COOLED TURBINE BLADE WITH LEADING EDGE FLOW REDIRECTION AND DIFFUSION

• US 20140093390A1
• Filed: 09/28/2012
• Published: 04/03/2014
• Est. Priority Date: 09/28/2012
• Status: Active Grant

• Current Assignee: Solar Turbines Incorporated (Caterpillar Incorporated)
• Original Assignee: Solar Turbines Incorporated (Caterpillar Incorporated)
• Inventors: Pointon, Stephen Edward, Meier, Andrew, Okpara, Nnawuihe Asonye, Luo, Jiang
• Granted Patent: US 9,228,439 B2
• Time in Patent Office: Days
• US Class Current: 416/97.R
• CPC Class Codes: F01D 5/147 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 Class Codes: F01D 5/18 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 Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F01D 5/188 with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
• CPC Class Codes: F05D 2240/303 related to the leading edge of a rotor blade

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.

Claims

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;

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; 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.

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:

at least one cooling air passageway in the base;

a single-bend heat exchange path within the airfoil, the single-bend heat exchange path interfacing with and beginning at the at least one cooling air passageway in the base, and terminating at the first trailing edge, the single-bend heat exchange path configured to redirect the cooling air from the at least one cooling air passageway in the base toward the first trailing edge;

wherein the single-bend heat exchange path is further configured to redirect the cooling air such that the cooling air is redirected in a single turn; and

wherein at least a portion of the single-bend heat exchange path is sub-divided by the inner spar.

10. 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.

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

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

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;

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; 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 deflector configured such that cooling air either passes along the turning side or along the diffusion side to leave the leading edge chamber, 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. The turbine blade of claim 13, wherein the leading edge air deflector is further configured to turn the cooling air between 80 and 100 degrees along the turning side;

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 of the leading edge air deflector; and

wherein the leading edge air deflector is further configured such that the cooling air passing along the turning side rejoins the cooling air passing along the diffusion side downstream of the second trailing edge.

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

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 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;

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 towards the 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 such that an inner gap is made between the leading edge air deflector and the leading edge rib, and an outer gap is made between the leading edge air deflector and both the skin of the airfoil and the tip end, the leading edge air deflector further positioned such that cooling air passes through either the inner gap or the outer gap to leave the leading edge chamber.

17. The gas turbine engine of claim 16, wherein a cross sectional area of the outer gap, normal to air flow, is greater at the second trailing edge than at the second leading edge.

18. The gas turbine engine of claim 17, wherein the cross sectional area of the outer gap, normal to air flow, is greater at the second trailing edge than at the second leading edge by a ratio of at least 1 to 4.5.

19. The gas turbine engine of claim 17, 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 leading edge and the trailing edge of the leading edge air deflector;

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;

wherein the leading edge air deflector has a maximum aerodynamic thickness;

wherein the second leading edge has a leading edge radius;

wherein the second trailing edge has a trailing edge radius;

wherein the maximum aerodynamic thickness is between 1.2 to 1.4 times the leading edge radius; and

wherein the maximum aerodynamic thickness is between 1.6 to 1.9 times the trailing edge radius.

20. The turbine blade of claim 16, further comprising:

a plurality of trailing edge cooling fins extending from the pressure side of the skin to the lift side of the skin;

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.