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  • Cell & Developmental Biology  (1,170)
  • Inorganic Chemistry  (773)
  • AERODYNAMICS  (684)
  • ddc:330
  • 1990-1994  (2,629)
  • 1950-1954
  • 1991  (2,629)
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Keywords
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  • 1990-1994  (2,629)
  • 1950-1954
Year
  • 101
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    Studies of shock/shock interaction on smooth and transpiration-cooled hemispherical nosetips in hypersonic flow (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Experimental studies are conducted to examine the utilization of transpiration cooling to reduce the peak-heating loads in areas of shock/shock interaction. Smooth and transpiration-cooled nosetip models, 12 inches in diameter, were employed in these studies, which focused on defining the pressure distributions and heat transfer in type III and IV interaction areas. Transpiration cooling was determined to significantly increase the size of the shock layer and to move the peak-heating point around the body. A transpiration-cooling rate of more than 30 percent of the freestream maximum flux did not lower the peak-heating level more than 10 percent, but the integrated heating loads were reduced.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1765
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  • 102
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    Stability of a compressible swirling jet (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Temporal linear stability of a compressible swirling axisymmetric jet is considered. It is found that with the addition of a modest amount of swirl, instability growth rates are substantially increased. Additionally, rotating jets are found to be highly unstable for disturbances with high aximuthal wave numbers. Such disturbances are absent for the case of non-swirling jets. Most importantly, it is found that the stabilizing influence of increasing Mach number is diminished with the introduction of swirl to the jet flow.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1770
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  • 103
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    3-D Navier-Stokes analysis of crossing, glancing shocks/turbulent boundary layer interactions (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Three dimensional viscous flow analysis is performed for a configuration where two crossing and glancing shocks interact with a turbulent boundary layer. A time marching 3-D full Navier-Stokes code, called PARC3D, is used to compute the flow field, and the solution is compared to the experimental data obtained at the NASA Lewis Research Center's 1 x 1 ft supersonic wind tunnel facility. The study is carried out as part of the continuing code assessment program in support of the generic hypersonic research at NASA Lewis. Detailed comparisons of static pressure fields and oil flow patterns are made with the corresponding solution on the wall containing the shock/boundary layer interaction in an effort to validate the code for hypersonic inlet applications.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1758
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  • 104
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    Pressure flux-split viscous solutions for subsonic diffusers (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A method for predicting flow in diffusers with inlet swirl has been developed. Solutions of the pressure flux-split Reduced Navier Stokes (RNS) equations are obtained for flows in axisymmetric diffusers. Viscous-Inviscid interactions and flow fields with toroidal recirculation regions are efficiently captured. The computational model is verified by comparision with experimental data and other computations. The extreme sensitivity to grid, turbulent closure model and inlet profiles is discussed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1745
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  • 105
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    Evaluation of a three-dimensional Navier-Stokes solver for transonic flow over a low-aspect-ratio wing mounted in a wind tunnel (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A three dimensional Navier-Stokes solver is evaluated for transonic flow over a thin, swept, low-aspect ratio wing. The computational study was undertaken in support of a wind tunnel experimental program. The computational results are compared to experimental surface pressure data obtained in a cryogenic wind tunnel with an adaptive wall test section. The results show favorable agreement over a wide range of conditions, further the numerical results provide additional data of the complex three-dimensional flow field. Differences in the predictions and experiment suggest a need to conduct further experiments to evaluate the adaptive wall testing technique, and to model the tunnel sidewall boundary layer in the computations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1725
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  • 106
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    An engineering assessment of gas/surface interactions in free-molecular aerodynamics (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Data available on the gas/surface interaction (GSI) phenomenon, particularly in the rarefied to near-free-molecular flow regime are collected, categorized, and analyzed with a purpose of developing a GSI model that could be used as a guide to spacecraft designers. The study shows that there are not enough useful data for building a high-confidence GSI model. However, sufficient results are obtained to suggest that continued reliance on the diffuse GSI model is inappropriate, particularly in the near-free-molecular regime.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1747
