EP4793546A1

EP4793546A1 - Combustion liner having high-density of cooling holes

Info

Publication number: EP4793546A1
Application number: EP26159149.9A
Authority: EP
Inventors: John OLS; Mary Gurak; Judith Brooks; Deanna Jindal; Roger Coffey
Original assignee: RTX Corp
Current assignee: RTX Corp
Priority date: 2025-02-17
Filing date: 2026-02-17
Publication date: 2026-08-19

Abstract

Apparatus and associated methods relate to a combustion liner (10) with a plurality of cooling holes (18). The combustion liner (10) includes a base-alloy substrate (12) having a bottom surface (22) and a top surface (26), a thermal barrier coat (TBC) metallic layer (14) on the top surface (26) of the base-alloy structure (12), and a TBC ceramic layer (16) on the TBC metallic layer (14). The TBC ceramic layer (16) has a top surface (26) configured to be exposed to combustion gases during operation within a gas turbine engine. A ratio of a combined hole volume of the plurality of cooling holes (18) to a combustion-liner volume of the region of the combustion liner (10) to which the plurality of cooling holes (18) belong is greater than 20 percent.

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

In an aspect of the present invention, a combustion liner with a plurality of cooling holes is provided. 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. A ratio of a combined hole volume of the plurality of cooling holes to a combustion-liner volume of the region of the combustion liner to which the plurality of cooling holes belong is greater than 20 percent.
In another aspect of the present invention, a method for creating a combustion liner is provided. 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. A ratio of a combined hole volume of the plurality of cooling holes to a combustion-liner volume of the region of the combustion liner to which the plurality of cooling holes belong is greater than 20 percent.

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 plan view of a region of a combustion liner and a longitudinal cross-sectional 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.
FIG. 3 are a lateral cross-sectional views of a cooling hole of the combustion liner.
FIG. 4 is a longitudinal cross-sectional view of a cooling hole of the combustion liner.
FIG. 5 is a flowchart for a method of producing holes in a combustion liner.
FIGS. 6A and 6B are a cross-sectional view and a plan view of a region of a combustion liner having a high-density of cooling holes formed therein.

