US-12618561-B1 - Combustion liner having cooling holes with stepped lateral sidewalls

Abstract

Apparatus and associated methods relate to geometry of cooling holes in a combustion liner for a gas turbine engine. The combustion liner includes a base-alloy substrate having a bottom surface and a top surface, a thermal barrier coat (TBC) metallic layer on the top surface of the base-alloy structure, and a TBC ceramic layer on the TBC metallic layer. The cooling holes are formed at oblique angles to the top surface of the TBC ceramic layer, thereby forming oval-characteristic exit apertures at the top surface of the TBC ceramic layer. The oval-characteristic exit apertures define a long axis and a short axis. The long axis extends between an upstream side and a downstream side and the short axis extends between opposite lateral sides, thereby defining opposite lateral sidewalls, which have stepped sidewall profiles, in the cooling holes.

Inventors

• John Ols
• Mary Gurak
• Judith Brooks
• Deanna Jindal
• Roger Coffey

Assignees

• RTX CORPORATION

Dates

Publication Date

20260505

Application Date

20250217

Claims (20)

1. 1 . A combustion liner comprising: a base-alloy substrate having a bottom surface and a top surface; a thermal barrier coat (TBC) metallic layer on the top surface of the base-alloy substrate; a TBC ceramic layer on the TBC metallic layer, thereby forming a TBC interface therebetween, the TBC ceramic layer having a top surface exposed to combustion gases during operation within a gas turbine engine; and a plurality of cooling holes, each extending from an entrance aperture at the bottom surface of the base-alloy substrate and an oval-characteristic exit aperture at the top surface of the TBC ceramic layer, each of the plurality of cooling holes formed at an oblique angle to the top surface of the TBC ceramic layer, thereby forming the oval-characteristic exit aperture at the top surface of the TBC ceramic layer, the oval-characteristic exit aperture defining a long axis and a short axis, the long axis extending between an upstream side and a downstream side and the short axis extending between opposite lateral sides, thereby defining upstream and downstream sidewalls and opposite lateral sidewalls in each of the base-alloy substrate, the TBC metallic layer, and the TBC ceramic layer, wherein each of the cooling holes has a stepped sidewall profile recessing the TBC ceramic layer from TBC metallic layer at each of the upstream and downstream sidewalls and the opposite lateral sidewalls, wherein, a first distance between the opposite lateral sidewall of the TBC ceramic layer is greater than a second distance between the opposite lateral sidewalls of the TBC metallic layer.
2. 2 . The combustion liner of claim 1 , wherein the stepped sidewall profile includes an exposed top surface of the TBC metallic layer, thereby forming an exposed metallic ledge as viewed from the exit aperture.
3. 3 . The combustion liner of claim 2 , wherein the first distance between the opposite lateral sidewalls of the TBC ceramic layer is greater than a third distance between the opposite lateral sidewalls of the base-alloy substrate.
4. 4 . The combustion liner of claim 3 , wherein the metallic ledge is an exposed surface of the TBC metallic layer.
5. 5 . The combustion liner of claim 2 , wherein the metallic ledge has a lateral ledge width, as measured in the direction of the short axis, that is between 5% and 40% of the first distance between the opposite lateral sidewalls of the TBC ceramic layer at the TBC interface.
6. 6 . The combustion liner of claim 5 , wherein the lateral ledge width, as measured in the direction of the short axis, that is between 10% and 30% of the first distance between the opposite lateral sidewalls of the TBC ceramic layer at the TBC interface between the TBC metallic layer and the TBC ceramic layer.
7. 7 . The combustion liner of claim 1 , wherein a portion of the stepped sidewall profile corresponding to the TBC ceramic layer is conic shaped or bell shaped, with the distance between the opposite lateral sidewalls of the TBC ceramic layer monotonically increasing as measured from the TBC interface to the top surface of the TBC ceramic layer.