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  • 107
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    On the quasi-conical flowfield structure of the swept shock wave-turbulent boundary layer interaction (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The swept oblique shock-wave/turbulent-boundary-layer interaction generated by a 20-deg sharp fin at Mach 4 and Reynolds number 21,000 is investigated via a series of computations using both conical and three-dimensional Reynolds-averaged Navier-Stokes equations with turbulence incorporated through the algebraic turbulent eddy viscosity model of Baldwin-Lomax. Results are compared with known experimental data, and it is concluded that the computed three-dimensional flowfield is quasi-conical (in agreement with the experimental data), the computed three-dimensional and conical surface pressure and surface flow direction are in good agreement with the experiment, and the three-dimensional and conical flows significantly underpredict the peak experimental skin friction. It is pointed out that most of the features of the conical flowfield model in the experiment are observed in the conical computation which also describes the complete conical streamline pattern not included in the model of the experiment.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1759
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  • 108
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    Aerodynamics of powered missile separation from a wing (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A 3D dynamic 'chimera' algorithm that solves the thin-layer Navier-Stokes equations over multiple moving bodies was modified to numerically simulate the aerodynamics, missile dynamics, and missile plume of a finless missile separating from a wing in transonic flow. A powered missile separation case was considered to examine the influence of the missile and plume on the wing. The wing and missile is at a two degree angle of attack. The computational results show the details of the flow field.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1663
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  • 109
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    Numerical study of juncture flows (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The present paper describes a computational study of laminar/turbulent and subsonic/supersonic horseshoe vortex systems generated by a cylindrical protuberance mounted on a flat plate. Various vortex structures are predicted and discussed. Low-speed laminar juncture flows are computed to determine the Reynolds number effect with the same incoming boundary-layer thickness. For a low subsonic laminar flow, the number of vortex arrays increases with the Reynolds number, in agreement with both experimental and numerical observations. Qualitative comparisons are made along with the computations, experimental observations, and analytical work. For incompressible flow, the relationships among pressure extrema, vorticity, and singular points in flow structure are discussed. A parametric study of the effect of the free-stream Mach number on the flow structure for laminar flow is conducted. The juncture flow when the incoming flow is turbulent and supersonic is computed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1660
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  • 110
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    Assessment of meanflow solutions for instability analysis of transitioning flows (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Meanflow solutions of 3D supersonic flow past a cone at incidence and a swept leading edge wing have been obtained by thre methods, viz., boundary-layer, parabolized Navier-Stokes, and thin shear-layer Navier-Stokes solvers. The smoothness and accuracy of the solution profiles are compared with a view to applying the meanflow solution to boundary-layer stability analysis.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1638
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  • 111
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    Trim angle measurements in free-flight facilities (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The aerodynamic cofficients and trim angle for an aerobrake at Mach 9.2 and 11.8 were found using a combination of experiment and computation. Free-flight tests were performed at NASA Ames Research Center's Hypervelocity Free-Flight Aerodynamic Facility, and the forebody pressure distribution was calculated using a three-dimensional Navier-Stokes code with an effective specific heat ratio. Using the computed drag, lift, and moments to prescribe the number of terms in the aerodynamic coefficient expansions and to specify the values of the higher order terms, the experimental aerodynamic coefficients and trim angle were found using a six-degree-of-freedom, weighted, least-squares analysis. The experimental and computed aerodynamic coefficients and trim angles are in good agreement. The trim angle obtained from the free-flight tests, 14.7 deg, differs from the design trim angle, 17 deg, and from the Langley wind tunnel results, 12 deg in air and 17 deg in CF4. These differences are attributable to real-gas effects.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1632
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  • 112