DETAILED DESCRIPTION

Apparatus and associated methods relate 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. A ratio of a combined hole volume of the plurality of cooling holes to a combustion-liner volume of the region of the combustion liner to which the plurality of cooling holes belong is greater than 20 percent.
FIGS. 1A-1B are a longitudinal cross-sectional view of the region of the combustion liner and a plan view of a region of a combustion liner, respectively. In FIG. 1A, a portion of combustion liner 10 includes base-alloy substrate 12, thermal barrier coat (TBC) metallic layer 14, and TBC ceramic layer 16. Although the entire combustion liner is not depicted in FIGS. 1A and 1B, base-alloy substrate 12, is typically formed into a geometry appropriate for the combustion chamber of a gas turbine engine. TBC metallic layer 14 is formed upon on side of base-alloy substrate 12 on an interior facing side of the combustion liner. TBC ceramic layer 16 is then formed upon TBC metallic layer 14. Thus, TBC ceramic layer 16 is configured to be exposed to combustion gases within the combustion chamber of the gas turbine engine. Cooling holes 18 are formed through combustion liner 10, extending from entrance aperture 20 at bottom surface 22 of base-alloy structure 12 and exit aperture 24 at top surface 26 of TBC ceramic layer 16. Each of cooling holes 18 is formed at an oblique angle θ with respect to top surface 26 of TBC ceramic layer 16. The central axes A of cooling holes 18 are typically greater than 45° from a normal vector N of top surface 26 of TBC ceramic layer 16. Such obliquely angled cooling holes 18 facilitate flow of cooling air to be directed as a film layer immediately adjacent to top surface 26 of TBC ceramic layer 16. Such a film layer of cooling air creates a thermal buffer between the hot gases within the combustion chamber and TBC ceramic layer 16.
The portion of combustion liner 10 depicted in cross-sectional view in FIG. 1A is also depicted in plan view in FIG. 1B. Annotated in FIG. 1B is cross-sectional line D-D, through which the cross-section depicted in FIG. 1A is depicted. Cooling holes 18 form oval-characteristic exit apertures 24 at top surface 26 of TBC ceramic layer 16. Such oval-characteristic exit apertures 24 are a result of the oblique angle θ of cooling holes 18 with respect to central axes A. FIG. 2 is a plan view of a single cooling hole, depicting various layers of the combustion liner. In FIG. 2, base-alloy substrate 12, TBC metallic layer 14, and TBC ceramic layer 16 can be seen through exit aperture 24 of cooling hole 18. The oval-characteristic exit aperture 24 defines a long axis L and a short axis S. The long axis L extends between an upstream side 30 and a downstream side 32 of the exit aperture 24, thereby defining upstream and downstream sidewalls in each of cooling holes 18. The short axis S extends between opposite lateral sides, thereby defining opposite lateral sidewalls in each of cooling holes 18. As will be described in more detail below, these sidewalls (the upstream sidewall, the downstream sidewall and the opposite lateral sidewalls ) have been created with profiles that facilitate formation of the film layer of cooling air.
FIGS. 3A and 34B are lateral cross-sectional views of a cooling hole of the combustion liner. In FIG. 3A, cooling hole 18 is cross-sectioned vertically, with respect to top surface 26 of TBC ceramic layer 16. In FIG. 3B, cooling hole 18 is cross-sectioned obliquely, along axis A of cooling hole 18. Because of the oblique angle of the cross section, layers appear thicker in FIG. 3B than those same layers appear in FIG. 3A. In FIGS. 3A and 3B, cooling hole 18 extends from bottom surface 22 of base-alloy structure 12 and to top surface 26 of TBC ceramic layer 16, thereby traversing through each of base-alloy substrate 12, TBC metallic layer 14, and TBC ceramic layer 16. Although TBC metallic layer 14 is thin, as compared with thicknesses of base-alloy substrate 12 and TBC ceramic layer 16, TBC metallic layer 14 can be readily seen from above exit aperture 24, as TBC ceramic layer 16 is recessed away from central axis A with respect to TBC metallic layer 14, thereby exposing a metallic ledge of a top surface of TBC metallic layer 14. A distance (i.e., a lateral ledge width), as measured in the direction of the short axis S, of the exposed surface of the TBC metallic layer 14 (i.e., a distance between each of opposite lateral sidewalls 28 of the TBC metallic layer 14 and the corresponding lateral sidewall 28 of TBC ceramic layer 16 proximate TBC interface 30) can be between 5% and 40%, or between 10% and 30%, of a distance between the opposite lateral sidewalls 28 of TBC metallic layer 14.
