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Publication numberUS20050220618 A1
Publication typeApplication
Application numberUS 10/813,131
Publication dateOct 6, 2005
Filing dateMar 31, 2004
Priority dateMar 31, 2004
Publication number10813131, 813131, US 2005/0220618 A1, US 2005/220618 A1, US 20050220618 A1, US 20050220618A1, US 2005220618 A1, US 2005220618A1, US-A1-20050220618, US-A1-2005220618, US2005/0220618A1, US2005/220618A1, US20050220618 A1, US20050220618A1, US2005220618 A1, US2005220618A1
InventorsXiuzhang Zhang, Mark Duer, Doyle Lewis
Original AssigneeGeneral Electric Company
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Counter-bored film-cooling holes and related method
US 20050220618 A1
Abstract
A turbine component includes a plurality of film-cooling holes formed in a region of the component to be cooled, the cooling holes having specified diameter, each hole at an exit thereof formed with a counter-bore of predetermined depth; the component having a coating applied thereto at least in the region, wherein the counter-bore provides an area for excess coating material to accumulate without reducing the specified diameter. A method of maintaining cooling efficiency of film-cooling holes in a turbine component where the film-cooling holes have specified diameters and the turbine component has a protective coating therein comprising: a) before coating, forming each film-cooling hole with a counter-bore and an exit end of the film-cooling hole; and b) spraying the coating onto the turbine component at least in areas surrounding the film-cooling holes such that excess coating material accumulates in the counter-bore without reducing the specified diameter of the cooling holes.
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Claims(13)
1. A turbine component having a plurality of film-cooling holes formed in a region of the component to be cooled, said cooling holes having specified diameter, each hole at an exit thereof formed with a counter-bore of predetermined depth; said component having a coating applied thereto at least in said region, wherein the counter-bore provides an area for excess coating material to accumulate without reducing the specified diameter.
2. The turbine component of claim 1 wherein, for a specified diameter of about 0.033 inch, the counter-bore has a diameter of about 0.053 inch.
3. The turbine component of claim 2 wherein the counter-bore has a depth of about 0.030 inch.
4. The turbine component of claim 2 wherein said coating comprises a first bondcoat layer and a second thermal barrier coating layer.
5. The turbine component of claim 4 wherein the bondcoat layer is an NiAl-based material.
6. The turbine component of claim 5 wherein the thermal barrier coating layer is a yttria stabilized zirconium material.
7. The turbine component of claim 1 wherein the turbine component comprises a gas turbine bucket having an airfoil portion and a shank portion, and wherein said region comprises the airfoil portion.
8. The turbine component of claim 7 wherein, for a specified diameter of about 0.033 inch, the counter-bore has a diameter of about 0.053 inch; and wherein the counter-bore has a depth of about 0.030 inch.
9. The turbine component of claim 7 wherein said coating comprises a first bondcoat layer and a second thermal barrier coating layer; and wherein the bondcoat layer is an NiAl-based material.
10. A gas turbine bucket having an airfoil portion and a shank portion, said airfoil portion having a plurality of film-cooling holes therein, each hole at an exit thereof formed with a counter-bore of predetermined depth; said component having a coating applied thereto at least in said region, wherein the counter-bore provides an area for excess coating material to accumulate without reducing the specified diameter; and wherein said coating comprises a first bondcoat layer and a second thermal barrier coating layer.
11. The gas turbine bucket of claim 10 wherein the bondcoat layer is an NiAl-based material.
12. The gas turbine bucket of claim 11 wherein the thermal barrier coating layer is a yttria-stabilized zirconium layer.
13. A method of maintaining cooling efficiency of film-cooling holes in a turbine component where the film-cooling holes have specified diameters and the turbine component has a protective coating therein comprising:
a) before coating, forming each film-cooling hole with a counter-bore and an exit end of the film-cooling hole; and
b) spraying the coating onto the turbine component at least in areas surrounding the film-cooling holes such that excess coating material accumulates in the counter-bore without reducing the specified diameter of the cooling holes.
Description
BACKGROUND OF THE INVENTION

This invention relates to the configuration of film-cooling holes utilized as part of the cooling circuit in the airfoil portion of a turbine blade or bucket.

Film-cooling has been a major aspect of gas turbine cooling for many years. The application of effective film-cooling techniques provides the first and best line of defense for hot gas path surfaces against the onslaught of extreme heat fluxes, serving to directly reduce the incident convective heat flux on the surface. As film-cooling holes first go into service, they are typically cleaned of all obstructions or unwanted debris. Film holes in this condition may also include certain protective coatings, either diffusion or thermal barrier coatings (TBC), for such purposes as oxidation protection. In operation, film holes and hot gas path surfaces see a multitude of conditions and environments which can result in the sudden or gradual blockage of holes to various degrees, thereby influencing the film-cooling performance to lesser or greater extents.

It has been discovered, however, that as a TBC coating is sprayed on the airfoil of the bucket, some of the coating material enters the exit of the film-cooling holes. Thus, the TBC adheres to the inside surface of the film-cooling holes, decreasing the effective exit area of the holes and reducing the film-cooling effect from the design intent.

BRIEF DESCRIPTION OF THE INVENTION

The present invention solves the partial obstruction of film-cooling holes due to TBC's sprayed on the airfoils of the buckets by changing the configuration of the film-cooling holes to include a counter-bore at the outlet or exit ends of the film-cooling holes. It is contemplated that the counter bores would be applicable to the holes along the leading edge shower head, gill holes and the holes around the bucket tip region. The counter-bore diameter and depth are specific to the design, and have been optimized for performance. For example, in one exemplary embodiment, a counter-bore of 0.053 inches on a 0.033 inch diameter through-hole extends 0.03 inches from the outlet surface of the airfoil on the minimal dimension.