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    Navier-Stokes solutions for an oscillating double-delta wing (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An upwind-biased implicit scheme is used to investigate steady-state and unsteady Navier-Stokes solutions of the vortical flow over a double-delta wing configuration. The governing equations are solved numerically with a fully upwind, implicit, iterative, and factorized numerical scheme. Steady-state solutions for fixed angles of attack and unsteady solutions for a sinusoidal oscillatory motion are obtained. The steady-state solutions on the baseline grid are in agreement with the experiment, and grid refinements show some improvements of the predictions. The higher-order accuracy of the present scheme yields equivalent solutions on smaller grid densities compared to solutions obtained with a second-order accurate method on larger grids. As the angle of attack increases, the grid resolution requirements for adequate resolution of the leeward-side vortical flowfield become very severe. The unsteady solutions are in general agreement with the measurements and show a qualitative correlation with the experiment.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1624
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  • 113
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    Study of cluster formation and its effects on Rayleigh and Raman scattering measurements in a Mach 6 wind tunnel (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Using a frequency-doubled Nd-YAG pulsed laser and a single-intensified CCD camera, Rayleigh scattering measurements have been performed to study the cluster formation in a Mach 6 wind tunnel at NASA Langley Research Center. These studies were conducted both in the free stream and in a model flow field for various flow conditions to gain an understanding of the dependence of the Rayleigh scattering (by clusters) on the local pressures and temperatures in the facility. Using the same laser system, simultaneous measurements of the local temperature have also been performed using the rotational Raman scattering of molecular nitrogen and determined the densities of molecular oxygen and nitrogen by using the vibrational Raman scattering from these species. Quantitative results are presented in detail with emphasis on the applicability of the Rayleigh scattering for obtaining quantitative measurements of molecular densities both in the free stream and in the model flow field.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1496
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  • 114
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    Reattachment studies of an oscillating airfoil dynamic stall flow field (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The reattaching flow over an oscillating airfoil executing large amplitude sinusoidal motion around a mean angle of attack of 10 degrees has been studied using the techniques of stroboscopic schlieren, two component laser Doppler velocimetry and point diffraction interferometry, for a free stream Mach number of 0.3 and a reduced frequency of 0.05. The results show that the dynamically stalled flow reattaches in a process that begins when the airfoil is very close to the static stall angle on its downward stroke and progresses over the airfoil through a large range of angles of attack as the airfoil decreases to about 6 degrees. The airfoil suction peak shows a dramatic rise as the static stall angle is approached and the velocity profiles develop such that the flow near the surface is accelerated. The process completes through the disappearance of a separation bubble that forms over the airfoil.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-3225
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  • 115
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    Reynolds number effects on supersonic asymmetrical flows over a cone at high angle of attack (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The supersonic viscous flow over a 5-degree half-angle cone at an angle of attack of four times the cone half-angle is studied computationally using both the conical and the three-dimensional Navier-Stokes equations. The numerical solutions were obtained with an implicit, upwind-biased algorithm. Asymmetrical flowfields of the absolute-instability type are found using the conical-flow equations which agree with published results. However, the absolute instabilities of the originally symmetric flow found with the conical equations do not occur in the three-dimensional simulations, although spurious asymmetric three-dimensional flows for symmetric bodies arise if the grid resolution is insufficient in the nose region. The asymmetric flows computed with the three-dimensional equations are convective instabilities and are possible if the local Reynolds number exceeds a critical value and a fixed geometric asymmetry is imposed. A continuous range of asymmetries can be developed, depending on the size of the disturbance and the Reynolds number. As the Reynolds number is increased, the asymmetries demonstrate a bistable behavior at levels of side force consistent with those predicted using the conical equations. Below a certain critical Reynolds number, any flow asymmetries arising from a geometrical asymmetry are damped with increasing distance downstream from the geometrical asymmetry.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-3295
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  • 116
