Design Tools for Combustion Stability

Award Information
Agency:
Department of Defense
Branch
n/a
Amount:
$746,333.00
Award Year:
2011
Program:
SBIR
Phase:
Phase II
Contract:
FA9300-11-C-3006
Award Id:
n/a
Agency Tracking Number:
F083-112-1394
Solicitation Year:
2008
Solicitation Topic Code:
AF083-112
Solicitation Number:
2008.3
Small Business Information
3495 Kent Avenue, Suite G100, West Lafayette, IN, -
Hubzone Owned:
N
Minority Owned:
N
Woman Owned:
N
Duns:
132073946
Principal Investigator:
B.J. Austin
President
(765) 775-2107
bjaustin@inspacellc.com
Business Contact:
Amy Austin
Business Manager
(765) 775-2107
aaustin@inspacellc.com
Research Institute:
Stub




Abstract
ABSTRACT: A joint experimental and computational project using a carefully designed, flexible test article is proposed to assess the velocity-coupled combustion response of a representative liquid rocket injector element to transverse acoustic disturbances. In the proposed Phase II, the response of an injector flowfield at supercritical pressure conditions will be simulated, measured, and reduced into a combustion response model applicable to current USAF interests in oxidizer-rich staged-combustion (ORSC) engines. The program directly builds upon Phase I feasibility demonstrations, where the experiment and companion high-resolution CFD simulations showed qualitative agreement on the strong interactions between the transverse waves and the injector flowfield. By taking advantage of our significant ongoing studies of the mechanisms that cause longitudinal instabilities in ORSC engines, this proposal offers an unprecedented opportunity for developing a comprehensive nonlinear combustion response model, validated at realistic conditions, that includes both pressure- and velocity-coupling mechanisms. Key tasks in the proposed effort include the design and fabrication of the experimental combustor; simultaneous and spatially-resolved measurements of the dynamic heat release and flow structures under simulated unstable conditions; hybrid RANS-LES simulations of the experimental configuration and their validation; reduction of the experimental and computational data into approximate combustion response models; and reporting. BENEFIT: With the exception of fuel-rich staged-combustion hydrogen-fueled rocket engines, combustion instability has been a problem for essentially every rocket engine. It may be the most severe technical risk for rocket engine development programs. We offer a revolutionary approach to mitigating this risk by exploiting and combining the exponential growth in computing power, advances in computational algorithms, and a recent experimental breakthrough that allows measurement of stability parameters under simulated unstable conditions to obtain and calculate information critically needed for accurate predictions. This reduction in risk could reduce the rocket engine development time by a factor of five, reduce the cost of engine development by over an order of magnitude, and, perhaps most importantly, help eliminate the severe aversion to risk that has hindered the development and use of truly innovative propulsion devices. If this approach is successful, it will represent a new paradigm in rocket engine development and be used by every engine developer. Related commercial applications that could use the technology demonstrated here include gas turbine propulsion and power generation, image analysis, and any other fields that use reduced-order modeling of complex physical phenomena.

* information listed above is at the time of submission.

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