8. 8 . The combustion liner of claim 7 , wherein a ratio of a first width between the opposite lateral sidewalls of the TBC ceramic layer, as measured at the TBC interface, to a second width between the opposite lateral sidewalls of the base-alloy substrate, is greater than 1.1.
9. 9 . The combustion liner of claim 8 , wherein a ratio of the first width between the opposite lateral sidewalls of the TBC ceramic layer, as measured at the TBC interface, to the second width between the opposite lateral sidewalls of the base-alloy substrate, is greater than 1.2.
10. 10 . A method for creating a combustion liner, the method comprising: providing a base-alloy substrate having a bottom surface and a top surface; depositing a thermal barrier coat (TBC) metallic layer on the top surface of the base-alloy substrate; depositing a TBC ceramic layer on the TBC metallic layer, thereby forming a TBC interface therebetween, the TBC ceramic layer having a top surface exposed to combustion gases during operation within a gas turbine engine; and forming a plurality of cooling holes, each extending from an entrance aperture at the bottom surface of the base-alloy substrate and an oval-characteristic exit aperture at the top surface of the TBC ceramic layer, each of the plurality of cooling holes formed at an oblique angles to the top surface of the TBC ceramic layer, thereby forming the oval-characteristic exit aperture at the top surface of the TBC ceramic layer, the oval-characteristic exit aperture defining a long axis and a short axis, the long axis extending between an upstream side and a downstream side and the short axis extending between opposite lateral sides, thereby defining upstream and downstream sidewalls and opposite lateral sidewalls in each of the base-alloy substrate, the TBC metallic layer, and the TBC ceramic layer, wherein each of the cooling holes has a stepped sidewall profile recessing the TBC ceramic layer from TBC metallic layer at each of the upstream and downstream sidewalls and the opposite lateral sidewalls, wherein a first distance between the opposite lateral sidewalls of the TBC ceramic layer is greater than a second distance between the opposite lateral sidewalls of the TBC metallic layer.
11. 11 . The method of claim 10 , wherein drilling the plurality of cooling holes is performed by a waterjet hole drill.
12. 12 . The method of claim 11 , wherein the waterjet hole drill is configured to erode more the TBC ceramic layer than the base-alloy substrate.
13. 13 . The method of claim 11 , wherein the waterjet hole drill is configured to erode more the TBC ceramic layer than the TBC metallic layer.
14. 14 . The method of claim 10 , wherein the stepped sidewall profile includes an exposed top surface of the TBC metallic layer, thereby forming an exposed metallic ledge as viewed from the exit aperture.
15. 15 . The method of claim 14 , wherein the first distance between the opposite lateral sidewalls of the TBC ceramic layer is greater than a third distance between the opposite lateral sidewalls of the base-alloy substrate.
16. 16 . The method of claim 15 , wherein the metallic ledge is an exposed surface of the TBC metallic layer.
17. 17 . The method of claim 14 , wherein the metallic ledge has a width, as measured in the direction of the short axis, that is between 5% and 40% of the first distance between the opposite lateral sidewalls of the TBC ceramic layer at the TBC interface.
18. 18 . The combustion liner of claim 14 , wherein the metallic ledge has a width, as measured in the direction of the short axis, that is between 5% and 20% of the second distance between the opposite lateral sidewalls of the base-alloy substrate.
19. 19 . The combustion liner of claim 10 , wherein a portion of the stepped sidewall profile corresponding to the TBC ceramic layer is conic shaped or bell shaped, with the distance between the opposite lateral sidewalls of the TBC ceramic layer monotonically increasing as measured from the TBC interface to the top surface of the TBC ceramic layer.
20. 20 . The method of claim 19 , wherein a ratio of a first width between the opposite lateral sidewalls of the TBC ceramic layer, as measured at the TBC interface, to a second width between the opposite lateral sidewalls of the base-alloy substrate, is greater than 1.1.