This recessing of TBC ceramic layer 16 away from central axis A with respect to TBC metallic layer 14 results in a stepped sidewall profile for each of opposite lateral sidewalls 28 of TBC ceramic layer 16 at TBC interface 30. Cooling hole 18 is substantially cylindrical through the combined layers of base-alloy substrate 12 and TBC metallic layer 14. Cooling hole 18 has a stepped expansion at TBC interface 30 between TBC metallic layer 14 and TBC ceramic layer 16. Cooling hole 18 then monotonically increases in lateral dimension as cooling hole 18 traverses from TBC interface 30 to top surface 26 of TBC ceramic layer 16. In the depicted embodiment, the opposite lateral sidewall profile of TBC ceramic layer 16 is bell shaped (i.e., opposite lateral sidewalls 28 are inwardly concave or outwardly convex). In other embodiments, the opposite lateral sidewall profile of TBC ceramic layer 16 can be made to be conic shaped (i.e., with straight angled opposite lateral sidewalls 28) or horn shaped (i.e., opposite lateral sidewalls 28 are inwardly convex or outwardly concave). In all these embodiments, the distance between opposite lateral sidewalls 28 of TBC ceramic layer 16 is greater than the distance between opposite lateral sidewalls 28 of TBC metallic layer 14 and of base-alloy layer 12. A ratio of the distance between the opposite lateral sidewalls 28 of TBC ceramic layer 16, as measured proximate TBC interface 30, to the distance between opposite lateral sidewalls 28 of TBC metallic layer 14, can be greater than 1.1, greater than 1.2 or even greater than 1.3. Such a stepped opposite lateral sidewall profile facilitates the production of a film layer of cooling are above top surface 26 of TBC ceramic layer 16.
FIG. 4 is a longitudinal cross-sectional view of a cooling hole of the combustion liner. In FIG. 4, cooling hole 18 extends from bottom surface 22 of base-alloy structure 12 and to top surface 26 of TBC ceramic layer 16, thereby traversing through each of base-alloy substrate 12, TBC metallic layer 14, and TBC ceramic layer 16. Although TBC metallic layer 14 is thin, as compared with thicknesses of base-alloy substrate 12 and TBC ceramic layer 16, TBC metallic layer 14 can be readily seen from above exit aperture 24, as TBC ceramic layer 16 is recessed away from central axis A with respect to TBC metallic layer 14, thereby exposing metallic ledge of a top surface of TBC metallic layer 14. A distance (i.e., an upstream ledge width), as measured in the direction of the long axis L, of the exposed surface of the TBC metallic layer 14 (i.e., a distance between upstream sidewall 32 of the TBC metallic layer 14 and the corresponding upstream sidewall 32 of TBC ceramic layer 16 proximate TBC interface 30) can be between 5% and 40%, or between 10% and 30%, of a distance between upstream sidewall 32 and downstream sidewall 34 of TBC metallic layer 14.
This recessing of TBC ceramic layer 16 away from central axis A with respect to TBC metallic layer 14 results in a stepped sidewall profile for each of upstream sidewall 32 and downstream sidewall 34, although it is much more pronounced in upstream sidewall 32. Again, from this cross-sectional perspective, cooling hole 18 is substantially cylindrical through the combined layers of base-alloy substrate 12 and TBC metallic layer 14. Cooling hole 18 has a stepped expansion at TBC interface 30 between TBC metallic layer 14 and TBC ceramic layer 16. Cooling hole 18 then monotonically increases in longitudinal dimension as cooling hole 18 traverses from TBC interface 30 to top surface 26 of TBC ceramic layer 16. In the depicted embodiment, the upstream sidewall profile of TBC ceramic layer 16 is a retrograde profile (i.e., upstream sidewall 32 is angled opposite the direction of cooling hole 18). Although depicted as unsmooth, in other embodiments, the upstream sidewall profile of TBC ceramic layer 16 can be made to be smooth (e.g., bell shaped, conic shaped, or horn shaped). A ratio of the distance between upstream sidewall 32 and downstream sidewall 34 of TBC ceramic layer 16, as measured the direction of the long axis L and proximate TBC interface 30, to the distance between upstream sidewall 32 and downstream sidewall 34 of TBC metallic layer 14, can be greater than 1.03, greater than 1.1, or even greater than 1.2. Such stepped upstream and downstream sidewall profiles facilitate the production of a film layer of cooling are above top surface 26 of TBC ceramic layer 16.
