Rotor and gas turbine engine including same

Patent metadata

Field	Value
Publication number	US-10082034-B2
Application number	US-201514755952-A
Country	US
Kind code	B2
Filing date	Jun 30, 2015
Priority date	Jul 3, 2014
Publication date	Sep 25, 2018
Grant date	Sep 25, 2018

Abstract

A rotor for a gas turbine engine includes a plurality of blades which extend from a rotor disk, adjacent ones of the plurality of blades are joined by a flexible web.

First claim

What is claimed:

1. A rotor for a gas turbine engine comprising:
a rotor disk defined along an axis of rotation;
a plurality of cooling airflow passages extending axially through the rotor disk;
a plurality of blades integrally formed with and extending from said rotor disk; and
a plurality of flexible webs integrally formed with the rotor disk, each of said plurality of flexible webs extending between two of said plurality of blades, wherein the plurality of flexible webs provide a seal to the plurality of cooling airflow passages.

2. The rotor of claim 1 , wherein each of said plurality of flexible webs comprises:
a first arm including a first arm proximal end operatively coupled to a first one of the plurality of blades, the first arm further including a first arm distal end; and
a second arm including a second arm proximal end operatively coupled to a second one of the plurality of blades, the second arm further including a second arm distal end operatively coupled to the first arm distal end.

3. The rotor of claim 2 , further comprising a radiused joint operatively coupling the first arm distal end to the second arm distal end.

4. The rotor of claim 2 , wherein the first arm and the second arm each extend radially inboard.

5. The rotor of claim 1 , wherein each of said plurality of blades includes a platform section.

6. The rotor of claim 5 , wherein each of said plurality of flexible webs comprises:
a first arm including a first arm proximal end operatively coupled to a first platform section of a first one of the plurality of blades, the first arm further including a first arm distal end; and
a second arm including a second arm proximal end operatively coupled to a second platform section of a second one of the plurality of blades, the second arm further including a second arm distal end operatively coupled to the first arm distal end.

7. The rotor of claim 6 , further comprising a radiused joint operatively coupling the first arm distal end to the second arm distal end.

8. The rotor of claim 6 , wherein the first arm and the second arm each extend radially inboard.

9. The rotor of claim 1 , wherein said rotor disk is manufactured of a first material and said plurality of blades are manufactured of a second material, said first material different than said second material.

10. The rotor of claim 1 , wherein said plurality of blades are high pressure compressor blades.

11. The rotor of claim 1 , wherein said rotor disk and said plurality of blades comprise an integrally bladed rotor.

12. A spool for a gas turbine engine comprising:
a compressor rotor disk defined along an axis of rotation;
a plurality of cooling airflow passages extending axially through the compressor rotor disk;
a plurality of compressor blades integrally formed with and extending from said compressor rotor disk; and
a first plurality of flexible webs integrally formed with the rotor disk, each of said first plurality of flexible webs extending between two of said plurality of compressor blades, wherein the first plurality of flexible webs provide a seal to the plurality of cooling airflow passages.

13. The spool of claim 12 , wherein each of said first plurality of flexible webs comprises:
a first arm including a first arm proximal end operatively coupled to a first one of the plurality of compressor blades, the first arm further including a first arm distal end; and
a second arm including a second arm proximal end operatively coupled to a second one of the plurality of compressor blades, the second arm further including a second arm distal end operatively coupled to the first arm distal end.

14. The spool of claim 13 , further comprising a radiused joint operatively coupling the first arm distal end to the second arm distal end.

15. The spool of claim 13 , wherein the first arm and the second arm each extend radially inboard.

16. The spool of claim 13 , wherein each of said plurality of compressor blades includes a platform section, the first arm proximal end operatively coupled to a first platform section of the first one of the plurality of compressor blades, and the second arm proximal end operatively coupled to a second platform section of the second one of the plurality of compressor blades.

17. The spool of claim 12 , wherein said compressor rotor disk is manufactured of a first material and said plurality of compressor blades are manufactured of a second material, said first material different than said second material.

18. The spool of claim 12 , wherein said plurality of compressor blades are high pressure compressor blades.

19. The spool of claim 12 , wherein said compressor rotor disk and said plurality of compressor blades comprise an integrally bladed rotor.

20. The spool of claim 12 , further comprising:
a turbine rotor disk defined along said axis of rotation;
a plurality of cooling airflow passages extending axially through the turbine rotor disk; and
a plurality of turbine blades integrally formed with and extending from said turbine rotor disk; and
a second plurality of flexible webs integrally formed with the turbine rotor disk, each of said second plurality of flexible webs extending between two of said plurality of turbine blades, wherein the second plurality of flexible webs provide a seal to the plurality of cooling airflow passages of the turbine rotor disk.

Assignees

• United Technologies Corp

Inventors

• Merry Brian D
• Suciu Gabriel L
• Hill James D

Classifications

• F05D2240/24

  for turbines · CPC title

• F01D5/34

  Rotor-blade aggregates of unitary construction {, e.g. formed of sheet laminae; (discs formed of sheet laminae F01D5/028; ceramic materials F01D5/284, composite materials F01D5/282)} · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F01D5/225 Primary

  by shrouding · CPC title

• F05D2260/941

  particularly aimed at mechanical or thermal stress reduction · CPC title