Method and apparatus to enhance laminar flow for gas turbine engine components

Patent metadata

Field	Value
Publication number	US-11466574-B2
Application number	US-202016983398-A
Country	US
Kind code	B2
Filing date	Aug 3, 2020
Priority date	May 16, 2016
Publication date	Oct 11, 2022
Grant date	Oct 11, 2022

Abstract

A gas turbine engine component has a component body configured to be positioned within a flow path of a gas turbine engine having an external pressure, and wherein the component body includes at least one internal cavity having an internal pressure. At least one inlet opening is formed in an outer surface of the component body to direct hot exhaust gas flow into the at least one internal cavity, and there is at least one outlet from the internal cavity. The internal pressure is less than an inlet external pressure at the inlet opening and the internal pressure is greater than an outlet external pressure at the outlet opening to controllably ingest hot exhaust gas via the inlet opening and expel the hot exhaust gas via the outlet opening to maintain a laminar boundary layer along the outer surface of the component body.

First claim

The invention claimed is:

1. A gas turbine engine component comprising:
a component body configured to be positioned within a flow path of a gas turbine engine having an external pressure, wherein the component body includes a leading edge, a trailing edge, and pressure and suction side walls extending from the leading edge to the trailing edge, and wherein the component body extends radially from a base, and wherein the component body includes at least one internal cavity having an internal pressure, and wherein the component body is positioned downstream of a combustor section and comprises a non-cooled component where the at least one internal cavity is free from receiving cooling flow;
a plurality of inlet openings formed in an outer surface of the pressure and suction side walls of the component body to direct hot exhaust gas flow into the at least one internal cavity, wherein the plurality of inlet openings are spaced apart from each other in a direction extending radially from the base, and wherein the leading edge is free from the plurality of inlet openings; and
at least one outlet from the at least one internal cavity formed at the trailing edge, wherein the internal pressure is less than an inlet external pressure at the plurality of inlet openings and the internal pressure is greater than an outlet external pressure at the at least one outlet to controllably ingest hot exhaust gas via the plurality of inlet openings and expel the hot exhaust gas flow via the at least one outlet to maintain a laminar boundary layer along the outer surface of the component body.

2. The gas turbine engine component according to claim 1 wherein the component body comprises a platform.

3. The gas turbine engine component according to claim 1 wherein the component body comprises an airfoil in a turbine, wherein the airfoil extends from the base to a tip.

4. The gas turbine engine component according to claim 3 wherein the at least one outlet comprises at least one opening to the outer surface that is located at the trailing edge.

5. The gas turbine engine component according to claim 3 wherein the at least one outlet comprises at least one opening to the outer surface that is located near or at the tip.

6. The gas turbine engine component according to claim 1 wherein the plurality of inlet openings are exposed to temperatures as high as 2000 degrees Celsius, and wherein the plurality of inlet openings provide a passage surface that is coated with at least one of a thermal barrier coating or environmental barrier coating.

7. The gas turbine engine component according to claim 6 wherein the plurality of inlet openings provide the passage surface that is coated with the thermal barrier coating and the environmental barrier coating to comprise a plurality of coatings.

8. The gas turbine engine component according to claim 7 wherein an outermost layer of the plurality of coatings is comprised of the thermal barrier coating.

9. The gas turbine engine component according to claim 1 wherein the plurality of inlet openings are exposed to temperatures as high as 2000 degrees Celsius, and wherein the at least one internal cavity is coated with at least one of a thermal barrier coating or environmental barrier coating.

10. The gas turbine engine component according to claim 9 wherein the at least one internal cavity and the plurality of inlet openings are coated with the thermal barrier coating and the environmental barrier coating.

11. The gas turbine engine component according to claim 10 wherein the component body is comprised of a non-metallic material.

12. The gas turbine engine component according to claim 11 wherein the non-metallic material is a ceramic matrix composite material.

13. The gas turbine engine component according to claim 1 wherein the component body comprises a transition duct.

14. The gas turbine engine component according to claim 1 wherein the plurality of inlet openings are spaced apart from each other in a first direction extending from the base to a tip, and wherein the plurality of inlet openings are spaced apart from each other in a second direction extending from the leading edge to the trailing edge.

15. A method of enhancing laminar flow for a gas turbine engine component comprising the steps of:
a) positioning a component body within a hot gas flow of a gas turbine engine having an external pressure, wherein the component body includes a leading edge, a trailing edge, and pressure and suction side walls extending from the leading edge to the trailing edge, and wherein the component body extends radially from a base, and wherein the component body includes at least one internal cavity having an internal pressure, and wherein the component body is positioned downstream of a combustor section and comprises one of an airfoil, a platform, or a transition duct in at least one of a mid-turbine frame or turbine exhaust case;
b) keeping the at least one internal cavity free from cooling flow;
c) providing a plurality of inlet openings formed in an external surface of the pressure and suction side walls of the component body to direct hot exhaust gas flow into the at least one internal cavity, wherein the plurality of inlet openings are spaced apart from each other in a direction extending radially from the base, and wherein the leading edge is free from the plurality of inlet openings, and providing at least one outlet from the internal cavity to external atmosphere, wherein the at least one outlet is at the trailing edge, and
d) maintaining the internal pressure to be less than the external pressure at the plurality of inlet openings and to be greater than the external pressure at the at least one outlet to controllably ingest a portion of the hot exhaust gas flow via the plurality of inlet openings and expel ingested hot exhaust gas flow via the at least one outlet to form a laminar boundary layer of a remaining portion of the hot gas flow along the external surface of the component body.

16. The method according to claim 15 including
forming the component body from a non-metallic material, and
coating at least one of the at least one internal cavity and the plurality of inlet openings with at least one of a thermal barrier coating or environmental barrier coating.

17. The method according to claim 15 wherein the plurality of inlet openings are exposed to temperatures as high as 2000 degrees Celsius, and including
forming the component body from a non-metallic material, and
coating a passage surface of each of the plurality of inlet openings with a plurality of coatings including at least a thermal barrier coating and an environmental barrier coating.

18. The method according to claim 17 wherein an outermost layer of the plurality of coatings is comprised of the thermal barrier coating.

19. The method according to claim 18 including coating the at least one internal cavity with a plurality of coatings including at least the thermal barrier coating and the environmental barrier coating.

20. The method according to claim 17 wherein the component body comprises a platform.

21. The method according to claim 17 wherein the component body comprises a transition duct.

22. The method according to claim 17 wherein the component body comprises an airfoil in a turbine.

23. The method according to claim 15 including spacing the plurality of inlet openings apart from each other in a first direction extending from the base to a tip, and spacing t

Assignees

• Raytheon Tech Corp

Inventors

• Xu Jinquan
• Auxier James Tilsley

Classifications

• F01D5/14

  Form or construction (selecting particular materials, measures against erosion or corrosion F01D5/28) · CPC title

• F05D2240/35

  Combustors or associated equipment · CPC title

• F05D2300/611

  Coating · CPC title

• F05D2240/80

  Platforms for stationary or moving blades · CPC title

• F01D5/18

  Hollow blades, {i.e. blades with cooling or heating channels or cavities (structure of hollow blades in general F01D5/147)}; Heating, heat-insulating or cooling means on blades · CPC title