FIG. 5 is a flowchart of a method for creating a combustion liner. In FIG. 6, method 40 begins at step 42, where base-alloy substrate 12 having bottom surface 22 and a top surface is provided. Method 40 then advances to step 44, where TBC metallic layer 14 is deposited on top surface of base-alloy structure 12. Method 40 then advances to step 46, where TBC ceramic layer 16 is deposited on TBC metallic layer 14, thereby creating a combustion liner. TBC ceramic layer 16 has top surface 26, which is configured to be exposed to combustion gases during operation within a gas turbine engine. Method 40 then advances to step 48, where combustion liner 10, so produced, is positioned in relation to a water jet drilling system to form an oblique-angled cooling hole. Method 40 then advances to step 50, where the cooling hole is formed, via the water jet drilling system, through the combustion liner, extending from an entrance aperture at bottom surface 22 of base-alloy structure 12 and exit aperture 24 at top surface 26 of TBC ceramic layer 16. The water jet drilling system is configured to create the cooling hole with a stepped sidewall profile at each of opposite lateral sidewalls 28, such that a first distance between opposite lateral sidewalls 28 of TBC ceramic layer 16 is greater than a second distance between opposite lateral sidewalls 28 of base-alloy structure 12. Method 40 then advances to step 52, where combustion liner 10 is repositioned relative to the water jet drilling system for forming another cooling hole. In some embodiments, either combustion liner 10 or the water jet drilling system is repositioned or both. Method 40 then returns to step 50, where the next cooling hole is formed. The water jet drilling system can be configured to produce the desired sidewall profiles that facilitate production of a film layer of cooling air. The water jet drilling system can produce such profiled cooling holes without causing damage to the material remaining in base-alloy substrate 12, thermal barrier coat (TBC) metallic layer 14, and TBC ceramic layer 16.
FIGS. 6A and 6B are a cross-sectional view and a plan view of a region of a combustion liner having a high-density of cooling holes formed therein. In FIGS. 6A and 6B, combustion liner 10 has cooling holes 18 arranged in a close-pack configuration, in which surrounded by six nearest-neighbor cooling holes 18 about each cooling hole 18. Nearest-neighbor cooling holes 18 are radially distributed about each cooling hole 18. By using a water-jet drilling system to form the holes, little, if any, damage is imparted to the base-alloy substrate 12, thermal barrier coat (TBC) metallic layer 14, and TBC ceramic layer 16. As such, cooling holes 18 can be drilled in a higher hole density than can be obtained in traditional combustion liners. Such density can be measured in various manners. For example, a distance between adjacent nearest neighbor cooling holes, a ratio of the hole diameter to the distance between nearest neighbors, and/or a ratio of volume of the hole to a total volume of the region of combustion liner can be used as metrics of hole density.
To measure a ratio of a combined hole volume of the plurality of cooling holes to the volume of the region of the combustion liner to which the plurality of cooling holes belong, an area is determined, and the volume of holes therein is determined. For example, the hole volume V_HOLE is given by: $V_{HOLE} = {lπr}^{2},$ where r is the radius of the hole and I is the distance of the hole through the combustion liner. For such a close-pack configuration, the volume V_LINER of the combustion liner associated with only one hole is given by: $V_{LINER} l \frac{\sqrt{3}}{2} s^{2},$ where s is the center-to-center spacing, as measured orthogonal to the axes A of the cooling holes, between adjacent cooling holes. A ratio V_HOLE /V_LINER of these volumes - the volume of a single cooling hole to a volume of the associated cooling liner (which includes the cooling hole and its volume is given by: $\frac{2 π}{\sqrt{3}} \cdot {(\frac{r}{s})}^{2} .$
For a cooling hole with a diameter of 0.011 inches (0.279 mm) , and a separation distance between adjacent colling holes of 0.05 inches (1.27 mm), the ratio V_HOLE /V_LINER of the volume of a cooling hole volume to the volume of the associated region of the combustion liner is given by: $\frac{V_{HOLE}}{V_{LINER}} \approx 0.176 .$
Such a ratio can be increased using water-jet drilling techniques. For example, ratios V_HOLE /V_LINER of greater than 0.20, 025, and 0.30 can be obtained, while maintaining structural integrity of combustion liner 10.
As indicated in equation (2) above, a ratio of the radius (or diameter) of the cooling hole to separation spacing between cooling holes is also a metric of hole density. A ratio of the radius of each of the plurality of cooling holes to a center-to-center spacing between adjacent cooling holes, as measured in a direction perpendicular to the central axes, can be made to be greater than 0.20, 0.25, and 0.30. To realize such high ratios, the center-to-center spacing of adjacent cooling holes can be less than 0.028 inches (0.711 mm), 0.024 inches (0.610 mm), 0.020 inches (0.508 mm), or 0.016 inches (0.406 mm). Such high densities of cooling holes 18 can improve film cooling of combustion liner 10.

