Turbine airfoil cooling passageway
An internally cooled gas turbine engine turbine vane has an outboard shroud and an airfoil extending from an outboard end at the shroud to an inboard end. A cooling passageway has an inlet in the shroud, a first turn at least partially within the airfoil, a first leg extending from the inlet inboard through the airfoil to the first turn, and a second leg extending from the first turn. A dividing wall is in the passageway and has an upstream end in an outboard half of a span of the airfoil and has a plurality of vents. The vane may be formed as a reengineering of a baseline configuration lacking the dividing wall.
1. An internally-cooled gas turbine engine turbine vane comprising:
an outboard shroud;
an airfoil extending from an outboard end at the shroud to an inboard end;
a cooling passageway having:
an inlet in the shroud;
a first turn at least partially within the airfoil;
a first leg extending from the inlet inboard through the airfoil to the first turn; and
a second leg extending from the first turn and separated from the first leg by a wall; and
a dividing wall in the passageway, dividing at least one of the first and second legs, and having:
an upstream end in an outboard half of a span of the airfoil; and
a plurality of vents.
2. The vane of claim 1 wherein:
there are no additional features extending between airfoil pressure and suction side walls along the first leg.
3. The vane of claim 2 wherein:
the dividing wall has a length within the first leg of at least half the span of the airfoil.
4. The vane of claim 1 wherein:
the dividing wall essentially locally divides the first leg into first and second flowpath portions, each having a cross-sectional area at least 35% of a combined cross-sectional area.
5. The vane of claim 1 wherein:
the dividing wall extends to a second end outboard of the airfoil inboard end and not downstream of a middle of the first turn.
6. The vane of claim 1 wherein:
the vane has a platform at the inboard end of the airfoil; and
the first turn is partially within the platform.
7. The vane of claim 1 wherein:
the first turn is in excess of 90°.
8. The vane of claim 1 wherein:
the cooling passageway extends to a trailing edge discharge slot.
9. The vane of claim 1 wherein the wall and the dividing wall are in a unitary casting.
10. An internally-cooled turbomachine element comprising:
an airfoil extending between inboard and outboard ends; and
internal surface portions defining a cooling passageway at least partially within the airfoil, wherein:
the cooling passageway has a first turn from an upstream first leg to a downstream second leg, a wall separating the upstream leg from the downstream leg;
a dividing wall bifurcates a section of the cooling passageway into first and second portions and extends within the passageway along a length from a wall first end in the first leg to a wall second end, the wall first end being in an upstream half of a portion of the first leg within the airfoil, there being no additional features extending between airfoil pressure and suction side walls along the first leg; and
the dividing wall has a plurality of apertures.
11. The element of claim 10 wherein:
the first and second portions each provide 35–65% of a cross-sectional area of the cooling passageway along said length of the dividing wall.
12. The element of claim 10 wherein:
the dividing wall second end is proximate an end of the first leg at the first turn.
13. The element of claim 10 wherein:
the passageway has a second turn from the second leg to a third leg;
the wall extends along a majority of an airfoil span.
14. The element of claim 10 wherein:
the passageway has a second turn from the second leg to a third leg;
the third leg is along a trailing edge discharge slot.
15. The element of claim 10 being a vane and having:
an inboard platform; and
an outboard shroud.
16. The element of claim 10 wherein:
the dividing wall first end is located between 10% and 30% of a spanwise distance from the airfoil outboard end to the airfoil inboard end.
17. The element of claim 10 wherein the element consists essentially of a unitary casting.
18. A method for reengineering a configuration for an internally-cooled turbomachine element from a baseline configuration to a reengineered configuration wherein the baseline configuration has an internal passageway through an airfoil and having first and second generally spanwise legs and a first turn therebetween, the method comprising:
adding a wall to bifurcate the passageway into first and second portions, the wall extending within the passageway along a length from a wall first end to a wall second end; and
otherwise essentially maintaining a basic shape of the first cooling passageway.
19. The method of claim 18 wherein:
the first turn is around an end of a second wall.
20. The method of claim 18 wherein:
the wall has a series of apertures.
21. The method of claim 18 wherein:
the wall extends within the first leg for at least 50% of a length of the first leg within the airfoil.
22. The method of claim 18 wherein;
no additional features are added along the first leg to span between pressure and suction side walls.
23. The method of claim 18 wherein:
the wall extends within the first leg.
24. The method of claim 18 wherein the adding comprises adding the wall to a casting.