Claims
- 1. 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.
- 2. A gas turbine combustor liner as claimed in claim 1 wherein said circumferential direction coincides with a predetermined swirl angle of the flow.
- 3. 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.
- 4. 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.
- 5. 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.
- 6. 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.
- 7. 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.
- 8. 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.
- 9. 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.
- 10. 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.
- 11. 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.
- 12. 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.
Parent Case Info
This application is a continuation of application Ser. No. 07/614,368, filed Nov. 15, 1990, now abandoned.
US Referenced Citations (17)
Foreign Referenced Citations (2)
Number |
Date |
Country |
9007087 |
Jun 1990 |
WOX |
2221979 |
Feb 1990 |
GBX |
Non-Patent Literature Citations (3)
Entry |
Multihole Cooling Film Effectiveness and Heat Transfer, by R. E. Mayle and F. J. Camarata--Transactions of the ASME--Nov., 1975. |
Alternate Cooling Configuration for Gas Turbine Combustion Systems, by D. A. Nealy, S. B. Reider, H. C. Mongia--Allison Gas Turbine Divn., Prepared by Advisory Group for Aerospace Research & Development 65th Meeting--May 6-10, 1985. |
NASA-CR-159656--Advanced Low--Emissions Catalytic--Combustor Program--Phase I Final Report by G. J. Sturgess--Jun. 1981 report. |
Continuations (1)
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Number |
Date |
Country |
Parent |
614368 |
Nov 1990 |
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