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  • 1
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    Full potential solutions of three-dimensional supersonic flows (1985)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 22; 881-887
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Streamtube analysis of a hydrogen-burning scramjet exhaust and simulation technique (1977)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: A scramjet/airframe integration program and a technique for simulating thermally perfect scramjet exhaust flows (freon/argon gas blends) is studied to extend the technique to more complicated flows approaching the actual exhaust flow in complexity. The state of the flow and the accuracy of the substitute gas simulation are analyzed in the case of a shock discontinuity present. Findings are: scramjet exhaust flow is essentially frozen throughout the expansion at Mach 6 and Mach 8; flow behind moderate shocks remains frozen; the technique can accurately track static distributions in scramjet exhaust flows (shocked or unshocked).
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: Journal of Aircraft; 14; Sept
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    Calculated heating on the afterbody nozzle of a hypersonic aircraft concept (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The present analysis was carried out to estimate the heating levels of the external nozzle of a scramjet/airframe-integrated research aircraft. A parametric examination of the effects of Mach number, reference length, and wall temperature showed that the heating rate distributions are independent of reference length and wall temperature. The initial heating rates obtained for a Mach 6 flight are in the (3 to 8) x 10 to the 5th power W/sq m range. Underlying the entire study is the question of nozzle boundary layer formation and growth, as well as the question of the reference length value that should be used in the computations. It is shown that the reference length is not the dominant factor setting the heating levels; an attempt to bound the actual length was made. A more detailed calculation of the rates requires further work to gain a better understanding of the combustor exit boundary layer.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 14; Dec. 197
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Two-dimensional viscous simulation of inlet/diffuser flows with terminal shocks (1985)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: Journal of Propulsion and Power (ISSN 0748-4658); 1; 103-108
    Format: text
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    NASA TECHNICAL REPORTS
  • 5
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    Application of a full potential method for computation of three-dimensional supersonic flows (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A nonlinear aerodynamic analysis technique based on the full potential equation in conservative form has been modified to permit treatment of supersonic flows with embedded subsonic regions (typically near the fuselage-canopy juncture and the wing leading edge). Solution procedures for the equations do not require any specific form of geometry or physical grid system. This results in the capability to analyze easily very complex geometries provided the posed problem lies within the isentropic restrictions of the full potential theory. Characteristic signal propagation theory is used to monitor the type dependent flow and a conservative switching scheme is employed to transition from the supersonic marching algorithm to a subsonic relaxation procedure and vice versa. An implicit approximate factorization scheme is used to solve the finite-difference equations. These modifications now permit analysis of fully three-dimensional flowfields including the interference effects due to lifting surface wakes. Improved grid generation capability allows analysis of complete complex aircraft geometries (fuselage, wing, tail, wing wake, and tail wake). Results are presented showing very good correlations with experimental surface pressure data and aerodynamic force data at both design and off-design operating points. Configurations examined include several waverider concepts, an arrow wing-body with wake, an advanced tactical fighter concept, and a fighter forebody-canard configuration.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0139
    Format: text
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  • 6
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    The use of interpractive graphic displays for interpretation of surface design parameters (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An interactive computer graphics technique known as the Graphic Display Data method has been developed to provide a convenient means for rapidly interpreting large amounts of surface design data. The display technique should prove valuable in such disciplines as aerodynamic analysis, structural analysis, and experimental data analysis. To demonstrate the system's features, an example is presented of the Graphic Data Display method used as an interpretive tool for radiation equilibrium temperature distributions over the surface of an aerodynamic vehicle. Color graphic displays were also examined as a logical extension of the technique to improve its clarity and to allow the presentation of greater detail in a single display.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA-TM-81963 , L-14112
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Numerical simulation of flow through inlets/diffusers with terminal shocks (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Two-dimensional Euler and Navier-Stokes solutions of the flow through three inlet/diffuser configurations with terminal shock systems are reported. Calculations without bleed indicate that the terminal shock location is very sensitive to the outflow back pressure. For cases where there are no available experimental results, it becomes difficult to estimate the back pressure that will result in a terminal shock. Estimates based on quasi-one-dimensional analysis are not found adequate for complex two-dimensional flows. It is found that since the flow downstream of the terminal shock is subsonic, and what happens at the outflow boundary affects the flow inside the inlet, enough of the subsonic diffuser must be modeled to accurately predict the terminal shock region. The diffuser portion should be fairly long with the outflow boundary occurring in a region of more or less uniform flow to be able to prescribe a uniform back pressure. The third configuration studied was investigated with and without incorporating bleed in the code. It is found that the use of bleed stabilizes the shock location and allows solutions which without bleed result in unstarting of the inlet. Comparisons are made with available experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1362
    Format: text
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  • 8
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    Flow field visualization about external axial corners (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental investigation was conducted to visualize the flow field about external axial corners. The investigation was initiated to provide answers to questions about the inviscid flow pattern for continuing numerical investigations. Symmetrical and asymmetrical corner models were tested at a Reynolds number per meter of 60,700,000. Oil-flow and vapor-screen photographs were taken for both models at angle of attack and yaw. The paper presents the results of the investigation in the form of oil-flow photographs and the surrounding shock wave location obtained from the vapor screens.
    Keywords: AERODYNAMICS
    Type: NASA-TM-78682
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    Scramjet exhaust simulation technique for hypersonic aircraft nozzle design and aerodynamic tests (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Current design philosophy for scramjet-powered hypersonic aircraft results in configurations with the entire lower fuselage surface utilized as part of the propulsion system. The lower aft-end of the vehicle acts as a high expansion ratio nozzle. Not only must the external nozzle be designed to extract the maximum possible thrust force from the high energy flow at the combustor exit, but the forces produced by the nozzle must be aligned such that they do not unduly affect aerodynamic balance. The strong coupling between the propulsion system and aerodynamics of the aircraft makes imperative at least a partial simulation of the inlet, exhaust, and external flows of the hydrogen-burning scramjet in conventional facilities for both nozzle formulation and aerodynamic-force data acquisition. Aerodynamic testing methods offer no contemporary approach for such vehicle design requirements. NASA-Langley has pursued an extensive scramjet/airframe integration R&D program for several years and has recently developed a promising technique for simulation of the scramjet exhaust flow for hypersonic aircraft. Current results of the research program to develop a scramjet flow simulation technique through the use of substitute gas blends are described in this paper.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 77-82 , Aerospace Sciences Meeting; Jan 24, 1977 - Jan 26, 1977; Los Angeles, CA
    Format: text
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  • 10
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    Thermodynamic properties of gaseous fluorocarbons and isentropic equilibrium expansions of two binary mixtures of fluorocarbons and argon (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Equations and computer code are given for the thermodynamic properties of gaseous fluorocarbons in chemical equilibrium. In addition, isentropic equilibrium expansions of two binary mixtures of fluorocarbons and argon are included. The computer code calculates the equilibrium thermodynamic properties and, in some cases, the transport properties for the following fluorocarbons: CCl2F, CCl2F2, CBrF3, CF4, CHCl2F, CHF3, CCL2F-CCl2F, CCLF2-CClF2, CF3-CF3, and C4F8. Equilibrium thermodynamic properties are tabulated for six of the fluorocarbons(CCl3F, CCL2F2, CBrF3, CF4, CF3-CF3, and C4F8) and pressure-enthalpy diagrams are presented for CBrF3.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TN-D-8405 , L-11274
    Format: application/pdf
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