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Full-Text Articles in Engineering
Direct Numerical Simulation Of Roughness Induced Hypersonic Boundary Layer Transition On A 7° Half-Angle Cone, Tara E. Crouch
Direct Numerical Simulation Of Roughness Induced Hypersonic Boundary Layer Transition On A 7° Half-Angle Cone, Tara E. Crouch
Theses and Dissertations
Direct numerical simulation (DNS) computational fluid dynamic (CFD) calculations were performed on a 30° slice of 7° half-angle cones with increasing nose radii bluntness at Mach 10 while simulating a distributed roughness pattern on the cone surface. These DNS computations were designed to determine if the non-modal transition behavior observed in testing performed at the Arnold Engineering Development Center (AEDC) Hypervelocity Wind Tunnel 9 was induced via distributed surface roughness. When boundary layer transition is dominated by second mode instabilities, an increase in nose radius delays the transition location downstream. However, blunt nose experiments indicated that as the nose radius …
Comparison Of Flow Field In The Proximity Of A Single Planar & Wrap-Around Fin, Nayhel Sharma, Palak Saini, Hrishabh Chaudhary, Gurteg Nagi, Rakesh Kumar Dr.
Comparison Of Flow Field In The Proximity Of A Single Planar & Wrap-Around Fin, Nayhel Sharma, Palak Saini, Hrishabh Chaudhary, Gurteg Nagi, Rakesh Kumar Dr.
International Journal of Aviation, Aeronautics, and Aerospace
Abstract
This paper analyses the results of the computational analysis between a single planar and a wrap-around fin mounted on a semi-cylindrical body. A free-stream Computational Fluid Dynamics (CFD) model was simulated for both cases in the Mach 0.4-3.0M range at 0°Angle of attack, in which, the behavior of flow around the fin was investigated using a turbulence model of higher order discretization. The post-processing shows all the possible views of the flow dynamics around the fins, as well as the missile body. The aerodynamic drag and the rolling moment characteristics of the planar and the wrap-around fin have been …
A Comparison Of The Aerodynamic Centers For Panel Code Compressible Corrections And Openfoam 5 For Mach 0.1 To 0.8, Dustin Weaver
A Comparison Of The Aerodynamic Centers For Panel Code Compressible Corrections And Openfoam 5 For Mach 0.1 To 0.8, Dustin Weaver
All Graduate Plan B and other Reports, Spring 1920 to Spring 2023
It is known that the aerodynamic center changes from quarter chord to half chord from incompressible to compressible flows on airfoils. Compressible corrections are derived and implemented in a vortex panel code. These results will be used to find the aerodynamic centers for the specified Mach range of 0.1 to 0.8 in 0.1 increments within - 6 to 6 degrees angle of attack. OpenFOAM 5 cases will be created with specific meshes and settings. The results calculated from OpenFOAM 5 will be compared to the results obtained from the compressible corrections.