Cooled turbine blade with leading edge flow redirection and diffusion

Patent metadata

Field	Value
Publication number	US-9228439-B2
Application number	US-201213631133-A
Country	US
Kind code	B2
Filing date	Sep 28, 2012
Priority date	Sep 28, 2012
Publication date	Jan 5, 2016
Grant date	Jan 5, 2016

Abstract

A cooled turbine blade having a base and an airfoil, the base including cooling air inlet and an internal cooling air passageway, and the airfoil including an internal heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a “skin” that encompasses a tip wall, an inner spar, a leading edge rib, and a leading edge air deflector. The leading edge rib is configured to form a leading edge chamber in conjunction with the leading edge of the skin. The leading edge air deflector is shaped and positioned such that cooling air leaving the leading edge chamber is both turned and diffused.

First claim

What is claimed is:

1. A turbine blade for use in a gas turbine engine, the turbine blade comprising:
a base;
an airfoil comprising a skin extending from the base and forming a first leading edge, a first trailing edge, a pressure side, and a lift side, the airfoil having tip end distal from the base;
a leading edge rib extending from the pressure side of the skin to the lift side of the skin, the leading edge rib extending from the base and terminating prior to reaching the tip end, the leading edge rib forming a leading edge chamber in conjunction with the first leading edge of the skin, the leading edge chamber extending from the base towards the tip end;
an inner spar extending between the base and the tip end, the inner spar located between the pressure side of the skin and the lift side of the skin, and further extending from the leading edge rib toward the trailing edge;
a section divider extending towards the first leading edge and between the inner spar and the skin, the section divider being offset from the tip wall and forming a heat exchange path therebetween, the heat exchange path extending along the tip wall towards the first trailing edge; and
a leading edge air deflector extending from the pressure side of the skin to the lift side of the skin and having a second leading edge, a second trailing edge, a turning side and a diffusion side, the leading edge air deflector located at least partially between the leading edge rib and the tip end, the leading edge deflector positioned with an inner gap between the leading edge air deflector and the leading edge rib, and an outer gap between the leading edge air deflector and both the skin of the airfoil and the tip end, the leading edge air deflector curving from the second leading edge to the second trailing edge and intersects the inner spar at the second trailing edge to turn cooling air passing through the inner gap and the outer gap towards the first trailing edge and to direct the cooling air into the heat exchange path.

2. The turbine blade of claim 1 , wherein a width of the outer gap is greater at the second trailing edge than at the second leading edge.

3. The turbine blade of claim 1 , wherein the width of the outer gap is greater at the second trailing edge than at the second leading edge by a ratio of at least 1 to 4.5.

4. The turbine blade of claim 1 , wherein the turning side of the leading edge air deflector forms a concave curve; and
wherein the diffusion side of the leading edge air deflector forms a convex curve.

5. The turbine blade of claim 4 , wherein the diffusion side of the leading edge air deflector is smoothly contoured such that there is no more than two percent pressure drop associated with separation losses from the diffusion side between the second leading edge and the second trailing edge.

6. The turbine blade of claim 1 , wherein the leading edge air deflector has a maximum aerodynamic thickness between 1.0 and 2.0 times the thickness of the skin of the airfoil.

7. The turbine blade of claim 1 , wherein the second leading edge has a leading edge radius;
wherein the leading edge air deflector has a maximum aerodynamic thickness; and
wherein the maximum aerodynamic thickness is between 1.2 to 1.4 times the leading edge radius.

8. The turbine blade of claim 1 , wherein the second trailing edge has a leading edge radius;
wherein the leading edge air deflector has a maximum aerodynamic thickness; and
wherein the maximum aerodynamic thickness is between 1.6 to 1.9 times the trailing edge radius.

9. The turbine blade of claim 1 , further comprising
a plurality of first inner spar cooling fins extending from the inner spar to the skin on the lift side of the airfoil, wherein the plurality of first inner spar cooling fins extend from the inner spar with a density of at least 80 fins per square inch; and
a plurality of second inner spar cooling fins extending from the inner spar to the skin on the pressure side of the airfoil, wherein the plurality of second inner spar cooling fins extend from the inner spar with a density of at least 80 fins per square inch.

10. The turbine blade of claim 1 , wherein the turbine blade is cast from a single material.

11. A gas turbine engine including a turbine having a turbine rotor assembly that includes a plurality of turbine blades of claim 1 .

12. The turbine blade of claim 1 , wherein the section divider includes a portion that extends up from the base towards the tip end that transitions to extending towards the first trailing edge, and wherein the heat exchange path extends upward from the base and includes a single-bend that turns the heat exchange path towards the first trailing edge.

13. A turbine blade for use in a gas turbine engine, the turbine blade comprising:
a base;
an airfoil comprising a skin extending from the base and forming a first leading edge, a first trailing edge, a pressure side, and a lift side, the airfoil having tip end distal from the base;
a leading edge rib extending from the pressure side of the skin to the lift side of the skin, the leading edge rib extending from the base and terminating prior to reaching the tip end, the leading edge rib defining a leading edge chamber in conjunction with the first leading edge of the skin, the leading edge chamber extending from the base towards the tip end;
an inner spar extending from the base toward the tip end, the inner spar located between the pressure side of the skin and the lift side of the skin, and further extending from the leading edge rib towards the first trailing edge;
a section divider extending towards the first leading edge and between the inner spar and the skin, the section divider being offset from the tip wall and forming a heat exchange path therebetween, the heat exchange path extending along the tip wall to the first trailing edge; and
a leading edge air deflector extending from the pressure side of the skin to the lift side of the skin and having a second leading edge facing towards the base, a second trailing edge facing towards the first trailing edge, the angle between the directions that the second trailing edge and the second leading edge face is from 80 to 100 degrees, a turning side extending from the second leading edge to the second trailing edge with a concave shape and offset from leading edge rib forming an inner gap there between, and a diffusion side extending from the second leading edge to the second trailing edge with a convex shape, the leading edge deflector configured such that cooling air either passes through the inner gap or along the diffusion side and is directed towards the first trailing edge through the heat exchange path, and wherein the cooling air passing along the diffusion side is diffused by a ratio of at least 1 to 4.5 between the second leading edge and the second trailing edge.

14. A gas turbine engine including a turbine having a turbine rotor assembly that includes a plurality of turbine blades of claim 13 .

15. The turbine blade of claim 13 , wherein the leading edge air deflector intersects the inner spar at the second trailing edge.

16. A gas turbine engine including a turbine having a turbine rotor assembly that includes a plurality of turbine blades, each turbine blade including
a base;
an airfoil comprising a skin extending from the base and forming a first leading edge, a first trailing edge, a pressure side, and a lift side, the airfoil having tip end distal from the base;
a leading edge rib extending from the pressure side of the skin to the lift side of the skin, the leading edge rib extending from the base and termina

Assignees

• Solar Turbines Inc

Inventors

• Pointon Stephen Edward
• Meier Andrew
• Okpara Nnawuihe Asonye
• Luo Jiang

Classifications

• F05D2240/303

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

• F01D5/186 Primary

  Film cooling (F01D5/187 takes precedence) · CPC title

• F01D5/187

  Convection cooling · CPC title

• F01D5/188

  with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall · CPC title

• F01D5/18

  Hollow blades, {i.e. blades with cooling or heating channels or cavities (structure of hollow blades in general F01D5/147)}; Heating, heat-insulating or cooling means on blades · CPC title