Discussion of Possible Embodiments

The following are non-exclusive descriptions of possible embodiments of the present invention.
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. 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 has a central axis that forms an oblique angle with the top surface of the TBC ceramic layer. A ratio of a combined hole volume of the plurality of cooling holes to a combustion-liner volume of the region of the combustion liner to which the plurality of cooling holes belong is greater than 20 percent.
The combustion liner of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
A further embodiment of the foregoing combustion liner, wherein the ratio of the combined hole volume of the plurality of cooling holes to the volume of the region of the combustion liner to which the plurality of cooling holes belong can be greater than 25 percent.
A further embodiment of any of the foregoing combustion liners, wherein the ratio of the combined hole volume of the plurality of cooling holes to the volume of the region of the combustion liner to which the plurality of cooling holes belong can be greater than 30 percent.
A further embodiment of any of the foregoing combustion liners, wherein the plurality of cooling holes can be arranged in a close-packed configuration.
A further embodiment of any of the foregoing combustion liners, wherein each of the plurality of cooling holes can have six nearest neighbors.
A further embodiment of any of the foregoing combustion liners, wherein the distance between pairs of adjacent holes, as measured in a direction perpendicular to the central axes can be less than .028 inches.
A further embodiment of any of the foregoing combustion liners, wherein the distance between pairs of adjacent holes, as measured in a direction perpendicular to the central axes can be less than .024 inches.
A further embodiment of any of the foregoing combustion liners, wherein the distance between pairs of adjacent holes, as measured in a direction perpendicular to the central axes can be less than .020 inches.
A further embodiment of any of the foregoing combustion liners, wherein the distance between pairs of adjacent holes, as measured in a direction perpendicular to the central axes can be less than .016 inches.
A further embodiment of any of the foregoing combustion liners, a ratio of a radius of each of the plurality of cooling holes to a center-to-center spacing between adjacent cooling holes, as measured in a direction perpendicular to the central axes, can be greater than 20 percent.
A further embodiment of any of the foregoing combustion liners, a ratio of a radius of each of the plurality of cooling holes to a center-to-center spacing between adjacent cooling holes, as measured in a direction perpendicular to the central axes, can be greater than 25 percent.
A further embodiment of any of the foregoing combustion liners, a ratio of a radius of each of the plurality of cooling holes to a center-to-center spacing between adjacent cooling holes, as measured in a direction perpendicular to the central axes, can be greater than 30 percent.
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 oblique angles to the top surface of the TBC ceramic layer. A ratio of a combined hole volume of the plurality of cooling holes to a combustion-liner volume of the region of the combustion liner to which the plurality of cooling holes belong is greater than 20 percent.
The method of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional steps:
A further embodiment of the foregoing method, wherein the ratio of the combined hole volume of the plurality of cooling holes to the volume of the region of the combustion liner to which the plurality of cooling holes belong can be greater than 25 percent.
A further embodiment of any of the foregoing methods, wherein the ratio of the combined hole volume of the plurality of cooling holes to the volume of the region of the combustion liner to which the plurality of cooling holes belong can be greater than 30 percent.
A further embodiment of any of the foregoing methods, wherein the plurality of cooling holes can be arranged in a close-packed configuration.
A further embodiment of any of the foregoing methods, wherein each of the plurality of cooling holes can have six nearest neighbors.
A further embodiment of any of the foregoing methods, wherein the distance between pairs of adjacent holes, as measured in a direction perpendicular to the central axes can be less than .028 inches.
A further embodiment of any of the foregoing methods, wherein the distance between pairs of adjacent holes, as measured in a direction perpendicular to the central axes can be less than .024 inches.
A further embodiment of any of the foregoing methods, wherein the distance between pairs of adjacent holes, as measured in a direction perpendicular to the central axes can be less than .020 inches.
It will be recognized that the invention is not limited to the implementations so described, but can be practiced with modification and alteration without departing from the scope of the appended claims. For example, the above implementations may include specific combination of features. However, the above implementations are not limited in this regard and, in various implementations, the above implementations may include the undertaking only a subset of such features, undertaking a different order of such features, undertaking a different combination of such features, and/or undertaking additional features than those features explicitly listed. The scope of the invention should, therefore, be determined with reference to the appended claims, along with the full scope of equivalents to which such claims are entitled.