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    Low-speed aerodynamics of the trailed-rotor high-speed rotorcraft concept in cruise configuration (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An experimental investigation was performed to determine the low-speed aerodynamic performance characteristics of the trailed-rotor high-speed rotorcraft concept in its cruise configuration. A 15 percent scale semispan model was tested at speeds up to 180 knots in the NASA Ames 7-by 10-Foot Subsonic Wind Tunnel. The objective of this investigation was to determine specific aerodynamic performance characteristics to assist in evaluating the trailed rotor as a high-speed rotorcraft. The aerodynamic influence of the following model configuration changes were determined: ailerons, flaps, wing/pod angle, number of trailed blades, trailed-blade twist and azimuth, and wing/pod filet radius. The low-speed performance objectives for the concept were met and results indicated that the trailed-rotor model had no significant adverse aerodynamic characteristics. The optimum low-speed cruise configuration was determined. Results suggest the trailed-rotor concept has better low-speed cruise performance characteristics than the folding tiltrotor configuration.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-3230
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  • 117
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    An experimental and computational study of 2-D blade-vortex interaction (1991)
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    Publication Date: 2019-06-28
    Description: An experimental and computational study is carried out to investigate the dominant physical factors of 2D parallel blade-vortex interaction (BVI) and its noise generation. A shock tube was used to generate a starting vortex which interacted with a target airfoil. Double-exposed holographic interferometry and airfoil surface pressure measurements were employed to obtain quantitative data during the BVI. As a numerical approach, thin-layer Navier-Stokes code, with a multizonal grid, was also used to resolve the phenomena occuring in the BVI, especially in the head-on collision case.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-3277
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  • 118
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    An efficient DSMC algorithm applied to a delta wing (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A new algorithm for 3D direct simulation Monte Carlo (DSMC) is tested and numerical results are compared with wind tunnel data and results obtained earlier with a more traditional DSMC code. The test case is the flowfield around a delta wing at incidence at Knudsen number of 0.016 and Mach number of 20.2. The results are shown to compare favorably with both experimental and earlier numerical results. The new algorithm is described with special emphasis placed on its distinctive features: Cartesian/unstructured combination grid, special body surface definition, discretization in physical space.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1316
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  • 119
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    Hypersonic rarefied flow about a delta wing - Direct simulation and comparison with experiment (1991)
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    Publication Date: 2019-06-28
    Description: Three-dimensional simulations of hypersonic rarefied flow about a delta wing are made using the direct simulation Monte Carlo (DSMC) method of Bird, and the results of the computations are compared with recent experimental data obtained in a vacuum wind tunnel at the DLR in Gottingen, Germany. The present study considers Mach 8.89 nitrogen flow for a range of conditions that include Knudsen numbers of 0.016 to 3.505 for an incidence angle of 30 deg, and angles of incidence of 15 to 60 deg for a constant Knudsen number of 0.389. The calculations provide details concerning the flowfield structure and surface quantities. Comparisons between the calculations and the available experimental measurements are made for aerodynamic and overall heat-transfer coefficients and recovery temperature. The agreement between the measured and calculated data are very good, well within the estimated measurement uncertainty. Comparisons are also made with modified Newtonian and free-molecule theories.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1315
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  • 120
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    Shock wave-boundary layer interactions in rarefied gas flows (1991)
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    Publication Date: 2019-06-28
    Description: A numerical study is presented, using the direct simulation Monte Carlo (DSMC) method, of shock wave-boundary layer interactions in low density supersonic flows. Test cases include two-dimensional, axially-symmetric and three-dimensional flows. The effective displacement angle of the boundary layer is calculated for representative flat plate, wedge, and cone flows. The maximum pressure, shear stress, and heat transfer in the shock formation region is determined in each case. The two-dimensional reflection of an oblique shock wave from a flat plate is studied, as is the three-dimensional interaction of such a wave with a sidewall boundary layer.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1312
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  • 121
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    Hypersonic rarefied flow about a compression corner - DSMC simulation and experiment (1991)
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    Publication Date: 2019-06-28
