Air cooled turbine blade

• US 7,988,417 B1
• Filed: 11/19/2007
• Issued: 08/02/2011
• Est. Priority Date: 11/19/2007
• Status: Expired due to Fees

• Current Assignee: Florida Turbine Technologies, Inc. (Kratos Defense & Security Solutions Incorporated)
• Original Assignee: Florida Turbine Technologies, Inc. (Kratos Defense & Security Solutions Incorporated)
• Inventors: Liang, George
• Primary Examiner(s): Kershteyn; Igor
• Application Number:
  US11/986,032
• Time in Patent Office: 1,352 Days
• Field of Search: 416/90.R, 416/96.R, 416/97.R, 415/115
• US Class Current: 416/90.00R
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F01D 5/20 Specially-shaped blade tips to seal space between tips and stator F01D5/225 takes precedence
• CPC Class Codes: F05D 2250/185 serpentine-like
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F05D 2260/2212 by creating turbulence
• CPC Class Codes: F05D 2260/2214 by increasing the heat transfer surface
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A turbine blade with a 5-pass serpentine flow cooling circuit for mid-blade cooling that is formed from a 3-pass forward flowing serpentine circuit followed by a 2-pass aft flowing serpentine circuit with the first leg or supply channel for the circuit located between the second leg and the fourth leg. A crossover channel connects the third leg to the fourth leg and passes underneath the blade tip to provide convective cooling and also discharges cooling air through tip cooling holes.

Claims

1. A turbine blade for a gas turbine engine comprising:

an airfoil having a leading edge and a trailing edge, and a pressure side wall and a suction side wall; and

,

a 5-pass serpentine flow cooling circuit which includes a 3-pass forward flowing serpentine circuit and a 2-pass aft flowing serpentine circuit downstream from the 3-pass serpentine circuit, and the five legs of the serpentine flow cooling circuit being connected in series.

2. The turbine blade of claim 1, and further comprising:

the first leg of the 5-pass serpentine circuit is positioned between the second leg and the fourth leg.

3. The turbine blade of claim 1, and further comprising:

a crossover channel joins the third leg to the fourth leg.

4. The turbine blade of claim 3, and further comprising:

the crossover channel passes underneath the blade tip.

5. The turbine blade of claim 4, and further comprising:

the crossover channel includes a plurality of tip cooling holes to discharge cooling air through the blade tip.

6. The turbine blade of claim 1, and further comprising:

the third leg includes a plurality of film cooling holes to discharge film cooling air onto the pressure side wall or the suction side wall.

7. The turbine blade of claim 1, and further comprising:

the 5-pass serpentine circuit extends substantially along the entire spanwise length of the airfoil from the platform to the tip.

8. The turbine blade of claim 1, and further comprising:

a leading edge cooling supply channel positioned in the leading edge region of the airfoil;

an impingement cavity connected to the leading edge supply channel; and

,

a showerhead arrangement of film cooling holes connected to the impingement cavity.

9. The turbine blade of claim 1, and further comprising:

a trailing edge cooling supply channel positioned in the trailing edge region of the airfoil;

a row of exit cooling slots or holes on the trailing edge; and

,

cooling air connecting means between the trailing edge cooling supply channel and the exit cooling slots or holes to discharge cooling air from the supply channel to the slots or holes.

10. The turbine blade of claim 1, and further comprising:

the five channels in the 5-pass serpentine flow circuit each extend between the pressure side wall and the suction side wall of the airfoil.

11. The turbine blade of claim 1, and further comprising:

the last leg of the 5-pass serpentine flow cooling circuit includes a row of film cooling holes one the pressure side or the suction side wall to discharge film cooling air just upstream of the trailing edge region of the airfoil.

12. The turbine blade of claim 6, and further comprising:

the last leg of the 5-pass serpentine flow cooling circuit includes a row of film cooling holes one the pressure side wall to discharge film cooling air just upstream of the trailing edge region of the airfoil.

13. A turbine blade for a gas turbine engine comprising:

a leading edge region and a trailing edge region extending between a mid-chord region of an airfoil;

a five-pass serpentine flow cooling circuit to cool the mid-chord region of the airfoil;

the five-pass serpentine flow cooling circuit includes a first leg located in the mid-chord region, a second leg located forward of the first leg, a third leg located forward of the second leg, a fifth leg located rearward of the first leg, and a fifth leg located rearward of the fourth leg; and

,

the five legs of the serpentine flow cooling circuit being connected in series.

14. The turbine blade of claim 13, and further comprising:

the five legs extend from a platform region of the blade to a tip region of the blade.

15. The turbine blade of claim 13, and further comprising:

a crossover channel joins the third leg to the fourth leg and forms a tip cooling channel for a tip of the blade.

16. The turbine blade of claim 13, and further comprising:

a leading edge cooling circuit located in the leading edge region;

a trailing edge cooling circuit located in the trailing edge region; and

,

the five pass serpentine flow circuit being a separate cooling circuit from the leading and trailing edge cooling circuits.

17. The turbine blade of claim 13, and further comprising:

the first leg of the five-pass serpentine circuit is positioned between the second leg and the fourth leg.