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  • AERODYNAMICS  (179)
  • 1990-1994  (179)
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  • 1992  (88)
  • 1991  (91)
  • 1934
  • 1
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    Aeroassist flight experiment heating-rate sensitivity study (1991)
    In:  Other Sources
    Details
    Publication Date: 2011-08-24
    Keywords: AERODYNAMICS
    Type: Journal of Thermophysics and Heat Transfer (ISSN 0887-8722); 5; 456-462
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Aerodynamic heating on AFE due to nonequilibrium flow with variable entropy at boundary layer edge (1991)
    In:  CASI
    Details
    Publication Date: 2013-08-29
    Description: A method of predicting the aerobrake aerothermodynamic environment on the NASA Aeroassist Flight Experiment (AFE) vehicle is described. Results of a three dimensional inviscid nonequilibrium solution are used as input to an axisymmetric nonequilibrium boundary layer program to predict AFE convective heating rates. Inviscid flow field properties are obtained from the Euler option of the Viscous Reacting Flow (VRFLO) code at the boundary layer edge. Heating rates on the AFE surface are generated with the Boundary Layer Integral Matrix Procedure (BLIMP) code for a partially catalytic surface composed of Reusable Surface Insulation (RSI) times. The 1864 kg AFE will fly an aerobraking trajectory, simulating return from geosynchronous Earth orbit, with a 75 km perigee and a 10 km/sec entry velocity. Results of this analysis will provide principal investigators and thermal analysts with aeroheating environments to perform experiment and thermal protection system design.
    Keywords: AERODYNAMICS
    Type: ESA, Aerothermodynamics for Space Vehicles; p 371-376
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  • 3
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    The NASA Langley laminar-flow-control experiment on a swept, supercritical airfoil: Suction coefficient analysis (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A swept supercritical wing incorporating laminar flow control at transonic flow conditions was designed and tested. The definition of an experimental suction coefficient and a derivation of the compressible and incompressible formulas for the computation of the coefficient from measurable quantities is presented. The suction flow coefficient in the highest velocity nozzles is shown to be overpredicted by as much as 12 percent through the use of an incompressible formula. However, the overprediction on the computed value of suction drag when some of the suction nozzles were operating in the compressible flow regime is evaluated and found to be at most 6 percent at design conditions.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4267 , L-16774 , NAS 1.15:4267
    Format: application/pdf
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  • 4
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    Evaluation of panel code predictions with experimental results of inlet performance for a 17-inch ducted prop/fan simulator operating at Mach 0.2 (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: An axisymmetric panel code was used to evaluate a series of ducted propeller inlets. The inlets were tested in the Lewis 9 by 15 Foot Low Speed Wind Tunnel. Three basic inlets having ratios of shroud length to propeller diameter of 0.2, 0.4, and 0.5 were tested with the Pratt and Whitney ducted prop/fan simulator. A fourth hybrid inlet consisting of the shroud from the shortest basic inlet coupled with the spinner from the largest basic inlet was also tested. This later configuration represented the shortest overall inlet. The simulator duct diameter at the propeller face was 17.25 inches. The short and long spinners provided hub-to-tip retios of 0.44 at the propeller face. The four inlets were tested at a nominal free stream Mach number of 0.2 and at angles of attack from 0 degrees to 35 degrees. The panel code method incorporated a simple two-part separation model which yielded conservative estimates of inlet separation.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-3354
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  • 5
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    Unsteady separation in sharp fin-induced shock wave/turbulent boundary layer interaction at Mach 5 (1992)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Fluctuating wall-pressure measurements are made in shock-wave/turbulent-boundary-layer interactions generated by sharp/unswept fins at angles of attack of 16, 18, 20, 22, 24, 26, and 28 degrees at Mach 5. The experiment was conducted under approximately adiabatic wall temperature conditions. The mean and rms pressure distributions can be collapsed in conical coordinates. The wall-pressure signal near separation is intermittent for all angles of attack (16-28 deg) and is qualitatively similar to that measured in unswept flows. However, the shock frequencies are higher - about 5 kHz compared to 0.5-1 kHz. Over the range of sweepbacks examined, from 25-55 deg, the spectral content of the fluctuating pressures does not change. Thus, the increase in separation-shock frequency from 1 to 5 kHz occurs at lower interaction sweepback and is not a continuous process with increasing sweepback. Power spectra at the position of maximum rms in the intermittent region for interactions in different incoming boundary layers have the same center frequency. The maximum rms in the intermittent region correlates with interaction sweepback, not with overall inviscid pressure rise.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 92-0748
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  • 6