    Description: The phenomena of shock/boundary-layer interactions and flow separation are investigated using both computational and experimental methods for low-density hypersonic flow about two-dimensional compression corners. The numerical calculations are made with the direct simulation Monte Carlo (DSMC) method. Experimental measurements provide information concerning the flowfield structure and surface flow patterns by means of gas glow discharge and oil flow pictures, respectively. Comparison of the two data sets provides a qualitative basis for assessing the ability of the DSMC method to describe such flows.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1313
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  • 122
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    Multiple shock-shock interference on a cylindrical leading edge (1991)
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    Publication Date: 2019-06-28
    Description: The details of an experimental study of shock wave interference heating on a cylindrical leading edge representative of the cowl of a rectangular hypersonic engine inlet are presented. This Mach 8 study has provided the first detailed pressure and heat transfer rate distributions on a cylinder resulting from a two-dimensional shockwave interference pattern created by two incident oblique shock waves intersecting the cylinder bow shock wave. The peak heat transfer rate was 38 times the undisturbed flow stagnation point level and occurred when the two oblique shock waves coalesced prior to intersecting the cylinder bow shock wave. Development of pressure deflection diagrams identified a new interference pattern consisting of concomitant supersonic jets separated from each other by a shear layer and submerged in the subsonic region between the bow shock wave and body.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1800
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  • 123
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    A comparative study of several compressibility corrections to turbulence models applied to high-speed shear layers (1991)
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    Publication Date: 2019-06-28
    Description: Several recently published compressibility corrections to the standard k-epsilon turbulence model are used with the Navier-Stokes equations to compute the mixing region of a large variety of high speed flows. These corrections, specifically developed to address the weakness of higher order turbulence models to accurately predict the spread rate of compressible free shear flows, are applied to two stream flows of the same gas mixing under a large variety of free stream conditions. Results are presented for two types of flows: unconfined streams with either (1) matched total temperatures and static pressures, or (2) matched static temperatures and pressures, and a confined stream.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1783
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  • 124
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    An experimental study of a three-dimensional shock wave/turbulent boundary-layer interaction at a hypersonic Mach number (1991)
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    Publication Date: 2019-06-28
    Description: Experimental data for a series of three-dimensional shock-wave/turbulent-boundary-layer interaction flows at Mach 8.2 are presented. The test bodies, composed of sharp fins fastened to a flat-plate test surface, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure, heat-transfer, and skin-friction distributions, as well as limited mean flowfield surveys both in the undisturbed and interaction regimes. The data were obtained for the purpose of validating computational models of these hypersonic interactions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1761
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  • 125
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    Hypersonic Shock/Boundary-Layer Interaction Database (1991)
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    Publication Date: 2019-06-28
    Description: Turbulence modeling is generally recognized as the major problem obstructing further advances in computational fluid dynamics (CFD). A closed solution of the governing Navier-Stokes equations for turbulent flows of practical consequence is still far beyond grasp. At the same time, the simplified models of turbulence which are used to achieve closure of the Navier-Stokes equations are known to be rigorously incorrect. While these models serve a definite purpose, they are inadequate for the general prediction of hypersonic viscous/inviscid interactions, mixing problems, chemical nonequilibria, and a range of other phenomena which must be predicted in order to design a hypersonic vehicle computationally. Due to the complexity of turbulence, useful new turbulence models are synthesized only when great expertise is brought to bear and considerable intellectual energy is expended. Although this process is fundamentally theoretical, crucial guidance may be gained from carefully-executed basic experiments. Following the birth of a new model, its testing and validation once again demand comparisons with data of unimpeachable quality. This report concerns these issues which arise from the experimental aspects of hypersonic modeling and represents the results of the first phase of an effort to develop compressible turbulence models.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1763
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  • 126
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    Hypersonic shock-wave turbulent-boundary-layer interaction flows - Experiment and computation (1991)
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    Publication Date: 2019-06-28
