Flow directing cover for engine component

Patent metadata

Field	Value
Publication number	US-9845694-B2
Application number	US-201514693359-A
Country	US
Kind code	B2
Filing date	Apr 22, 2015
Priority date	Apr 22, 2015
Publication date	Dec 19, 2017
Grant date	Dec 19, 2017

Abstract

An assembly for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end. The assembly further includes a cover having at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway.

First claim

The invention claimed is:

1. An assembly for a gas turbine engine, comprising:
an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end; and
a cover including at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway;
wherein:
the cover includes a first turning cavity and a second turning cavity;
the airfoil includes a third passageway having an inlet at the radial end;
the first turning cavity is configured to direct a first portion of the fluid expelled from the outlet of the first passageway into the inlet of the second passageway; and
the second turning cavity is configured to direct a second portion of the fluid expelled from the outlet of the first passageway into the inlet of the third passageway.

2. The assembly as recited in claim 1 , wherein the cover includes a flow divider between the first turning cavity and the second turning cavity.

3. The assembly as recited in claim 2 , wherein, when viewed from an interior of the cover, the flow divider is substantially convex and the first and second turning cavities are substantially concave.

4. The assembly as recited in claim 3 , wherein the first and second turning cavities are substantially semi-circular in cross-section.

5. The assembly as recited in claim 1 , wherein the first passageway is inward of the second and third passageways relative to an exterior wall of an airfoil.

6. The assembly as recited in claim 5 , wherein the second passageway extends along one of a pressure and a suction side wall of the airfoil, and wherein the third passageway extends along the other of the pressure and the suction side wall of the airfoil.

7. A gas turbine engine, comprising:
a source of cooling fluid;
an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end, the first passageway fluidly coupled to the source of cooling fluid; and
a cover including at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway;
wherein:
the cover includes a first turning cavity and a second turning cavity;
the airfoil includes a third passageway having an inlet at the radial end;
the first turning cavity is configured to direct a first portion of the fluid expelled from the outlet of the first passageway into the inlet of the second passageway; and
the second turning cavity is configured to direct a second portion of the fluid expelled from the outlet of the first passageway into the inlet of the third passageway.

8. The engine as recited in claim 7 , wherein the airfoil is a stator vane within a turbine section of the engine.

9. The engine as recited in claim 7 , wherein the source of cooling fluid is a compressor of the engine.

10. The engine as recited in claim 7 , wherein the cover includes a flow divider between the first turning cavity and the second turning cavity.

11. The engine as recited in claim 10 , wherein, when viewed from an interior of the cover, the flow divider is substantially convex and the first and second turning cavities are substantially concave.

12. The engine as recited in claim 7 , wherein the first and second turning cavities are radially spaced-apart from a core airflow path of the gas turbine engine.

Assignees

• United Technologies Corp

Inventors

• Hagan Benjamin F
• Waite Ryan Alan
• Dube Bryan P
• Thornton Lane

Classifications

• F01D5/186

  Film cooling (F01D5/187 takes precedence) · CPC title

• F05D2240/125

  related to the tip of a stator vane · CPC title

• F01D9/041

  using blades (F01D5/148 takes precedence) · CPC title

• F01D5/187

  Convection cooling · CPC title

• F05D2250/75

  given by its similarity to a letter, e.g. T-shaped · CPC title