Description

BACKGROUND Gas turbine engines can operate at very high temperatures for long periods of time. Various components of gas turbine engines can be exposed to very hot gases, such as the gases produced in the combustion chamber of gas turbine engines. These products of combustion provide high thermal exposure to various components, such as, for example, combustion liners, turbine blades, and nozzle guide vanes. Insufficient cooling of these components can result in local thermal cracks and can reduce the strength of the components' materials. Various cooling technologies can be used to protect these components, so as to extend the life of these components. To protect surfaces of these components from exposure to temperatures higher than the component's safe thermal-exposure specification, a secondary flow can be introduced by means of holes over surfaces resulting in formation of a film of cooling air flowing thereover. This film of cooling air operates as a protection layer between high temperature gases and the components' surfaces. Such a cooling technique is called effusion cooling or film cooling. SUMMARY Some embodiments are related to a combustion liner with a plurality of cooling holes. The combustion liner includes a base-alloy substrate having a bottom surface and a top surface, a thermal barrier coat (TBC) metallic layer on the top surface of the base-alloy structure, and a TBC ceramic layer on the TBC metallic layer. The TBC ceramic layer has a top surface configured to be exposed to combustion gases during operation within a gas turbine engine. The plurality of cooling holes is formed, each through the combustion liner extending from an entrance aperture at the bottom surface of the base-alloy structure and an exit aperture at the top surface of the TBC ceramic layer. Each of the plurality of cooling holes is formed at an oblique angle to the top surface of the TBC ceramic layer, thereby forming an oval-characteristic exit aperture at the top surface of the TBC ceramic layer. Each of the oval-characteristic exit apertures defines a long axis and a short axis. The long axis extends between an upstream side and a downstream side and the short axis extends between opposite lateral sides, thereby defining opposite lateral sidewalls in each of the base-alloy substrate, the TBC metallic layer, and the TBC ceramic layer. Each of the cooling holes has a stepped sidewall profile at each of the opposite lateral sidewalls. A first distance between the opposite lateral sidewall of the TBC ceramic layer is greater than a second distance between the opposite lateral sidewalls of the TBC metallic layer. Some embodiments relate to a method for creating a combustion liner. In the method, a base-alloy substrate having a bottom surface and a top surface is provided. A thermal barrier coat (TBC) metallic layer is deposited on the top surface of the base-alloy structure. A TBC ceramic layer is deposited on the TBC metallic layer. The TBC ceramic layer has a top surface configured to be exposed to combustion gases during operation within a gas turbine engine. A plurality of cooling holes is formed through the combustion liner, each extending from an entrance aperture at the bottom surface of the base-alloy structure and an exit aperture at the top surface of the TBC ceramic layer. Each of the plurality of cooling holes is formed at an oblique angles to the top surface of the TBC ceramic layer, thereby forming an oval-characteristic exit aperture at the top surface of the TBC ceramic layer. Each of the oval-characteristic exit apertures defines a long axis and a short axis. The long axis extends between an upstream side and a downstream side, and the short axis extends between opposite lateral sides, thereby defining opposite lateral sidewalls in each of the base-alloy substrate, the TBC metallic layer, and the TBC ceramic layer. Each of the cooling holes has a stepped sidewall profile at each of the opposite lateral sidewalls. A first distance between the opposite lateral sidewalls of the TBC ceramic layer is greater than a second distance between the opposite lateral sidewalls of the TBC metallic layer. BRIEF DESCRIPTION OF THE DRAWINGS The material described herein is illustrated by way of example and not by way of limitation in the accompanying figures. For simplicity and clarity of illustration, elements illustrated in the figures are not necessarily drawn to scale. For example, the dimensions of some elements may be exaggerated relative to other elements for clarity. Further, where considered appropriate, reference labels have been repeated among the figures to indicate corresponding or analogous elements. In the figures: FIGS. 1A-1B are a longitudinal cross-sectional view of a region of a combustion liner and a plan view of the region of the combustion liner, respectively. FIG. 2 is a plan view of a single cooling hole, depicting various layers of the combustion liner. FIGS. 3A and 3B are lateral cross-sectional v