PSE Analysis of a Quiet Tunnel Hypersonic Boundary Layer Flow PDF Download
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Author: Melissa L. Manning
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ISBN:
Category :
Languages : en
Pages :
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Book Description
Author: Melissa L. Manning
Publisher:
ISBN:
Category :
Languages : en
Pages :
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Book Description
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages :
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Book Description
A computational evaluation of two stability experiments conducted in the NASA Langley Mach 6 axisymmetric quiet nozzle test chamber facility is conducted. Navier-Stokes analysis of the mean flow and linear stability theory analysis of boundary layer disturbances is performed in the computations. The effects of adverse pressure gradient and wall cooling are examined. Calculated pressure, temperature and boundary layer thickness distributions show very good overall agreement with experimental measurements. Computed mass flux and total temperature profiles show very good quantitative agreement with uncalibrated hot-wire measurements obtained with the hot-wire operated in high and low overheat modes respectively. Comparisons between calibrated hot-wire data and mean flow computations show excellent agreement in the early stages of the transitional flow. However, examination of the wire Reynolds number and mass flux and total temperature eigenfunction profiles suggest that when operated in high overheat mode the sensitivity of the hot-wire to total temperature is significant. Thus, while uncalibrated hot-wire measurements are useful to characterize the overall features of the flow, calibrated hot-wire measurements are necessary for quantitative comparison with stability theory. Computations show that adverse pressure gradient and wall cooling decrease the boundary layer thickness and increase the frequency and amplification rate of the unstable second mode disturbances; these findings are consistent with the experimental observations.
Author: Heath B. Johnson
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Category :
Languages : en
Pages :
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Author: Melissa L. Manning
Publisher:
ISBN: 9781423527459
Category :
Languages : en
Pages : 118
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Book Description
A computational evaluation of two stability experiments conducted in the NASA Langley Mach 6 axisymmetric quiet nozzle test chamber facility is conducted. Navier- Stokes analysis of the mean flow and linear stability theory analysis of boundary layer disturbances is performed in the computations. The effects of adverse pressure gradient and wall cooling are examined. Calculated pressure, temperature and boundary layer thickness distributions show very good overall agreement with experimental measurements. Computed mass flux and total temperature profiles show very good quantitative agreement with uncalibrated hot- wire measurements obtained with the hot-wire operated in high and low overheat modes respectively. Comparisons between calibrated hot-wire data and mean flow computations show excellent agreement in the early stages of the transitional flow. However, examination of the wire Reynolds number and mass flux and total temperature eigenfunction profiles suggest that when operated in high overheat mode the sensitivity of the hot-wire to total temperature is significant. Thus, while uncalibrated hot-wire measurements are useful to characterize the overall features of the flow, calibrated hot- wire measurements are necessary for quantitative comparison with stability theory. Computations show that adverse pressure gradient and wall cooling decrease the boundary layer thickness and increase the frequency and amplification rate of the unstable second mode disturbances; these findings are consistent with the experimental observations.
Author: National Aeronautics and Space Administration (NASA)
Publisher: Createspace Independent Publishing Platform
ISBN: 9781722024642
Category :
Languages : en
Pages : 30
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A flared-cone model under adiabatic and cooled-wall conditions was placed in a calibrated, low-disturbance Mach 6 flow and the stability of the boundary layer was investigated using a prototype constant-voltage anemometer. The results were compared with linear-stability theory predictions and good agreement was found in the prediction of second-mode frequencies and growth. In addition, the same 'N = 10' criterion used to predict boundary-layer transition in subsonic, transonic, and supersonic flows under low freestream noise conditions was found to be applicable for the hypersonic flow regime as well. Under cooled-wall conditions, a unique set of spectral data was acquired that documents the linear, nonlinear, and breakdown regions associated with the transition of hypersonic flow under low-noise conditions. Blanchard, Alan E. and Selby, Gregory V. and Wilkinson, Stephen P. Langley Research Center NCC1-180...
Author: Doyle D. Knight
Publisher:
ISBN:
Category : Hypersonic wind tunnels
Languages : en
Pages : 18
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Book Description
The report summarizes a research program focused on two main issues related to the achievement of high Reynolds number quiet flow in hypersonic tunnels, namely, 1) automated optimal design of quiet hypersonic tunnels, and 2) tunnel startup in the presence of blunt models. In the former case, a fully automated optimal design methodology was developed to determine the optimal shape of the supersonic nozzle to achieve laminar flow on the nozzle walls and hence quiet flow in the test section. In the latter case, detailed time-accurate simulations were performed for a cone model of different angles in the Purdue Quiet Tunnel. Results demonstrated the limiting size of the model that permitted startup of the tunnel.
Author: John C. Adams
Publisher:
ISBN:
Category : Aerodynamics, Hypersonic
Languages : en
Pages : 98
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An analytical approach toward numerical calculation of the three-dimensional turbulent boundary layer on a sharp cone at incidence under supersonic and hypersonic flow conditions is presented. The theoretical model is based on implicit finite-difference integration of the governing three-dimensional turbulent boundary-layer equations in conjunction with a three-dimensional scalar eddy-viscosity model of turbulence. Comparison is made of present theory with detailed experimental measurements of the three-dimensional turbulent boundary-layer structure (velocity and temperature profiles), the surface streamline direction (obtained via an oil-flow technique) and surface heat-transfer rate.
Author: Ndaona Chokani
Publisher:
ISBN:
Category :
Languages : en
Pages :
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Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 13
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This grant was redirected to focus on the search for high-Reynolds-number quiet flow in the Boeing/AFOSR Mach-6 Quiet Tunnel at Purdue. Quiet flow with freestream noise levels comparable to flight requires maintaining laminar nozzle-wall boundary layers; this becomes increasingly difficult, and increasingly useful, as the Reynolds number increases. After nearly five years of shakedown, quiet flow was finally achieved to a freestream unit Reynolds number of 2.8 million per foot, in early 2006. Although this is 90% of the prefabrication design value, it is achieved only intermittently. The masimum feasible quiet-flow Reynolds number remains to be determined, along with the conditions for achieving it reliably. Nevertheless, the facility is presently the only hypersonic quiet tunnel, anywhere in the world, and affordable operating costs have been maintained.
Author: National Aeronautics and Space Administration (NASA)
Publisher: Createspace Independent Publishing Platform
ISBN: 9781724242556
Category :
Languages : en
Pages : 232
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Book Description
Hypersonic boundary layer measurements over a flared cone were conducted in a Mach 6 quiet wind tunnel at a freestream unit Reynolds number of 2.82 million/ft. This Reynolds number provided laminar-to-transitional flow over the cone model in a low-disturbance environment. Four interchangeable nose-tips, including a sharp-tip, were tested. Point measurements with a single hot-wire using a novel constant voltage anemometer were used to measure the boundary layer disturbances. Surface temperature and schlieren measurements were also conducted to characterize the transitional state of the boundary layer and to identify instability modes. Results suggest that second mode disturbances were the most unstable and scaled with the boundary layer thickness. The second mode integrated growth rates compared well with linear stability theory in the linear stability regime. The second mode is responsible for transition onset despite the existence of a second mode subharmonic. The subharmonic disturbance wavelength also scales with the boundary layer thickness. Furthermore, the existence of higher harmonics of the fundamental suggests that nonlinear disturbances are not associated with 'high' free stream disturbance levels. Nose-tip radii greater than 2.7% of the base radius completely stabilized the second mode. Lachowicz, Jason T. and Chokani, Ndaona Langley Research Center NASA-CR-198272, NAS 1.26:198272 NCC1-183; RTOP-505-59-50-02...
