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Experimental, Theoretical, and Computational Investigation of Separated Nozzle FlowsA detailed experimental, theoretical, and computational study of separated nozzle flows has been conducted. Experimental testing was performed at the NASA Langley 16-Foot Transonic Tunnel Complex. As part of a comprehensive static performance investigation, force, moment, and pressure measurements were made and schlieren flow visualization was obtained for a sub-scale, non-axisymmetric, two-dimensional, convergent- divergent nozzle. In addition, two-dimensional numerical simulations were run using the computational fluid dynamics code PAB3D with two-equation turbulence closure and algebraic Reynolds stress modeling. For reference, experimental and computational results were compared with theoretical predictions based on one-dimensional gas dynamics and an approximate integral momentum boundary layer method. Experimental results from this study indicate that off-design overexpanded nozzle flow was dominated by shock induced boundary layer separation, which was divided into two distinct flow regimes; three- dimensional separation with partial reattachment, and fully detached two-dimensional separation. The test nozzle was observed to go through a marked transition in passing from one regime to the other. In all cases, separation provided a significant increase in static thrust efficiency compared to the ideal prediction. Results indicate that with controlled separation, the entire overexpanded range of nozzle performance would be within 10% of the peak thrust efficiency. By offering savings in weight and complexity over a conventional mechanical exhaust system, this may allow a fixed geometry nozzle to cover an entire flight envelope. The computational simulation was in excellent agreement with experimental data over most of the test range, and did a good job of modeling internal flow and thrust performance. An exception occurred at low nozzle pressure ratios, where the two-dimensional computational model was inconsistent with the three-dimensional separation observed in the experiment. In general, the computation captured the physics of the shock boundary layer interaction and shock induced boundary layer separation in the nozzle, though there were some differences in shock structure compared to experiment. Though minor, these differences could be important for studies involving flow control or thrust vectoring of separated nozzles. Combined with other observations, this indicates that more detailed, three-dimensional computational modeling needs to be conducted to more realistically simulate shock-separated nozzle flows.
Document ID
20040090513
Acquisition Source
Langley Research Center
Document Type
Other
Authors
Hunter, Craig A. (NASA Langley Research Center Hampton, VA, United States)
Date Acquired
September 7, 2013
Publication Date
January 1, 2004
Subject Category
Fluid Mechanics And Thermodynamics
Report/Patent Number
AIAA Paper 98-3107
Meeting Information
Meeting: 34th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit
Location: Cleveland, OH
Country: United States
Start Date: July 13, 1998
End Date: July 15, 1998
Sponsors: Society of Automotive Engineers, Inc., American Society of Mechanical Engineers, American Society for Electrical Engineers, American Inst. of Aeronautics and Astronautics
Distribution Limits
Public
Copyright
Public Use Permitted.

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