    Description: Solutions of the Reynolds-averaged Navier-Stokes equations are presented and compared with experimental surface data for a series of hypersonic shock-wave/turbulent-boundary-layer interaction flows. The turbulence models used include the standard k-epsilon two-equation eddy viscosity model, a two-layer modification to this model, and a third model with several modifications to account for compressibility effects. Both modified models gave significant improvements for all the test flows.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1760
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  • 127
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    Computational and experimental aftbody flow fields for hypersonic, airbreathing configurations with scramjet exhaust flow simulation (1991)
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    Publication Date: 2019-06-28
    Description: Computational results are presented for three issues pertinent to hypersonic, airbreathing vehicles employing scramjet exhaust flow simulation. The first issue consists of a comparison of schlieren photographs obtained on the aftbody of a cruise missile configuration under powered conditions with two-dimensional computational solutions. The second issue presents the powered aftbody effects of modeling the inlet with a fairing to divert the external flow as compared to an operating flow-through inlet on a generic hypersonic vehicle. Finally, a comparison of solutions examining the potential of testing powered configurations in a wind-off, instead of a wind-on, environment, indicate that, depending on the extent of the three-dimensional plume, it may be possible to test aftbody powered hypersonic, airbreathing configurations in a wind-off environment.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1709
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  • 128
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    Super/hypersonic aerodynamic characteristics for a transatmospheric vehicle concept having a minimum drag forebody (1991)
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    Publication Date: 2019-06-28
    Description: Experimental longitudinal and lateral-directional aerodynamic characteristics were obtained for a generic transatmospheric vehicle concept having a replaceable minimum drag forebody shape. The alternate forebody tested was a 1/4-power series body. Tests were made over a range of Mach numbers from 2 to 10 at a nominal Reynolds number, based on a length of 2.3 x 10 to the 8th and angles of attack from -4 to 20 deg. The minimum drag forebody provided significant improvements in minimum drag and L/D for the configuration as well as a longitudinally stabilizing increment. Although the baseline configuration is longitudinally unstable, the L/D improvements at low to moderate angles of attack would enhance cruise performance. Varying wing incidence angles was demonstrated as an effective horizontal trim device without significant trim drag penalties.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1694
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  • 129
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    Unsteady fluid dynamic model for propeller induced flow fields (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A potential flow based three-dimensional panel method was modified to treat time dependent flow conditions in which the body's geometry may vary with time. The main objective of this effort was the study of a flow field due to a propeller rotating relative to a nonrotating body which is otherwise moving at a constant forward speed. Calculated surface pressure, thrust and torque coefficient data for a four-bladed marine propeller/body compared favorably with previously published experimental results.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1664
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    Goertler instability and hypersonic quiet nozzle design (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A concept for nozzle design which incorporates slow expansion rates and a radial flow region is developed for use on a hypersonic (Mach 6) nozzle. It is shown that the boundary-layer suction slot upstream of the nozzle throat is necessary to remove upstream turbulent boundary layers and to initialize a new laminar boundary layer on the downstream nozzle wall. In effect, the laminar boundary-layer flow can be extended more effectively farther downstream by a slow expansion contour based on the present design concept than by a rapid expansion contour used in previous pilot quiet nozzles.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1648
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  • 131
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    The linear stability of three-dimensional boundary-layers over axisymmetric bodies at incidence (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The linear stability of the fully three-dimensional boundary layer formed over a 6:1 prolate spheroid at 10 deg incidence is investigated using a modified version of the linear stability code COSAL. For this case, both Tollmien-Schlichting and cross flow disturbances are relevant in the transition process. The predicted location of the onset of transition using the e exp N method compares favorably with experimental results of Meier and Kreplin (1980). Using a value of N = 10, the predicted location is located approximately 10 percent upstream of the experimentally determined location. Results also indicate that the direction of disturbance propagation is dependent upon the type of disturbance, and hence the dimensional frequency.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1640
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  • 132