The general concept of incorporating a counter-bore or flared shape can be applied to all film-cooling holes for various gas turbine buckets, nozzles and shrouds with TBC application on those parts.

Accordingly, in one aspect, the invention relates to a turbine component having a plurality of film-cooling holes formed in a region of the component to be cooled, the cooling holes having specified diameter, each hole at an exit thereof formed with a counter-bore of predetermined depth; the component having a coating applied thereto at least in the region, wherein the counter-bore provides an area for excess coating material to accumulate without reducing the specified diameter.

In another aspect, the invention relates to a gas turbine bucket having an airfoil portion and a shank portion, the airfoil portion having a plurality of film-cooling holes therein, each hole at an exit thereof formed with a counter-bore of predetermined depth; the component having a coating applied thereto at least in the region, wherein the counter-bore provides an area for excess coating material to accumulate without reducing the specified diameter; and wherein the coating comprises a first bondcoat layer and a second thermal barrier coating layer.

In still another aspect, the invention relates to a method of maintaining cooling efficiency of film-cooling holes in a turbine component where the film-cooling holes have specified diameters and the turbine component has a protective coating therein comprising: a) before coating, forming each film-cooling hole with a counter-bore and an exit end of the film-cooling hole; and b) spraying the coating onto the turbine component at least in areas surrounding the film-cooling holes such that excess coating material accumulates in the counter-bore without reducing the specified diameter of the cooling holes.

The invention will now be described in detail in connection with the drawings identified below.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a gas turbine bucket with film-cooling holes along the leading edge of the bucket airfoil; and

FIG. 2 is a sketch through the centerline of a film-cooling hole in a test plate showing the build-up of a TBC coating in the hole; and

FIG. 3 is a cross section view of a film-cooling hole on the leading edge of a bucket in accordance with one embodiment of the invention.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to FIG. 1, there is illustrated a turbine bucket 10 constructed in accordance with the present invention including an airfoil 12 mounted on a platform 14. The turbine bucket also includes forward and aft wheel space seals, i.e., angel wings 16, 18, respectively. The buckets 10 are adapted for mounting on the turbine wheel in conventional fashion. The airfoil 12 has a profile including a compound curvature with pressure and suction sides 20, 22, respectively, as well as a leading edge 24 and trailing edge 26.

It is known to apply a thermal barrier coating (TBC) to various regions of the bucket including adjacent the leading edge 24. Typical TBC's include a first bondcoat layer and a second ceramic coating layer. The bondcoat layer may be an NiAl-based bondcoat, and the thermal barrier coating layer may be a yttria-stabilized zirconium layer. It is also known to provide film-cooling holes 28 in various regions of the bucket including but not limited to the leading edge 24. FIG. 1 illustrates a few representative film-cooling holes 28 for purposes of ease of understanding of the invention.

FIG. 2 is a sketch of a conventionally-shaped film-cooling hole 28 formed in a test plate along a simulated leading edge 24. It can be seen that excess TBC material 30 that coats the simulated leading edge area accumulates within a portion of the hole 28, resulting in a decreased effective diameter of the hole at the outlet thereof. In a real cooling hole on a bucket, this condition decreases cooling efficiency.

FIG. 3 illustrates in detail one example of a film-cooling hole in accordance with the invention. Film-cooling hole 32 is shown to be located at the radially outer end of a bucket 34, along the leading edge 36. The film-cooling hole has a nominal diameter d that extends outwardly from an internal region 38 of the bucket. In accordance with the invention, film-cooling hole 32 is counter-bored to a diameter d1 from its outlet on the leading edge 36 inwardly to a predetermined depth h. In a typical example, the nominal diameter d of the film cooing hole 32 is 0.033 in. For this size cooling hole, the counter-bore 40 has been formed with a diameter d1, of 0.053 in. It will be appreciated that the dimensional relationships and dimensions themselves may be varied to suit different size buckets. It will further be appreciated that the invention is applicable to cooling holes in any areas on any components that are coated. The counter-bore will provide adequate space to accommodate excess TBC coating material without reduction of the effective cooling flow.

While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.

Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US7216485 *Sep 3, 2004May 15, 2007General Electric CompanyAdjusting airflow in turbine component by depositing overlay metallic coating
US7625178Aug 30, 2006Dec 1, 2009Honeywell International Inc.High effectiveness cooled turbine blade
US7997867Oct 17, 2007Aug 16, 2011Iowa State University Research Foundation, Inc.Momentum preserving film-cooling shaped holes
US8066478Oct 17, 2007Nov 29, 2011Iowa State University Research Foundation, Inc.Preventing hot-gas ingestion by film-cooling jet via flow-aligned blockers
US8672613Aug 31, 2010Mar 18, 2014General Electric CompanyComponents with conformal curved film holes and methods of manufacture
US20100192588 *Feb 1, 2010Aug 5, 2010Rolls-Royce Deutschland Ltd & Co KgMethod for the provision of a cooling-air opening in a wall of a gas-turbine combustion chamber as well as a combustion-chamber wall produced in accordance with this method
Classifications
U.S. Classification416/97.00R
International ClassificationB63H1/14, F01D5/18
Cooperative ClassificationF05D2260/202, F01D5/186
European ClassificationF01D5/18F
Legal Events
DateCodeEventDescription
Mar 31, 2004ASAssignment
Owner name: GENERAL ELECTRIC COMPANY, NEW YORK
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:ZHANG, XIUZHANG JAMES;DUER, MARK GERARD;LEWIS, DOYLE CLYDE;REEL/FRAME:015171/0968;SIGNING DATES FROM 20030325 TO 20040330