Internally cooled gas turbine engine airfoil

Patent metadata

Field	Value
Publication number	US-10221695-B2
Application number	US-201615170110-A
Country	US
Kind code	B2
Filing date	Jun 1, 2016
Priority date	Sep 25, 2012
Publication date	Mar 5, 2019
Grant date	Mar 5, 2019

Abstract

A gas turbine engine airfoil has a hollow airfoil section extending chordwise between a leading edge and a trailing edge. The airfoil has a leading edge cooling passage and a separate serpentine passage for cooling a remaining portion of the airfoil. The serpentine passage has at least three segment serially interconnected in fluid flow communication. The leading edge cooling passage and the serpentine cooling passage have separate coolant inlets. The coolant inlet of the serpentine passage comprises a primary inlet branch connected in fluid flow communication with a first one of the segments of the serpentine passage and a secondary inlet branch connected in flow communication with a last one of the segments, thereby providing for a portion of the flow passing through the coolant inlet of the serpentine passage to be directly fed into the last segment of the serpentine passage.

First claim

What is claimed is:

1. An internally cooled airfoil for a gas turbine engine, comprising: an airfoil body extending chordwise between a leading and a trailing edge and spanwise between a root and a tip; a leading edge cooling passage extending spanwise through said airfoil body, said leading edge cooling passage having a leading edge coolant inlet defined in said root and a leading edge coolant outlet comprising film holes distributed along the leading edge; a serpentine cooling passage extending through said airfoil body, said serpentine cooling passage being separate from said leading edge cooling passage and having at least a first spanwise segment disposed adjacent to the leading edge cooling passage, a second spanwise segment connected in fluid flow communication with the first spanwise segment and generally disposed in a mid-chord region of the airfoil body, and a third spanwise segment connected in fluid flow communication with the second spanwise segment and generally located in a trailing edge region of the airfoil body, the serpentine cooling passage further comprising a serpentine coolant inlet defined in the root of the airfoil and in fluid flow communication with the first spanwise segment, and at least one serpentine coolant outlet for discharging coolant from the third spanwise segment; and a bypass opening defined in the airfoil body for directing a portion of the coolant admitted into the serpentine coolant inlet directly into the third spanwise segment, the third spanwise segment including a row of spanwise distributed V-shaped trip strips, the apex of the V-shaped trip strips pointing towards the bypass opening, the V-shaped trip strips located at a tip end of the third spanwise segment being smaller that the V-shaped trip strips at a root end of the third spanwise segment, wherein the third spanwise segment has a rear crossover wall, and wherein the serpentine coolant outlet comprises a series of spanwise distributed crossover openings defined through the rear crossover wall and in fluid flow communication with a series of trailing-edge exit slots defined along the trailing edge of the airfoil body, and wherein the third spanwise segment becomes narrower in a spanwise direction from said root to said tip.

2. The internally cooled airfoil defined in claim 1 , wherein the shape of the crossover openings located at a radially inner end of the third spanwise segment is different than the shape of the crossover openings located at a radially outer end of the third spanwise segment.

3. The internally cooled airfoil defined in claim 1 , wherein the size of the crossover openings located at a radially inner end of the third spanwise segment is different than the size of the crossover openings located at a radially outer end of the third spanwise segment.

4. The internally cooled airfoil defined in claim 1 , wherein the trailing-edge exit slots have a size and/or a density which varies as a function of their spanwise location.

5. The internally cooled airfoil defined in claim 1 , wherein the serpentine coolant inlet comprises a plurality of intake openings spaced-apart along the chordwise direction of the airfoil body, said spaced-apart intake openings being all connected in fluid flow communication with an inlet end of said first spanwise segment.

6. The internally cooled airfoil defined in claim 5 , wherein said plurality of intake openings are separated from said second and third spanwise segments by a rear internal partition wall extending in a chordwise direction at the root of the airfoil body, and wherein said bypass opening extends through said rear internal partition wall.

7. The internally cooled airfoil defined in claim 1 , wherein the third spanwise segment becomes gradually narrower in a spanwise direction away from the bypass opening.

8. An internally cooled airfoil for a gas turbine engine, the airfoil comprising: an airfoil section extending chordwise between a leading edge and a trailing edge, a leading edge cooling passage extending radially through said airfoil section for cooling the leading edge of the airfoil section; film holes distributed along the leading edge; a serpentine passage defined in said airfoil section for cooling a remaining portion of the airfoil section, the serpentine passage including at least three radially extending segments serially interconnected in fluid flow communication, wherein the coolant inlet of the serpentine passage comprises a primary inlet branch in fluid flow communication with a first one of the at least three spanwise segments of the serpentine passage and a secondary inlet branch in flow communication with both the first one and a last one of the at least three spanwise segments, thereby providing for a portion of the flow passing through the coolant inlet of the serpentine passage to be directly fed into the last segment of the serpentine passage, the last segment narrowing down in a spanwise direction towards a tip of the airfoil, the last segment being separated from an adjacent one of the at least three radially extending segments by a partition wall which diverges away from the trailing edge as it extends towards a root of the airfoil, wherein the secondary inlet branch includes a bypass opening defined in an internal partition wall of the airfoil, the bypass opening being generally aligned with a set of V-shaped trip strips distributed along the spanwise direction, each V-shaped trip strip having an apex pointing towards the bypass opening, the V-shaped trip strips at a tip end of the last segment being smaller that the V-shaped trip strips at a root end of the last segment, wherein said last segment is delimited on a rear side thereof by an internal crossover wall extending radially through the airfoil section, the crossover wall defining a plurality of radially spaced-apart crossover openings along the length thereof, and wherein the crossover openings have a size varying as a function of their spanwise location.

9. The internally cooled airfoil defined in claim 8 , wherein the cross-over wall narrows down in a spanwise direction towards the tip of the airfoil.

10. The internally cooled airfoil defined in claim 8 , wherein the leading edge cooling passage and the serpentine cooling passage have separate coolant inlets.

11. The internally cooled airfoil defined in claim 8 , wherein said internal partition wall separates said coolant inlet of said serpentine passage from the last segment of the serpentine passage.

12. The internally cooled airfoil defined in claim 8 , wherein the last segment of the serpentine passage extends radially in a trailing edge region of the airfoil section.

13. The internally cooled airfoil defined in claim 8 , wherein the coolant flowing through the internal crossover wall is discharged from the airfoil section via a series of radially spaced-apart exit slots defined in the trailing edge of the airfoil section, and wherein the trailing-edge exit slots have a size and/or a density which varies as a function of their spanwise location.

14. The internally cooled airfoil defined in claim 8 , wherein the film holes are angled to discharge coolant with an axially forward and radially outward component.

15. The internally cooled airfoil defined in claim 8 , wherein the airfoil is a turbine blade.

16. The internally cooled airfoil defined in claim 8 , wherein trip strips extend adjacent to a back side of the leading edge cooling passage.

Assignees

• Pratt & Whitney Canada

Inventors

• Papple Michael
• Plante Ghislain

Classifications

• F05D2260/2212

  by creating turbulence · CPC title

• F01D5/187 Primary

  Convection cooling · CPC title

• F05D2260/202

  by film cooling · CPC title

• F05D2250/185

  serpentine-like · CPC title

• F05D2220/32

  in gas turbines · CPC title