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  • 1
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    High-speed code validation (1987)
    In:  CASI
    Details
    Publication Date: 2006-06-13
    Description: The topics are presented in viewgraph form and include the following: NFL body experiment; high-speed validation problems; 3-D Euler/Navier-Stokes inlet code; two-strut inlet configuration; pressure contours in two longitudinal planes; sidewall pressure distribution; pressure distribution on strut inner surface; inlet/forebody tests in 60 inch helium tunnel; pressure distributions on elliptical missile; code validations; small scale test apparatus; CARS nonintrusive measurements; optimized cone-derived waverider study; etc.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA. Ames Research Center, NASA CFD Validation Workshop; p 210-243
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    A numerical model for supersonic reacting mixing layers (1987)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: A current research effort is underway at the NASA Langley Research Center to achieve a detailed understanding of important phenomena present when a supersonic flow undergoes a chemical reaction. A computer program has been developed to study the details of such flows. The program has been constructed to consider the multicomponent diffusion and convection of important species, the finite-rate reaction of these species, and the resulting interaction between the fluid mechanics and chemistry. Code results from the analysis of a spatially developing and reacting mixing layer are presented, and conclusions are drawn regarding the structure of the evolving layer and its associated flame.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: Computer Methods in Applied Mechanics and Engineering (ISSN 0045-7825); 64; 39-60
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    NASA TECHNICAL REPORTS
  • 3
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    Effects of test facility contaminants on supersonic hydrogen-air diffusion flames (1986)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: Results are presented from an analytical assessment of the effects of NO and H2O in air on the ignition and reaction of hydrogen. Three flow models of the burning process were used, each including a 32 reaction chemistry mechanism. These models were a plug flow reactor to determine ignition and reaction times; a well-stirred reactor to assess flame stability or blow-off; and a parabolic Navier-Stokes code to examine the diffusion flame combustion. The diffusion flame was computed for a cold supersonic H2 jet coaxial with a supersonic air stream at a static temperature of 1100 K and a pressure of 1 atmosphere.
    Keywords: PROPELLANTS AND FUELS
    Type: Johns Hopkins Univ., The 23rd JANNAF Combustion Meeting, Volume 1; p 377-390
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Initial conditions for parabolic calculation of reacting flows (1986)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: Journal of Propulsion and Power (ISSN 0748-4658); 2; 235-240
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  • 5
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    Numerical predictions of a rearward-facing-step flow in a supersonic combustor (1989)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: Journal of Propulsion and Power (ISSN 0748-4658); 5; 158-164
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  • 6
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    Flow establishment in a generic scramjet combustor (1990)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The establishment of a quasi-steady flow in a generic scramjet combustor was studied for the case of a time varying inflow to the combustor. Such transient flow is characteristic of the reflected shock tunnel and expansion tube test facilities. Several numerical simulations of hypervelocity flow through a straight duct combustor with either a side wall step fuel injector or a centrally located strut injector are presented. Comparisons were made between impulsively started but otherwise constant flow conditions (typical of the expansion tube or tailored operations of the reflected shock tunnel) and the relaxing flow produced by the 'undertailored' operations of the reflected shock tunnel. Generally the inviscid flow features, such as the shock pattern and pressure distribution, were unaffected by the time varying inlet conditions and approached steady state in approx. the times indicated by experimental correlations. However, viscous features, such as heat transfer and skin friction, were altered by the relaxing inlet flow conditions.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-CR-187467 , ICASE-90-77 , NAS 1.26:187467 , AD-A229670
    Format: application/pdf
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  • 7
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    Three-dimensional computation of mixing of transverse injector in a ducted supersonic airstream (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Numerical solutions of the three-dimensional mass-averaged elliptic Navier-Stokes equations, including species transport, are obtained for nonreacting, turbulent, mixing flow fields for the case of transverse sonic injection of a secondary gas into a supersonic airstream through a circular orifice injector. Results are presented for flow through a constant area duct and through a duct with a rearward-facing step upstream of the injector. The equations are numerically integrated using MacCormack's explicit method and turbulence is included using the Baldwin-Lomax algebraic eddy viscosity model. In the species transport and energy equations, diffusion coefficients based on Fick's law and an assumption of unit Lewis number are applied. The computations were performed on a CDC-VPS-32 (extended version of Cyber-205) using a grid consisting of approximately 200,000 points. The computed results are compared with experimentally observed penetration and spreading boundaries for an injected gas at two dynamic pressure ratios. Three-dimensional flow field structures are dipicted in terms of static pressure, mass fractions of species and velocity vectors.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: AIAA PAPER 86-1423
    Format: text
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    NASA TECHNICAL REPORTS
  • 8
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    A detailed numerical model of a supersonic reacting mixing layer (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A current research effort is underway at the NASA Langley Research Center to achieve a detailed understanding of important phenomena present when a supersonic flow undergoes chemical reaction. A computer program has been developed to study the details of such flows. The program has been constructed to consider the multicomponent diffusion and convection of important species, the finite-rate reaction of these species, and the resulting interaction between the fluid mechanics and chemistry. Code results from the analysis of a spatially developing and reacting mixing layer are presented, and conclusions are drawn regarding the structure of the evolving layer and its associated flame.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: AIAA PAPER 86-1427
    Format: text
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    NASA TECHNICAL REPORTS
  • 9
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    A comparative study of scramjet injection strategies for high Mach numbers flows (1992)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A simple method for predicting the axial distribution of supersonic combustor thrust potential is described. A complementary technique for illustrating the spatial evolution and distribution of thrust potential and loss mechanisms in reacting flows is developed. Wall jet cases and swept ramp injector cases for Mach 17 and Mach 13.5 flight enthalpy inflow conditions are numerically modeled and analyzed using these techniques. The visualization of thrust potential in the combustor for the various cases examined provides a unique tool for increasing understanding of supersonic combustor performance potential.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: AIAA PAPER 92-3287
    Format: text
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    NASA TECHNICAL REPORTS
  • 10
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    Premixed shock-induced combustion studies in the hypulse facility (1992)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Engine performance calculations strongly suggest that the oblique detonation-wave engine can outperform the diffusive-burning scramjet engine in the high hypervelocity flight regime; i.e., at flight Mach numbers exceeding about 14. Research is described that examines and characterizes the behavior of oblique detonation waves (ODW). The effort is both computational and experimental, the latter utilizing the NASA Hypulse expansion tube. The experimental work is hampered by preignition of the hydrogen-oxygen-helium mixtures in either the intermediate or acceleration sections of the facility. The experiments are described and possible causes of this preignition are discussed. Initial results obtained from a parallel computation effort indicate that temporally steady ODW are achievable in the Hypulse facility.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: AIAA PAPER 92-3427
    Format: text
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