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    Numerical study of the X-31 high angle-of-attack flow characteristics (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The application of CFD methods to high-angle-of-attack flight regimes is studied. The X-31 configuration was chosen as the model due to its inherent designed high-angle-of-attack characteristics and capabilities, the existence of a wind tunnel database, and the ongoing flight test program. The flow characteristics of the X-31 at high angle-of-attack are presented and the CFD capability for studying the effects of canard deflection on aerodynamic control during poststall is revealed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1630
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  • 133
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    Comparison of cryogenic wind tunnel, flight, and computational results for a thin, low-aspect-ratio wing (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A research program was conducted at NASA Langley Research Center to build and test a thin, pressure instrumented wing. The wing chosen was the canard of the X-29, which has a maximum thickness of 5 percent of chord. The wing has 90 pressure taps and was built utilizing an advanced laminated metal technique. It was tested in the 0.3-Meter Transonic Cryogenic Tunnel at transonic Mach numbers and over a wide range of Reynolds number. The data are compared with flight data and Navier-Stokes computational results.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1626
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    The effects of shock wave precursors ahead of hypersonic entry vehicles (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A model has been developed to predict the magnitude and characteristics of the shock wave precursor ahead of a hypervelocity vehicle. This model includes both chemical and thermal nonequilibrium, utilizes detailed mass production rates for the photodissociation and photoionization reactions, and accounts for the effects of radiative absorption and emission on the individual internal energy modes of both atomic and diatomic species. Comparison of the present results with shock tube data indicates that the model is reasonably accurate. A series of test cases representing earth aerocapture return from Mars indicate that there is significant production of atoms, ions and electrons ahead of the shock front due to radiative absorption and that the precursor is characterized by an enhanced electron/electronic temperature and molecular ionization. However, the precursor has a negligible effect on the shock layer flow field.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1465
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  • 135
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    Prediction of internal performance for two-dimensional convergent-divergent nozzles (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A three-dimensional code (PAB3D) has been under development to simulate underexpanded and overexpanded supersonic jet exhaust effects on a nonaxisymmetric nozzle/afterbody. The code is a multiblock/multizone technique solving the simplified Navier-Stokes equations. This method was applied to the solution of two nozzle configurations to verify the recent implementation of a nozzle performance calculation package. Static pressure distributions, discharge coefficient and thrust ratio quantities were calculated for on-design and off-design operating conditions for rectangular convergent-divergent nozzles with different throat designs. Approaches involved in computing the flows for these nozzles are presented. Computation of performance parameters were typically within 1 percent of experimental on-design and off-design data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-2369
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  • 136
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    Mixing studies of helium in air at Mach 6 (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Concentration and mean flow measurements with nanoshadowgraphs and surface flow visualization have been obtained by a study of mixing phenomena during gas injection into hypersonic flow. At Mach 6, a comparison of the matched-pressure injection case with the underexpanded case indicated a greater injectant core penetration rate and a greater concentration-decay rate, leading to a shorter distance for the injectant core to reach an H2-air stoichiometric ratio. The entire injectant plume remained supersonic, and only moderate pressure losses were found. While injector yaw did not increase the mixing rates, it led to an increase in overall injectant plume cross-section, with consequent increase in the size of the mixing region.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-2268
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  • 137
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    Supersonic jet mixing enhancement by vortex generators (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Experiments were conducted to explore the effects of vortex generators, in the form of tabs projecting normally into the jet, on the mixing and the far-field noise characteristics of a jet. A converging-diverging nozzle with a design Mach number of 1.36 was used in the experiments. The flow regimes from subsonic to highly underexpanded supersonic conditions were studied. One, two, and four tabs were used and some of the findings of previous investigators were examined and confirmed. The tabs eliminated screech noise from moderately overexpanded cases to highly underexpanded cases. Detailed flow visualizations and measurements showed that two tabs bifurcated the jet at all Mach numbers. While the effect of two tabs was persistent and the jet remained bifurcated, the distortions produced by one and four tabs disappeared by a streamwise distance of approximately 16 jet diameters. Two and four tabs significantly increased the entrainment of ambient air into the jet.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-2263
