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  • 1
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    Applied aspects of laminar-flow technology (1990)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: An account is given of the development status and performance levels achieved with natural laminar flow (NLF), extended wing chord laminar flow control (LFC), and hybrid laminar flow control (HLFC) concepts combining NLF and partial-chord LFC in the leading-edge region. Attention is given to NLF wing structure construction methods capable of achieving the requisite surface-irregularity tolerances, LFC through wing surface suction slots or perforated skins, and the deleterious effects of insects, ice crystals, and noise disturbance inputs on the ability of NLF, LFC, and HLFC wings to maintain effective laminar flow operation.
    Keywords: AERODYNAMICS
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  • 2
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    A programmable heater control circuit for spacecraft (1994)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Spacecraft thermal control is accomplished for many components through use of multilayer insulation systems, electrical heaters, and radiator systems. The heaters are commanded to maintain component temperatures within design specifications. The programmable heater control circuit (PHCC) was designed to obtain an effective and efficient means of spacecraft thermal control. The hybrid circuit provides use of control instrumentation as temperature data, available to the spacecraft central data system, reprogramming capability of the local microprocessor during the spacecraft's mission, and the elimination of significant spacecraft wiring. The hybrid integrated circuit has a temperature sensing and conditioning circuit, a microprocessor, and a heater power and control circuit. The device is miniature and housed in a volume which allows physical integration with the component to be controlled. Applications might include alternate battery-powered logic-circuit configurations. A prototype unit with appropriate physical and functional interfaces was procured for testing. The physical functionality and the feasibility of fabrication of the hybrid integrated circuit were successfully verified. The remaining work to develop a flight-qualified device includes fabrication and testing of a Mil-certified part. An option for completing the PHCC flight qualification testing is to enter into a joint venture with industry.
    Keywords: AERODYNAMICS
    Type: NASA-TM-108459 , NAS 1.15:108459
    Format: application/pdf
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  • 3
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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
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  • 4
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    High Alpha Technology Program (HATP) ground test to flight comparisons (1994)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: This status paper reviews the experimental ground test program of the High Alpha Technology Program (HATP). The reasons for conducting this ground test program had their origins during the 1970's when several difficulties were experienced during the development programs of both the F-18 and F-16. A careful assessment of ground test to flight correlations appeared to be important for reestablishing a high degree of confidence in our ground test methodology. The current paper will then focus on one aspect of the HATP program that is intended to improve the correlation between ground test and flight, high-alpha gritting. The importance of this work arises from the sensitivity of configurations with smooth-sided forebodies to Reynolds number. After giving examples of the effects of Reynolds number, the paper will highlight efforts at forebody gritting. Finally, the paper will conclude by summarizing the charter of the HATP Experimental Aerodynamics Working Group and future experimental testing plans.
    Keywords: AERODYNAMICS
    Type: NASA. Dryden Flight Research Center, Fourth High Alpha Conference, Volume 1; 25 p
    Format: application/pdf
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  • 5
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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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  • 6
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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
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  • 7
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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
    Format: text
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  • 8
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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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  • 9
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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
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  • 10
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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
    Format: text
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