor defining a circumferential direction, wherein the rotor includes a slot, the slot including a first neck portion, a first blade and a second blade disposed circumferentially apart in the slot, and a block disposed in the slot circumferentially between the first blade and the second blade, the block including second neck portion, the first neck portion to at least partially interface the second neck portion.

First claim

What is claimed is:

1. A compressor comprising:
a rotor defining a circumferential direction, wherein the rotor includes a slot, the slot including a first neck portion defining axial flanges;
a first blade and a second blade disposed circumferentially apart in the slot, each of the first and second blades including a platform and a dovetail;
an integral block disposed in the slot circumferentially between the first blade and the second blade, the block including a second neck portion to interface the platform and the dovetail of the first blade and to interface the platform and the dovetail of the second blade;
the block including protrusions extending radially inward at radially outer axial ends of the block, the protrusions to be received by seal glands; and
the block including fore and aft faces, the fore face in a first contact with a first axial face of the axial flanges, the aft face in a second contact with a second axial face of the axial flanges, the first and second contacts to reduce airflow leakage,
wherein, during operation of the compressor, the block is to move radially outward to couple the block to the axial flanges of a dovetail receiving portion of the slot.

2. The compressor of claim 1 , wherein the compressor is a high-pressure, multi-stage compressor.

3. The compressor of claim 1 , wherein the dovetail portions of the first and second blades include second protrusions defining a volumetric space circumferentially therebetween.

4. The compressor of claim 3 , wherein the dovetail portion occupies at least a portion of the volumetric space to reduce air flow leakage into the slot.

5. The compressor of claim 4 , wherein the block is hollow.

6. A gas turbine comprising:
a compressor rotor including a slot, the slot including first and second axial flanges;
a first blade and a second blade disposed in the slot, each of the first and second blades including a platform and a dovetail;
an integral block disposed in the slot between the first and second blades, the block to interface the platform and the dovetail of the first blade and the platform and the dovetail of the second blade;
the block including protrusions extending radially inward at radially outer axial ends of the block, the protrusions to be received by seal glands; and
the block including fore and aft faces, the fore face in contact with an axial face of the first axial flange and the aft face in contact with an axial face of the second axial flange to prevent airflow from bypassing the block,
wherein, during operation of the gas turbine, the block is to move radially outward to couple the block to the axial flanges of a dovetail receiving portion of the slot.

7. The gas turbine of claim 6 , wherein the gas turbine is of a propulsion system of an aircraft.

8. The gas turbine of claim 6 , wherein the slot includes first and second grooves and the block includes first and second protrusions, the first groove to receive the first protrusion and the second groove to receive the second protrusion.

9. The gas turbine of claim 6 , wherein the block includes a platform portion to interface with the platforms of the first and second blades, the platform portion including an opening.

10. The gas turbine of claim 9 , wherein the gas turbine defines axial and radial directions, an axially aft portion of the platform portion disposed radially outward from an axially fore portion of the platform portion.

11. The gas turbine of claim 6 , wherein at least a portion of the block is hollow.

12. An apparatus comprising:
a rotor defining a radial direction and a circumferential direction, the rotor including a slot with a neck defining axial flanges;
a first blade and a second blade disposed in the slot, the first blade having a first protrusion and a first platform, and the second blade having a second protrusion and a second platform;
an integral block disposed circumferentially between the first blade and the second blade in the slot, the block to interface i) the first platform and the first protrusion of the first blade and ii) the second platform and the second protrusion of the second blade,
the block including third protrusions extending radially inward at radially outer axial ends of the block, the third protrusions to be received by seal glands; and
the block including fore and aft faces, the fore face in contact with a first axial face of a first axial flange of the axial flanges and the aft face in contact with a second axial face of a second axial flange of the axial flanges to reduce airflow bypass,
wherein, during operation of the apparatus, the block is to move radially outward to be radially retained by a dovetail receiving portion of the axial flanges of the neck.

13. The apparatus of claim 12 , wherein the block is hollow.

14. The apparatus of claim 12 , wherein edges of the block are substantially aligned with edges of the first and second blades.

15. The apparatus of claim 12 , wherein the platforms of the first and second blades extend a greater length in the circumferential direction then do the protrusions of the first and second blades.

16. The apparatus of claim 15 , wherein the block includes an opening.

17. The apparatus of claim 12 , wherein the block is manufactured using an additive manufacturing process.

Assignees

• Gen Electric

Inventors

• N Arvind

Classifications

• F01D5/3038 Primary

  the slot having inwardly directed abutment faces on both sides · CPC title

• F04D29/322

  Blade mountings · CPC title

• F02C3/064

  the compressor having concentric stages · CPC title

• F01D5/32

  Locking, e.g. by final locking blades or keys · CPC title

• F05D2220/3219

  for the last stage of a compressor or a high pressure compressor · CPC title