Integrated circuit cooled turbine blade

Patent metadata

Field	Value
Publication number	US-9745853-B2
Application number	US-201514840572-A
Country	US
Kind code	B2
Filing date	Aug 31, 2015
Priority date	Aug 31, 2015
Publication date	Aug 29, 2017
Grant date	Aug 29, 2017

Abstract

A turbine rotor blade includes at least two integrated cooling circuits that are formed within the blade that include a leading edge circuit having a first cavity and a second cavity and a trailing edge circuit that includes at least a third cavity located aft of the second cavity. The trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity. The last cavity is located along a trailing edge of the blade. A tip axial cooling channel connects to the first cavity of the leading edge circuit and the penultimate cavity of the trailing edge circuit. At least one crossover hole connects the penultimate cavity to the last cavity substantially near the tip end of the blade.

First claim

What is claimed is:

1. A turbine rotor blade comprising:
a leading edge and a trailing edge joined by a pressure side and a suction side, a tip end, and a root end;
at least two cooling circuits formed within the blade to provide cooling for the blade comprising;
a leading edge circuit comprising a first cavity located along the leading edge of the blade and a second cavity positioned aft of the first cavity in an axial direction, wherein the leading edge circuit only flows forward so that the second cavity impinges forward directly into the first cavity;
a trailing edge circuit comprising at least a third cavity located in a mid-chord area of the blade aft of the second cavity, wherein the trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity, wherein the last cavity is located along a trailing edge of the blade; and
a tip axial cooling channel comprising a first opening and a second opening, wherein the first opening connects to the first cavity and the second opening connects to the penultimate cavity, wherein the tip axial cooling channel connects the leading edge circuit to the trailing edge circuit, wherein the at least two cooling circuits are integrated only through the first opening and the second opening of the tip axial cooling channel connection; and
at least one crossover hole connecting the penultimate cavity to the last cavity substantially near the tip end of the blade.

2. The turbine rotor blade according to claim 1 , wherein the last cavity further comprises a plurality of trailing edge pins and/or a plurality of trailing edge exit holes along the trailing edge of the blade.

3. The turbine rotor blade according to claim 2 , wherein each of the plurality of trailing edge pins comprises an airfoil shape.

4. The turbine rotor blade according to claim 1 , further comprises an inboard squealer tip along the tip end of the blade.

5. The turbine rotor blade according to claim 1 , further comprises broken offset turbulators along the last cavity and broken staggered turbulators along cavities forward of the last cavity.

6. The turbine rotor blade according to claim 1 , further comprising a plurality of helical mini-grooves along the leading edge of the blade.

7. The turbine rotor blade according to claim 1 , wherein the trailing edge circuit is a 3-pass serpentine cooling circuit.

8. The turbine rotor blade according to claim 1 , wherein the trailing edge circuit is a 5-pass serpentine cooling circuit.

9. The turbine rotor blade according to claim 1 , wherein the trailing edge circuit is a 7-pass serpentine cooling circuit.

10. A method for increasing cooling to a trailing edge tip corner of a turbine blade, comprising:
providing a tip axial cooling channel comprising a first opening and a second opening);
connecting the first opening of the tip axial cooling channel to an end of a first cavity of a leading edge circuit of at least two cooling circuits formed within the turbine blade, wherein the leading edge circuit comprises a first cavity located along a leading edge of the blade and a second cavity positioned aft of the first cavity in an axial direction, wherein the leading edge circuit only flows forward so that the second cavity impinges forward directly into the first cavity;
connecting the second opening of the tip axial cooling channel to a trailing edge circuit, wherein the trailing edge circuit comprises at least a third cavity located in a mid-chord area of the blade aft of the second cavity, wherein the trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity, wherein the last cavity is located along a trailing edge of the blade, wherein the at least two integrated cooling circuits further comprise at least one crossover hole connecting the penultimate cavity to the last cavity substantially near the tip end of the blade, wherein the at least two cooling circuits are integrated only through the first opening and the second opening of the tip axial cooling channel connection;
sending cooling air through the second cavity of the leading edge circuit and the third cavity of the trailing edge circuit, wherein the cooling air flowing through the leading edge circuit then flows through the tip axial cooling channel and into the penultimate cavity of the trailing edge circuit, merging with the cooling air entering into the penultimate cavity in the trailing edge circuit, wherein a portion of the cooling air flows through the at least one crossover hole into the trailing edge tip corner, and a portion of the cooling air flows through the rest of the penultimate cavity into and up through the last cavity through the rest of the trailing edge circuit into the trailing edge tip corner and/or out through the trailing edge of the turbine blade.

11. The method according to claim 10 , further comprising introducing refresher air into the last cavity along the root end of the blade.

12. The method according to claim 10 , wherein the last cavity further comprises a plurality of trailing edge pins and/or a plurality of trailing edge exit holes along the trailing edge of the blade.

13. The method according to claim 12 , wherein each of the plurality of trailing edge pins comprise an airfoil shape.

14. The method according to claim 10 , further comprises an inboard squealer tip along the tip end of the blade.

15. The method according to claim 10 , further comprises broken offset turbulators along the last cavity and broken staggered turbulators along cavities forward of the last cavity.

16. The method according to claim 10 , further comprising a plurality of helical mini-grooves along the leading edge of the blade.

17. The method according to claim 10 , wherein the trailing edge circuit is a 3-pass serpentine cooling circuit.

18. The method according to claim 10 , wherein the trailing edge circuit is a 5-pass serpentine cooling circuit.

19. The method according to claim 10 , wherein the trailing edge circuit is a 7-pass serpentine cooing circuit.

Assignees

• Siemens Energy Inc

Inventors

• Lee Ching-Pang
• Jiang Nan
• Um Jae Y
• Holloman Harry
• Koester Steven

Classifications

• F05D2260/2212

  by creating turbulence · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F05D2240/304

  related to the trailing edge of a rotor blade · CPC title

• F05D2260/202

  by film cooling · CPC title

• F01D5/147

  Construction, i.e. structural features, e.g. of weight-saving hollow blades (F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18) · CPC title