US5233828A
Expired / Lapsed
US
⏱️ 0.0 years remaining
Combustor liner with circumferentially angled film cooling holes
Assignee: General Electric Co
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Filing: 1992-09-24
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Granted: 1993-08-10
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Expires: 2012-09-24
Abstract
Abstract A gas turbine engine combustor is provided, having a an annular single wall sheet metal liner which is generally annular in shape and having disposed therethrough a multi-hole film cooling means which includes at least one pattern of small closely spaced film cooling holes angled sharply in the downstream direction and angled in a circumferential direction wherein the circumferential angle generally coincides with the swirl angle of the flow along the surface of the liner. Another embodiment provides a corrugated aircraft engine sheet metal combustor liner which forms an axially extending wavy wall to help resist buckling, particularly useful for outer liners in the combustion section of the engine and in the exhaust section of gas turbine engines incorporating afterburners.
Inventors
- 👤 Phillip D. Napoli
Bibliographic Details
Application Number
—
Jurisdiction
US
Data Source
google_patents
Last Synchronized
2026-11-04 17:54
CPC Classifications
F; F23; F23R; F23R3/00; F23R3/002; F02; F02K; F02K1/00; F02K1/78; F02K1/82; F02K1/822; F05; F05D; F05D2260/00; F05D2260/20; F05D2260/202; F23R2900/00; F23R2900/00005; Y; Y02; Y02T; Y02T50/00; Y02T50/60
Claims Analysis
Total Claims: 14
Claim 1
Independent Claim
119 words
An annular gas turbine combustor liner for containing a hot combustor flow, said liner comprising: a single wall annular shell having a hot surface and a cold surface and at least one continuous pattern of small closely spaced film cooling holes angled sharply in the downstream direction from said cold surface to said hot surface, said continuous pattern effective to produce a cooling film extending substantially over the entire length of said shell, said film cooling holes having a hole diameter, a downstream slant angle, and spaced at least sufficiently close enough together to effect a cooling film on said hot surface of said shell during combustor operation, and said film cooling holes being angled in a circumferential direction.
Claim 2
Dependent Claim
119 words
An annular gas turbine combustor liner for containing a hot combustor flow, said liner comprising: a single wall annular shell having a hot surface and a cold surface and at least one continuous pattern of small closely spaced film cooling holes angled sharply in the downstream direction from said cold surface to said hot surface, said continuous pattern effective to produce a cooling film extending substantially over the entire length of said shell, said film cooling holes having a hole diameter, a downstream slant angle, and spaced at least sufficiently close enough together to effect a cooling film on said hot surface of said shell during combustor operation, and said film cooling holes being angled in a circumferential direction.
Claim 3
Dependent Claim
23 words
A gas turbine combustor liner as claimed in claim 1 wherein said circumferential direction coincides with a predetermined swirl angle of the flow.
Claim 4
Dependent Claim
38 words
A gas turbine combustor liner as claimed in claim 1 wherein the circumferential biased angle is in a range of between 30 and 65 degrees measured generally from a downstream component of the flow's direction in the combustor.
Claim 5
Dependent Claim
25 words
A gas turbine combustor liner as claimed in claim 1 wherein a portion of said shell is corrugated to form a shallow wavy wall cross-section.
Claim 6
Dependent Claim
46 words
A gas turbine combustor liner as closed in claim 3 wherein said film cooling holes have a downstream angle slanted from said cold surface of said shell to said hot surface of said shell and wherein said downstream angle has a value of about twenty degrees.
Claim 7
Dependent Claim
49 words
A gas turbine combustor liner as claimed in claim 3 wherein said film cooling holes have a downstream angle slanted from said cold surface of said shell to said hot surface of said shell and wherein said downstream angle has a value in a range of about twenty degrees.
Claim 8
Dependent Claim
25 words
A gas turbine combustor liner as claimed in claim 6 wherein a portion of said shell is corrugated to form a shallow wavy wall cross-section.
Claim 9
Dependent Claim
128 words
An afterburning gas turbine engine exhaust section combustor liner for containing a hot combustor flow, said exhaust section combustor liner comprising: a single wall sheet metal shell having a hot surface and a cold surface wherein a portion of said shell is corrugated to form a shallow wavy wall cross-section and at least one pattern of small closely spaced sharply downstream angled film cooling holes disposed through said shell having a downstream angle slanted from said cold surface of said shell to said hot surface of said shell wherein said downstream angle has a value of about twenty degrees and said film cooling holes are angled in a circumferential direction, said continuous pattern effective to produce a cooling film extending substantially over the entire length of said shell.
Claim 10
Dependent Claim
128 words
An afterburning gas turbine engine exhaust section combustor liner for containing a hot combustor flow, said exhaust section combustor liner comprising: a single wall sheet metal shell having a hot surface and a cold surface wherein a portion of said shell is corrugated to form a shallow wavy wall cross-section and at least one pattern of small closely spaced sharply downstream angled film cooling holes disposed through said shell having a downstream angle slanted from said cold surface of said shell to said hot surface of said shell wherein said downstream angle has a value of about twenty degrees and said film cooling holes are angled in a circumferential direction, said continuous pattern effective to produce a cooling film extending substantially over the entire length of said shell.
Claim 11
Dependent Claim
30 words
An afterburning gas turbine engine exhaust section combustor liner as claimed in claim 8 wherein said circumferential direction coincides with a predetermined swirl direction of the flow in the combustor.
Claim 12
Dependent Claim
38 words
A gas turbine combustor liner as claimed in claim 9 wherein the circumferential biased angle is in a range of between 30 and 65 degrees measured generally from a downstream component of the flow's direction in the combustor.
Claim 13
Dependent Claim
50 words
An afterburning gas turbine engine exhaust section combustor liner as claimed in claim 8 wherein said film cooling holes have a downstream angle slanted from said cold surface of said shell to said hot surface of said shell and wherein said downstream angle has a value of about fifteen degrees.
Claim 14
Dependent Claim
57 words
An afterburning gas turbine engine exhaust section combustor liner as claimed in claim 8 wherein said film cooling holes have a downstream angle slanted from said cold surface of said shell to said hot surface of said shell and wherein said downstream angle has a value in a preferred range of about between ten and twenty degrees.
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USPTO Annuity & Maintenance Windows
| Annuity Stage | Window Opens | Payment Deadline | Surcharge Period Until | Estimated Fee | Status |
|---|---|---|---|---|---|
| 3.5 Years (Year 4) | 1996-08-09 | 1997-02-08 | 1997-08-10 | $2000.00 | Window Closed / Lapsed |
| 7.5 Years (Year 8) | 2000-08-09 | 2001-02-08 | 2001-08-10 | $3760.00 | Window Closed / Lapsed |
| 11.5 Years (Year 12) | 2004-08-09 | 2005-02-08 | 2005-08-10 | $7700.00 | Window Closed / Lapsed |