Claims

A combustion liner (10) comprising:
a base-alloy substrate (12) having a bottom surface (22) and a top surface (26);

a thermal barrier coat (TBC) metallic layer (14) on the top surface (26) of the base-alloy structure (12);

a TBC ceramic layer (16) on the TBC metallic layer (14), the TBC ceramic layer (16) having a top surface (26) configured to be exposed to combustion gases during operation within a gas turbine engine; and

a plurality of cooling holes (18) formed through the combustion liner (10), each extending from an entrance aperture (20) at the bottom surface (22) of the base-alloy structure (12) and an exit aperture (24) at the top surface (26) of the TBC ceramic layer (16), each of the plurality of cooling holes (18) having a central axis that forms an oblique angle with the top surface (26) of the TBC ceramic layer (16),

wherein a ratio of a combined hole volume of the plurality of cooling holes (18) to a combustion-liner volume of the region of the combustion liner (10) to which the plurality of cooling holes (18) belong is greater than 20 percent.
The combustion liner (10) of claim 1, wherein the ratio of the combined hole volume of the plurality of cooling holes (18) to the volume of the region of the combustion liner (10) to which the plurality of cooling holes (18) belong is greater than 25 percent.
The combustion liner (10) of claim 1, wherein the ratio of the combined hole volume of the plurality of cooling holes (18) to the volume of the region of the combustion liner (10) to which the plurality of cooling holes (18) belong is greater than 30 percent.
The combustion liner (10) of claim 1, 2 or 3, wherein the plurality of cooling holes (18) is arranged in a close-packed configuration, and/or wherein each of the plurality of cooling holes (18) has six nearest neighbors.
The combustion liner (10) of any preceding claim, wherein the distance between pairs of adjacent holes (18), as measured in a direction perpendicular to the central axes is less than .028 inches (.711 mm).
The combustion liner (10) of any of claims 1 to 4, wherein the distance between pairs of adjacent holes (18), as measured in a direction perpendicular to the central axes is less than .016 inches (.406 mm).
The combustion liner (10) of any preceding claim, a ratio of a radius of each of the plurality of cooling holes (18) to a center-to-center spacing between adjacent cooling holes (18), as measured in a direction perpendicular to the central axes, is greater than 20 percent.
The combustion liner (10) of any of claims 1 to 6, a ratio of a radius of each of the plurality of cooling holes (18) to a center-to-center spacing between adjacent cooling holes (18), as measured in a direction perpendicular to the central axes, is greater than 25 percent.
The combustion liner (10) of any of claims 1 to 6, a ratio of a radius of each of the plurality of cooling holes (18) to a center-to-center spacing between adjacent cooling holes (18), as measured in a direction perpendicular to the central axes, is greater than 30 percent.
A method for drilling holes in a combustion liner (10), the method comprising:
providing a base-alloy substrate (12) having a bottom surface (22) and a top surface (26);

depositing a thermal barrier coat (TBC) metallic layer (14) on the top surface (26) of the base-alloy structure (12);

depositing a TBC ceramic layer (16) on the TBC metallic layer (14), the TBC ceramic layer (16) having a top surface (26) configured to be exposed to combustion gases during operation within a gas turbine engine; and

forming a plurality of cooling holes (18) formed through the combustion liner (10) extending from an entrance aperture (20) at the bottom surface (22) of the base-alloy structure (12) and an exit aperture (24) at the top surface (26) of the TBC ceramic layer (16), each of the plurality of cooling holes (18) formed at oblique angles to the top surface (26) of the TBC ceramic layer (16),

wherein a ratio of a combined hole volume of the plurality of cooling holes (18) to a combustion-liner volume of the region of the combustion liner (10) to which the plurality of cooling holes (18) belong is greater than 20 percent.
The method of claim 10, wherein the ratio of the combined hole volume of the plurality of cooling holes (18) to the volume of the region of the combustion liner (10) to which the plurality of cooling holes (18) belong is greater than 25 percent.
The method of claim 10, wherein the ratio of the combined hole volume of the plurality of cooling holes (18) to the volume of the region of the combustion liner (10) to which the plurality of cooling holes (18) belong is greater than 30 percent.
The method of claim 10, 11 or 12, wherein the plurality of cooling holes (18) is arranged in a close-packed configuration.
The method of any of claims 10 to 13, wherein each of the plurality of cooling holes (18) has six nearest neighbors.
The method of any of claims 10 to 14, wherein the distance between pairs of adjacent holes (18), as measured in a direction perpendicular to the central axes is less than .028 inches (.711 mm), or less than .020 inches (.508 mm).