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  • 1
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    An LDA investigation of the normal shock wave boundary layer interaction (1990)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: Nonintrusive measurements have been made of two normal shock wave-boundary layer interactions. Two-dimensional measurements were made throughout the interaction region while three-dimensional measurements were made in the vicinity of the shock wave. The measurements were made in the corner of the test section of a continuous flow supersonic wind tunnel in which a normal shock wave had been stabilized. LDA, surface pressure measurement and flow visualization techniques were employed for two freestream Mach number test cases: 1.6 and 1.3. The former contained separated flow regions and a system of shock waves. The latter was found to be far less complicated. The reported results define the flowfield structure in detail for each case.
    Keywords: AERODYNAMICS
    Type: (ISSN 0379-380X); 2, 19
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Heat transfer, velocity-temperature correlation, and turbulent shear stress from Navier-Stokes computations of shock wave/turbulent boundary layer interaction flows (1991)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The properties of 2-D shock wave/turbulent boundary layer interaction flows were calculated by using a compressible turbulent Navier-Stokes numerical computational code. Interaction flows caused by oblique shock wave impingement on the turbulent boundary layer flow were considered. The oblique shock waves were induced with shock generators at angles of attack less than 10 degs in supersonic flows. The surface temperatures were kept at near-adiabatic (ratio of wall static temperature to free stream total temperature) and cold wall (ratio of wall static temperature to free stream total temperature) conditions. The computational results were studied for the surface heat transfer, velocity temperature correlation, and turbulent shear stress in the interaction flow fields. Comparisons of the computational results with existing measurements indicated that (1) the surface heat transfer rates and surface pressures could be correlated with Holden's relationship, (2) the mean flow streamwise velocity components and static temperatures could be correlated with Crocco's relationship if flow separation did not occur, and (3) the Baldwin-Lomax turbulence model should be modified for turbulent shear stress computations in the interaction flows.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: Computational Fluid Dynamics Symposium on Aeropropulsion; p 429-456
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    An LDA (Laser-Doppler Anemometry) investigation of three-dimensional normal shock wave boundary-layer interactions (1989)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: Nonintrusive measurements were made of a normal shock wave/boundary layer interaction. Two dimensional measurements were made throughout the interaction region while 3-D measurements were made in the vicinity of the shock wave. The measurements were made in the corner of the test section of a continuous supersonic wind tunnel in which a normal shock wave had been stabilized. Laser Doppler Anemometry, surface pressure measurement and flow visualization techniques were employed for two freestream Mach number test cases: 1.6 and 1.3. The former contained separated flow regions and a system of shock waves. The latter was found to be far less complicated. The results define the flow field structure in detail for each case.
    Keywords: AERODYNAMICS
    Type: NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 2; p 741-764
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  • 4
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    Development of seeding techniques for small supersonic wind tunnel (1985)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The NASA Lewis 1x1 foot supersonic wind tunnel is used to experimentally verify computational methods. This tunnel, which is continuous running, operates from laboratory-wide high pressure air and vacuum systems. As such, the air does not recirculate but makes a single pass through the tunnel. The Mach number is varied with interchangeable nozzle blocks and has a range from Mach 1.6 to 4.0. Dry and filtered air is available up to pressures of 3 atmospheres. The air enters the tunnel system through a plenum having flow straighteners and 6 fine mesh screens. The exit of the plenum provides smooth contraction with an area ratio of approximately 20 that, along with the screens, provides a uniform flow for the nozzle.
