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  • 1
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    Viscous three-dimensional analyses for nozzles for hypersonic propulsion (1990)
    In:  CASI
    Details
    Publication Date: 2013-08-29
    Description: A Navier-Stokes computer code was validated using a number of two- and three-dimensional configurations for both laminar and turbulent flows. The validation data covers a range of freestream Mach numbers from 3 to 14, including wall pressures, velocity pressure, and skin friction. Nozzle flow fields computed for a generic scramjet nozzle from Mach 3 to 20, wall pressures, wall skin friction values, heat transfer values, and overall performance are presented. In addition, three-dimensional solutions obtained for two asymmetric, single expansion ramp nozzles at a pressure ratio of 10 consists of the internal expansion region in the converging/diverging sections and the external superonic exhaust in a quiescent ambient environment. The fundamental characteristics that were captured successfully include expansion fans; Mach wave reflections; mixing layers; and nonsymmetrical, multiple inviscid cell, supersonic exhausts. Comparison with experimental data for wall pressure distributions at the center planes shows good agreement.
    Keywords: AERODYNAMICS
    Type: AGARD, Hypersonic Combined Cycle Propulsion; 18 p
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Numerical simulation of supersonic flow using a new analytical bleed boundary condition (1995)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A new analytical bleed boundary condition is used to compute flowfields for a strong oblique shock wave/boundary layer interaction with a baseline and three bleed rates at a freestream Mach number of 2.47 with an 8 deg shock generator. The computational results are compared to experimental Pitot pressure profiles and wall static pressures through the interaction region. An algebraic turbulence model is employed for the bleed and baseline cases, and a one equation model is also used for the baseline case where the boundary layer is separated.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-198368 , E-9807 , NAS 1.26:198368
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Three-dimensional compressible turbulent computations for a diffusing S-duct (1992)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The purpose of the present study was to evaluate the capability of the computational fluid dynamics computer program PARC3D to model flow in a typical diffusing subsonic S-duct, with strong secondary flows. This evaluation is needed to provide confidence in the analysis of aircraft inlets, which have similar geometries. The performance predictions include total pressure profiles, static pressures, velocity profiles, boundary layer data, and skin friction data. Flow in the S-duct is subsonic, and the boundary layers are assumed to be turbulent. The results for both H and O grid solutions, are compared with existing test data.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4392 , E-6172 , NAS 1.26:4392
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Three-dimensional compressible turbulent computations for a nondiffusing S-duct (1992)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The PARC3D code was used to compute the compressible turbulent flow within a three dimensional, nondiffusing S-duct. A frame of reference is provided for future computational fluid dynamics studies of internal flows with strong secondary flows and provides an understanding of the performance characteristics of a typical S-duct with attached flow. The predicted results, obtained with both H- and O-grids, are compared with the experimental wall pressure, static and total pressure fields, and velocity vectors. Additionally, computed boundary layer thickness, velocity profiles in wall coordinates, and skin friction values are presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4391 , E-6173 , NAS 1.26:4391
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Characterization of real gas properties for space shuttle main engine fuel turbine and performance calculations (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Real thermodynamic and transport properties of hydrogen, steam, the SSME mixture, and air are developed. The SSME mixture properties are needed for the analysis of the space shuttle main engine fuel turbine. The mixture conditions for the gases, except air, are presented graphically over a temperature range from 800 to 1200 K, and a pressure range from 1 to 500 atm. Air properties are given over a temperature range of 320 to 500 K, which are within the bounds of the thermodynamics programs used, in order to provide mixture data which is more easily checked (than H2/H2O). The real gas property variation of the SSME mixture is quantified. Polynomial expressions, needed for future computer analysis, for viscosity, Prandtl number, and thermal conductivity are given for the H2/H2O SSME fuel turbine mixture at a pressure of 305 atm over a range of temperatures from 950 to 1140 K. These conditions are representative of the SSME turbine operation. Performance calculations are presented for the space shuttle main engine (SSME) fuel turbine. The calculations use the air equivalent concept. Progress towards obtaining the capability to evaluate the performance of the SSME fuel turbine, with the H2/H2O mixture, is described.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-175066 , E-2938 , NAS 1.26:175066
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Two-dimensional viscous flow computations of hypersonic scramjet nozzle flowfields at design and off-design conditions (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The PARC2D code has been selected to analyze the flowfields of a representative hypersonic scramjet nozzle over a range of flight conditions from Mach 3 to 20. The flowfields, wall pressures, wall skin friction values, heat transfer values and overall nozzle performance are presented.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: AIAA PAPER 88-3280
    Format: text
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    NASA TECHNICAL REPORTS
  • 7
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    Hypersonic turbulent wall boundary layer computations (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The Baldwin-Lomax (1978) algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer-layer eddy-viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith (1974) model for a flat plate at hypersonic speed, the new values of the coefficients were obtained. The results show that the values of C(cp) and C(kleb) are functions of both Mach number and wall temperature ratio. The C(cp) and C(kleb) variations with Mach number and wall temperature were used for the calculations of both a 4-deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin-layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2829
    Format: text
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    NASA TECHNICAL REPORTS
  • 8
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    Three-dimensional viscous flow computations of a circular jet in subsonic and supersonic cross flow (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Three-dimensional viscous flow computations are presented for 90 deg injection angle jets in subsonic and supersonic cross flow. Comparisons with experimental data include jet centerline and vortex trajectories for the subsonic cross flow, and surface pressure measurement for the supersonic crossflow case. The vortices induced in the jet/freestream interaction are computed and illustrated. The vortices persist in subsonic flow and die out quickly in supersonic flow. The structure of the shocks in the unconfined supersonic flow is illustrated.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-CR-182153 , E-4223 , NAS 1.26:182153
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    Two-dimensional viscous flow computations of hypersonic scramjet nozzle flowfields at design and off-design conditions (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The PARC2D code has been selected to analyze the flowfields of a representative hypersonic scramjet nozzle over a range of flight conditions from Mach 3 to 20. The flowfields, wall pressures, wall skin friction values, heat transfer values and overall nozzle performance are presented.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-182150 , E-4190 , NAS 1.26:182150
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    Hypersonic turbulent wall boundary layer computations (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The Baldwin-Lomax algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer layer eddy viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith model for a flat plate at hypersonic speed, the new values of the coefficient were obtained. The results show that the values of C sub cp and C sub kleb are functions of both Mach number and wall temperature ratio. The C sub cp and C sub kleb variations with Mach number and wall temperature were used for the calculations of both a 4 deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-CR-182147 , E-4172 , NAS 1.26:182147 , AIAA PAPER 88-2829
    Format: application/pdf
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