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    The hypersonic shock wave-turbulent boundary layer interaction generated by a sharp fin at Mach 8.2 (1992)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A combined experimental and numerical study has been conducted on the hypersonic shock-wave turbulent-boundary layer interaction at Mach 8.2 generated by a single fin of angles alpha = 10 and 15 deg. Three models are considered: (1) the 3D compressible Reynolds-averaged Navier-Stokes (RANS) equations using the k-epsilon turbulence model, (2) the 3D RANS using the Rodi turbulence model, and (3) the conical RANS using the Baldwin-Lomax algebraic turbulence model. The computations are compared with various experimental data. The computations using models (1) and (2) show quantitatively very similar results and very good agreement with experimental data for surface pressure and skin friction. Comparison with boundary layer profiles of pitot pressure and yaw angle are also generally good, but the peak surface heat transfer is overestimated by up to 48 percent. The effect of the laminar boundary layer on the fin is restricted to the immediate vicinity of the fin surface. Conical calculations using model (3) show substantially poorer agreement with experiment.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 92-0747
    Format: text
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  • 7
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    Crossing shock wave-turbulent boundary layer interactions (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Three-dimensional interactions between crossing shock waves generated by symmetric sharp fins and a turbulent boundary layer on a flat plate are investigated experimentally and theoretically at Mach number 2.95 and freestream unit Reynolds number 1.96 x 10 to the 7th/ft. The incoming boundary layer has a thickness of 4 mm at the location of the fin leading edges. A comparison of experimental and computational results for two sets of fin angles (11 x 11 and 9 x 9 deg) shows general agreement with regard to surface pressure measurements and surface streamline patterns. The principal feature of the streamline structure is a collision of counterrotating vortical structures emanating from near the fin leading edges and meeting at the geometric centerline of the interaction.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0649
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  • 8
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    Evaluation of panel code predictions with experimental results of inlet performance for a 17-inch ducted prop/fab simulator operating at Mach 0.2 (1991)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An axisymmetric panel code was used to evaluate a series of ducted propeller inlets. The inlets were tested in the Lewis 9 by 15 Foot Low Speed Wind Tunnel. Three basic inlets having ratios of shroud length to propeller diameter of 0.2, 0.4, and 0.5 were tested with the Pratt and Whitney ducted prop/fan simulator. A fourth hybrid inlet consisting of the shroud from the shortest basic inlet coupled with the spinner from the largest basic inlet was also tested. This later configuration represented the shortest overall inlet. The simulator duct diameter at the propeller face was 17.25 inches. The short and long spinners provided hub-to-tip ratios of 0.44 at the propeller face. The four inlets were tested at a nominal free stream Mach number of 0.2 and at angles of attack from 0 degrees to 35 degrees. The panel code method incorporated a simple two-part separation model which yielded conservative estimates of inlet separation.
    Keywords: AERODYNAMICS
    Type: NASA-TM-104428 , E-6261 , NAS 1.15:104428 , AIAA PAPER 91-3354 , Joint Propulsion Conference; Jun 24, 1991 - Jun 27, 1991; Sacramento, CA; United States
    Format: application/pdf
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  • 9
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    Longitudinal vortex control - Techniques and applications (The 32nd Lanchester Lecture) (1992)
    In:  Other Sources
    Details
    Publication Date: 2011-08-24
    Description: A summary is presented of vortex control applications and current techniques for the control of longitudinal vortices produced by bodies, leading edges, tips and intersections. Vortex control has up till now been performed by many approaches in an empirical fashion, assisted by the essentially inviscid nature of much of longitudinal vortex behavior. Attention is given to Reynolds number sensitivities, vortex breakdown and interactions, vortex control on highly swept wings, and vortex control in juncture flows.
    Keywords: AERODYNAMICS
    Type: Aeronautical Journal (ISSN 0001-9240); 96; 958; p. 293-312.
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  • 10
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    Supersonic laminar flow control (1992)
    In:  Other Sources
    Details
    Publication Date: 2011-08-24
    Description: A development status evaluation is presented for the theoretical understanding and design conceptualization of boundary layer control (BLC) systems applicable to supersonic transports, such as the currently envisioned NASA High Speed Civil Transport. By reducing fuel burned, supersonic BLC techniques could expand ranges to Pacific-crossing scales, while lowering sonic boom effects and upper-atmosphere pollution and even reducing skin friction temperature. The critical consideration for supersonic BLC is the presence of wave effects.
    Keywords: AERODYNAMICS
    Type: In: Natural laminar flow and laminar flow control (A93-41776 17-02); p. 233-245.
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