Advanced Turbine Blade Cooling Techniques

Award Information
Agency:
National Aeronautics and Space Administration
Branch
n/a
Amount:
$99,321.00
Award Year:
2011
Program:
SBIR
Phase:
Phase I
Contract:
NNX11CD93P
Award Id:
n/a
Agency Tracking Number:
105328
Solicitation Year:
2010
Solicitation Topic Code:
A2.10
Solicitation Number:
n/a
Small Business Information
CA, Huntington Beach, CA, 92647-7738
Hubzone Owned:
N
Minority Owned:
N
Woman Owned:
N
Duns:
001557268
Principal Investigator:
DavidUnderwood
Principal Investigator
(714) 847-9945
daveunderwood@microcoolingconcepts.com
Business Contact:
JackFryer
Business Official
(714) 847-9945
jayfryer@microcoolingconcepts.com
Research Institute:
Stub




Abstract
Gas turbine engine technology is constantly challenged to operate at higher combustor outlet temperatures. In a modern gas turbine engine, these temperatures can exceed the blade and disk material limits by 600<SUP>o</SUP>F or more, necessitating both internal and film cooling schemes in addition to the use of thermal barrier coatings. Internal convective cooling is inadequate in many blade locations, and both internal and film cooling approaches can lead to significant performance penalties in the engine.Micro Cooling Concepts has developed a turbine blade cooling concept that provides enhanced internal impingement cooling effectiveness via the use of micro-structured impingement surfaces. These surfaces significantly increase the cooling capability of the impinging flow, as compared to a conventional untextured surface. This approach can be combined with microchannel cooling and external film cooling to tailor the cooling capability per the external heating profile. The cooling system can then be optimized to minimize impact on engine performance.

* information listed above is at the time of submission.

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