    Keywords: INSTRUMENTATION AND PHOTOGRAPHY
    Type: NASA. Langley Research Center Wind Tunnel Seeding Systems for Laser Velocimeters; p 189-202
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Progress toward synergistic hypermixing nozzles (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Mean flow measurements have been obtained for air-to-air mixing downstream of swept and unswept ramp wall-mounted hypermixing nozzle configurations. Aside from the sweep of the ramps, the two nozzle configurations investigated are identical. The nozzles inject three parallel supersonic jets at a 15-deg angle (relative to the wind tunnel wall) into a supersonic freestream. Mach number and volume fraction distributions in a transverse plane 11.1 nozzle heights downstream from the nozzle exit plane were measured. Data are presented for a freestream Mach number of three at a matched static pressure condition and also at an underexpanded static pressure condition (pressure ratio equal to 5). Surface oil flow visualization was used to investigate the near-wall flow behavior. The results indicate that the swept ramp injectors produce stronger and larger vortex pairs than the unswept ramp injectors. The increased interaction between the swept ramp model's larger vortex pairs yields better mixing characteristics for this model.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: AIAA PAPER 91-2264
    Format: text
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    NASA TECHNICAL REPORTS
  • 6
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    The NASA Lewis Research Center Internal Fluid Mechanics Facility (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental facility specifically designed to investigate internal fluid duct flows is described. It is built in a modular fashion so that a variety of internal flow test hardware can be installed in the facility with minimal facility reconfiguration. The facility and test hardware interfaces are discussed along with design constraints of future test hardware. The plenum flow conditioning approach is also detailed. Available instrumentation and data acquisition capabilities are discussed. The incoming flow quality was documented over the current facility operating range. The incoming flow produces well behaved turbulent boundary layers with a uniform core. For the calibration duct used, the boundary layers approached 10 percent of the duct radius. Freestream turbulence levels at the various operating conditions varied from 0.64 to 0.69 percent of the average freestream velocity.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: NASA-TM-105187 , E-6487 , NAS 1.15:105187
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    NASA TECHNICAL REPORTS
  • 7
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    An experimental trace gas investigation of fluid transport and mixing in a circular-to-rectangular transition duct (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: An ethylene trace gas technique was used to map out fluid transport and mixing within a circular-to-rectangular transition duct. Ethylene gas was injected at several points in a cross stream plane upstream of the transition duct. Ethylene concentration contours were determined at several cross stream measurement planes spaced axially within the duct. The flow involved a uniform inlet flow at a Mach number level of 0.5. Statistical analyses were used to quantitatively interpret the trace gas results. Also, trace gas data were considered along with aerodynamic and surface flow visualization results to ascertain transition duct flow phenomena. Convection of wall boundary layer fluid by vortices produced regions of high total pressure loss in the duct. The physical extent of these high loss regions is governed by turbulent diffusion.
    Keywords: INSTRUMENTATION AND PHOTOGRAPHY
    Type: AIAA PAPER 91-2370
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  • 8
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    Surface and flow field measurements in a symmetric crossing shock wave/turbulent boundary layer flow (1992)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Results of an experimental investigation of a symmetric crossing shock/turbulent boundary layer interaction are presented for a Mach number of 3.44 and deflections angles of 2, 6, 8 and 9 deg. The interaction strengths vary from weak to strong enough to cause a large region of separated flow. Measured quantities include surface static pressure and flowfield Pitot pressures. Pitot profiles in the plane of symmetry through the interaction region are shown for various deflection angles. Oil flow visualization and the results of a trace gas streamline tracking technique are also presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 92-2634
    Format: text
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    NASA TECHNICAL REPORTS
  • 9
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    Flow visualization of shock-boundary layer interaction (1982)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Two and three-dimensional shock-boundary layer interaction data were obtained from supersonic wind tunnel tests. These interactions are studied both with and without boundary layer bleed. The data verify computational fluid dynamic codes. Surface static pressure, pitot pressure, flow angularity, and bleed rates, are studied by flow visualization techniques. Surface oil flow using fluorescent dye and laser sheet using water droplets as the scattering material are used for flow visualization.
    Keywords: INSTRUMENTATION AND PHOTOGRAPHY
    Type: NASA. Langley Research Center Flow Visualization and Laser Velocimetry for Wind Tunnels; p 101-108
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  • 10
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    The effect of varying Mach number on crossing, glancing shocks/turbulent boundary-layer interactions (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Two crossing side-wall shocks interacting with a supersonic tunnel wall boundary layer have been investigated over a Mach number range of 2.5 to 4.0. The investigation included a range of equal shock strengths produced by shock generators at angles from 4.0 to 12.0 degrees. Results of flow visualization show that the interaction is unseparated at the low shock generator angles. With increasing shock strength, the flow begins to form a separated region that grows in size and moves forward and eventually the model unstarts. The wall static pressures show a symmetrical compression that merges on the centerline upstream of the inviscid shock locations and becomes more 1D downstream. The region of the 1D pressure gradient moves upstream with increasing shock strengths until it coincides with the leading edge of the shock generators at the limit before model unstart. At the limiting conditions the wall pressure gradients are primarily in the axial direction throughout.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-2157
    Format: text
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