Tip leakage flow directionality control

Patent metadata

Field	Value
Publication number	US-9957817-B2
Application number	US-201314143072-A
Country	US
Kind code	B2
Filing date	Dec 30, 2013
Priority date	Jul 3, 2012
Publication date	May 1, 2018
Grant date	May 1, 2018

Abstract

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil having a tip wall that joins outer spanwise ends of a suction sidewall and a pressure sidewall, a tip leakage control channel recessed into an outer surface of the tip wall, a tip leakage control vane integrally formed on the tip wall adjacent to the tip leakage control channel and an air seal radially outward of the tip wall and positioned to minimize leakage at the tip wall.

First claim

What is claimed is:

1. A gas turbine engine, comprising:
an airfoil having a tip wall that joins outer spanwise ends of a suction sidewall and a pressure sidewall and including a tip rib extending chordwise along a pressure side of said tip wall, wherein said tip wall includes a tip shelf disposed at a junction between said pressure sidewall and said tip rib;
a plurality curved tip leakage control channels recessed into an outer surface of said tip wall, wherein each of said plurality of curved tip leakage control channels is adjacent to
one of a plurality of curved tip leakage control vanes integrally formed on said tip wall contiguous with a suction side surface of said tip rib, said plurality of curved tip leakage channels extending toward said trailing edge while extending toward said suction sidewall;
a first cooling aperture that feeds airflow to said tip leakage control channel;
a second cooling aperture that feeds airflow to said tip shelf; and
an air seal radially outward of said tip wall and positioned to minimize leakage at said tip wall.

2. The gas turbine engine as recited in claim 1 , wherein said air seal is a grooved blade outer air seal (BOAS).

3. The gas turbine engine as recited in claim 2 , wherein said grooved BOAS includes a plurality of projections that extend inwardly from a radially inner surface of said grooved BOAS.

4. The gas turbine engine as recited in claim 3 , comprising a plurality of grooves that extend between adjacent ones of said plurality of projections to define a circuitous flow path at said radially inner surface.

5. The gas turbine engine as recited in claim 1 , wherein said tip leakage control channel includes an inlet and an outlet recessed into said outer surface of said tip wall, said inlet beginning proximate a junction of said pressure sidewall and said tip wall, and said outlet terminating at a recessed portion of a junction of said tip wall and said suction sidewall.

6. The gas turbine engine as recited in claim 1 , wherein said tip leakage control vane projects radially outward in a spanwise direction from a tip floor of said tip wall, a leading portion of said tip leakage control vane beginning proximate a junction of said pressure sidewall and said tip wall, and a trailing portion of said tip leakage control vane terminating proximate a junction of said suction sidewall and said tip wall.

7. The gas turbine engine as recited in claim 1 , comprising a winglet that extends from said tip leakage control vane at said suction sidewall.

8. The gas turbine engine as recited in claim 1 , wherein at least one of said first cooling aperture and said second cooling aperature includes a curvature.

9. The gas turbine engine as recited in claim 1 , wherein said airfoil is a turbine airfoil.

10. The gas turbine engine as recited in claim 1 , wherein at least one of said plurality of curved tip leakage control channels includes an inlet having a first chordwise width and an outlet having a second chordwise width that is different from said first chordwise width.

11. The gas turbine engine as recited in claim 1 , wherein said tip rib is disposed circumferentially between said pressure sidewall and an inlet of at least one of said plurality of curved tip leakage control channels.

12. The gas turbine engine as recited in claim 1 , wherein at least one of said plurality of curved tip leakage control channels includes an inlet having a channel entrance interior angle relative to said pressure sidewall and an outlet having a channel exit interior angle relative to said suction sidewall, and said channel entrance interior angle is different from said channel exit interior angle.

13. The gas turbine engine as recited in claim 12 , wherein said channel entrance interior angle is between about 80° and about 95°.

14. The gas turbine engine as recited in claim 1 , wherein at least one of said plurality of curved tip leakage control vanes includes a leading portion forming a leading interior angle relative to said pressure sidewall and a trailing portion forming a trailing interior angle relative to said suction sidewall, and said trailing interior angle is different from said leading interior angle.

15. A gas turbine engine, comprising:
an airfoil having a tip wall including a plurality of curved tip leakage control channels recessed into an outer surface of said tip wall, said plurality of curved tip leakage control channels extending toward a trailing edge of said airfoil while extending toward a suction sidewall of said airfoil a tip leakage control vane formed with the outer surface of said tip wall and adjacent the plurality of curved tip leakage control channels; a winglet formed at a junction between said suction sidewall and said tip leakage control vane and localized at said tip leakage control vane; and
a grooved blade outer air seal (BOAS) positioned radially outward of said tip wall and configured to minimize clearances at said tip wall, wherein said grooved BOAS includes a plurality of projections that protrude from a radially inner surface of said grooved BOAS and extend away from said radially inner surface in a direction toward said tip wall.

16. The gas turbine engine as recited in claim 15 , comprising a plurality of grooves that extend between adjacent ones of said plurality of projections to define a circuitous flow path at said radially inner surface.

17. The gas turbine engine as recited in claim 15 , wherein said plurality of projections extend inwardly at an angle relative to the radially inner surface.

18. The gas turbine engine as recited in claim 17 , wherein a first projection of said plurality of projections extends at a first angle and a second projection of said plurality of projections extends at a second, different angle.

Assignees

• Zelesky Mark F
• Aggarwala Andrew S
• United Technologies Corp

Inventors

• Zelesky Mark F
• Aggarwala Andrew S

Classifications

• Y02T50/673

  Cross-Sectional Technologies · mapped topic

• F05D2240/55

  Seals · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F01D11/08

  for sealing space between rotor blade tips and stator (specially-shaped blade tips therefor F01D5/20) · CPC title

• F05D2250/294

  grooved · CPC title