High pressure rotor disk

Patent metadata

Field	Value
Publication number	US-RE49382-E
Application number	US-202017011472-A
Country	US
Kind code	E1
Filing date	Sep 3, 2020
Priority date	Sep 28, 2012
Publication date	Jan 24, 2023
Grant date	Jan 24, 2023

Abstract

A rotor disk for a gas turbine engine is disclosed and formed to enable operation at high rotational speeds in a high temperature environment. The rotor disk is formed to include a bore, a live rim diameter and an outer diameter related to each other according to defined relationships.

First claim

What is claimed is:

1. A gas turbine engine comprising:

a compressor section including a high pressure compressor and a low pressure compressor;
a combustor in fluid communication with the compressor section;
a turbine section in fluid communication with the combustor, wherein the turbine section includes a high pressure turbine driving the high pressure compressor and a low pressure turbine driving the low pressure compressor, wherein at least one of the high pressure turbine and the high pressure compressor includes a disk having a bore diameter (D) related to a bore width (W) according to a ratio (D/W) between 1.25 and 1.65.

2. The gas turbine engine as recited in claim 1 , wherein the ratio (DAY) is between 1.35 and 1.55.

3. The gas turbine engine as recited in claim 1 , wherein the ratio (D/W) is 1.45.

4. The gas turbine engine as recited in claim 1 A gas turbine engine comprising:
a compressor section including a high pressure compressor and a low pressure compressor;
a combustor in fluid communication with the compressor section; and
a turbine section in fluid communication with the combustor, wherein the turbine section includes a high pressure turbine driving the high pressure compressor and a low pressure turbine driving the low pressure compressor, wherein at least one of the high pressure turbine and the high pressure compressor includes a disk having a bore diameter (D) related to a bore width (W) according to a ratio (D/W) between 1.25 and 1.65, and wherein the disk includes an outer diameter (OD) related to the bore diameter (D) according to a ratio (OD/D) that is between 2.95 and 3.25.

5. The gas turbine engine as recited in claim 4 , wherein the ratio (OD/D) is between 3.04 and 3.20.

6. The gas turbine engine as recited in claim 4 , wherein the ratio (OD/D) is 3.15.

7. The gas turbine engine as recited in claim 1 A gas turbine engine comprising:
a compressor section including a high pressure compressor and a low pressure compressor;
a combustor in fluid communication with the compressor section; and
a turbine section in fluid communication with the combustor, wherein the turbine section includes a high pressure turbine driving the high pressure compressor and a low pressure turbine driving the low pressure compressor, wherein at least one of the high pressure turbine and the high pressure compressor includes a disk having a bore diameter (D) related to a bore width (W) according to a ratio (D/W) between 1.25 and 1.65, and wherein the disk includes a live rim diameter (d) related to the bore diameter (D) according to a ratio (d/D) that is between 2.25 and 3.00.

8. The gas turbine engine as recited in claim 7 , wherein the ratio (d/D) is between 2.50 and 2.75.

9. The gas turbine engine as recited in claim 7 , wherein the ratio (d/D) is 2.69.

10. A rotor disk for a gas turbine engine comprising:

an outer diameter (OD) related to a bore diameter (D) according to a ratio (OD/D) that is between 2.95 and 3.25.

11. The rotor disk as recited in claim 10 , wherein the ratio (OD/D) is between 3.04 and 3.20.

12. The rotor disk as recited in claim 10 , wherein the ratio (OD/D) is 3.15.

13. The rotor disk as recited in claim 10 , wherein the bore diameter (D) is related to a bore width (W) according to a ratio (D/W) between 1.25 and 1.65.

14. The rotor disk as recited in claim 13 , wherein the ratio (D/W) is between 1.53 and 1.55.

15. The rotor disk as recited in claim 13 , wherein the ratio (DAY) is 1.45.

16. The rotor disk as recited in claim 10 , wherein the disk includes a live rim diameter (d) related to the bore diameter (D) according to a ratio (d/D) that is between 2.25 and 3.00.

