High Performance Compressor Cooling

Award Information
Agency:
National Aeronautics and Space Administration
Branch
n/a
Amount:
$70,000.00
Award Year:
2001
Program:
SBIR
Phase:
Phase I
Contract:
n/a
Award Id:
52069
Agency Tracking Number:
NASA1865
Solicitation Year:
n/a
Solicitation Topic Code:
n/a
Solicitation Number:
n/a
Small Business Information
45 Manning Road, Billerica, MA, 01821
Hubzone Owned:
N
Minority Owned:
N
Woman Owned:
N
Duns:
n/a
Principal Investigator:
David Stickler
Executive Vice President
(978) 663-9500
dstick@aerodyne.com
Business Contact:
Charles Kolb
President
(978) 663-9500
kolb@aerodyne.com
Research Institution:
n/a
Abstract
High pressure ratio compressor operation is enabled by blade cooling, based on an evaporation / condensation cycle. The blade or blisk temperature is defined by the vaporization temperature of a working fluid internal to the blade, with heat transferred to an integral condenser at each blade root. This approach provides heat rejection to a cool, low pressure fluid, such as engine lubricating oil, with subsequent transfer to the fuel flow for complete energy recovery. Conventional film cooling technology, as developed for turbine blades using compressor bleed air, is not applicable to high the pressure compressor stage, due to lack of an acceptable high pressure cooler air supply. This approach to compressor cooling essentially eliminates the usual blade structure temperature constraint on overall compression ratio. This basic cooling technology has also been explored for application to the rotating turbine stages, with substantial cycle benefits identified. This blade cooling approach thus enables turbomachinery operation at high flight mach number, at high overall compression ratio, with very high turbine inlet temperature. The result is high efficiency air breathing propulsion at high flight Mach number.

* information listed above is at the time of submission.

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