17. The rotor disk as recited in claim 16 , wherein the ratio (d/D) is between 2.50 and 2.75.

18. The rotor disk as recited in claim 16 , wherein the ratio (d/D) is 2.69.

19. A method of fabricating a rotor disk for a gas turbine engine comprising;

forming a bore including a bore diameter (D) and a live rim diameter (d) with a ratio (d/D) of the live rim diameter (d) to the bore diameter (D) being between 2.25 and 3.00
forming at least one lug for mounting a blade at the live rim diameter (d); and
forming an outer diameter (OD).

20. The method as recited in claim 19 , including forming the disk to include a ratio (OD/D) of the outer diameter (OD) to the bore diameter (D) between 2.95 and 3.25.

21. The method as recited in claim 19 , including forming a bore including a bore diameter (D) and a bore width (W) in a direction parallel to an axis of intended rotation, wherein the bore diameter (D) is related to the bore width (W) according to a ratio (D/W) that is between 1.25 and 1.65.

22. The gas turbine engine as set forth in claim 4, wherein the ratio (D/W) is between 1.35 and 1.55.

23. The gas turbine engine as recited in claim 4, wherein the ratio (D/W) is 1.45.

24. The gas turbine engine as set forth in claim 7, wherein the ratio (D/W) is between 1.35 and 1.55.

25. The gas turbine engine as recited in claim 7, wherein the ratio (D/W) is 1.45.

26. A gas turbine engine, comprising:
a high speed shaft mounted for rotation about a longitudinal axis; a low speed shaft coupled to a fan of a fan section, the low speed shaft mounted for rotation concentrically with the high speed shaft, the low speed shaft configured to rotate at a lower speed than the high speed shaft; a rotor disk of a turbine attached to the high speed shaft, the rotor disk of the turbine having a first bore diameter (D 1 ) related to a first bore width (W 1 ) according to a first ratio (D 1 /W 1 ) between 1.25 and 1.65; and a rotor disk of a compressor attached to the high speed shaft, the rotor disk of the compressor having a second bore diameter (D 2 ) related to a compressor outer diameter (OD 2 ) according to a second ratio (OD 2 /D 2 ) between 2.95 and 3.25;
wherein the fan has less than 26 fan blades; and
further comprising a speed change device, wherein a low pressure turbine attached to the low speed shaft is configured to drive the fan through the speed change device at a lower speed than the low speed shaft.

27. The gas turbine engine as recited in claim 26, wherein the fan section has a low fan pressure ratio measured across the fan blades alone of less than 1.45 at a flight condition of 0.8 Mach and 35,000 ft., wherein the fan section has a low corrected fan tip speed less than 1,150 ft./sec at the flight condition of 0.8 Mach and 35,000 ft., and wherein the low corrected fan tip speed is an actual fan tip speed divided by ((Tram ° R)/(518.7° R)) 0.5 .

28. The gas turbine engine as set forth in claim 26, wherein the turbine is a two stage turbine.

29. The gas turbine engine as set forth in claim 26, wherein the ratio (D 1 /W 1 ) is between 1.35 and 1.55.

30. The gas turbine engine as recited in claim 26, wherein the ratio (D 1 /W 1 ) is 1.45.

31. The gas turbine engine as recited in claim 26, wherein the ratio (OD 2 /D 2 ) is between 3.04 and 3.20.

32. The gas turbine engine as recited in claim 26, wherein the ratio (OD 2 /D 2 ) is 3.15.

33. A gas turbine engine, comprising:
a high speed shaft moun

Assignees

• Raytheon Tech Corp

Inventors

• Virkler Scott D

Classifications

• F05D2240/24

  for turbines · CPC title

• F01D5/3007 Primary

  of axial insertion type · CPC title

• F05D2200/00

  Mathematical features · CPC title

• F05D2240/20

  Rotors · CPC title

• F01D5/02 Primary

  Blade-carrying members, e.g. rotors (rotors of non-bladed type F01D1/34; stators F01D9/00 {; selecting particular materials F01D5/28}) · CPC title