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  • 138
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    Periodic blade loads of a high speed propeller at small angle of attack (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A code has been developed to predict the periodic aerodynamic loads of an advanced turboprop propeller. The analytical formulation accounts for flow three-dimensionality and flow periodicity due to the propeller inclination. The flow past the blade sections is computed using a thin layer Navier-Stokes solver. An iterative procedure is used to account for the induced axial and rotational velocities. The viscous periodic results are obtained for an eight-bladed Hamilton Standard SR-7L advanced propeller at a cruise Mach number of 0.813 and 35,000 ft. altitude. The results are shown for flow field quantities and performance parameters during the blade passage in the plane of rotation illustrating the periodic nature of blade flow separation and shocks. The time averaged coefficients of thrust and power are computed and compared with available flight test data. The results obtained show excellent agreement at cruise conditions for small nacelle angles of attack.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-2250
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  • 139
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    Use of the PARC code for generic NASP nozzles operating at off-design transonic conditions (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A three-dimensional Navier-Stokes solver, PARC3D, has been investigated for applications to the wind tunnel testing of high-speed nozzle configurations, particularly with extended expansion ramps, operating at transonic off-design conditions. Numerous cases have been run and compared with existing wind tunnel data. The code has been shown to be an efficient and robust tool in the analysis of complex nozzle configurations whose flows are substantially determined by three-dimensional effects. Although the method requires that a grid, generated independently of the code, be supplied to it, the robust nature of the code does not require that this grid be excessively detailed, nor does it seem to require that the grid be fine-tuned for individual cases. The directness and flexibility of the boundary condition specifications allow a variety of configurations and flow conditions to be handled very efficiently.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-2154
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  • 140
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    The effect of bleed configuration on shock/boundary layer interactions (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The flow field characteristics are simulated numerically in an oblique shock wave/laminar boundary layer interaction for three different bleed configurations. The numerical solution for the flow field is obtained for the strong conservation-law form of the two-dimensional compressible Navier-Stokes equations using an implicit scheme. The computations model the flow in the interaction region and inside the bleed slot for an impinging oblique shock of sufficient strength to cause boundary layer separation on a flat plate boundary layer. The computed results are presented for a normal bleed slot at three different locations, upstream, across and downstream of the shock impingement point. The detailed flow characteristics in the interaction zone and inside the bleed slot are compared for the three cases. The resulting surface pressure and shear stress distributions are also compared for the three bleed cases and for the shock wave boundary layer interaction without bleed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-2014
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  • 141
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    Flutter control of incompressible flow turbomachine blade rows by splitter blades (1991)
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    Publication Date: 2019-06-28
    Description: Splitter blades as a passive flutter control technique is investigated by developing a mathematical model to predict the stability of an aerodynamically loaded splittered-rotor operating in an incompressible flow field. The splitter blades, positioned circumferentially in the flow passage between two principal blades, introduce aerodynamic and/or combined aerodynamic-structural detuning into the rotor. The two-dimensional oscillating cascade unsteady aerodynamics, including steady loading effects, are determined by developing a complete first-order unsteady aerodynamic analysis together with an unsteady aerodynamic influence coefficient technique. The torsion mode flutter of both uniformly spaced tuned rotors and detuned rotors are predicted by incorporating the unsteady aerodyamic influence coefficients into a single-degree-of-freedom aeroelastic model. This model is then utilized to demonstrate that incorporating splitters into unstable rotor configurations results in stable splittered-rotor configurations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1900
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  • 142
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    Active unsteady aerodynamic suppression of rotating stall in an incompressible flow centrifugal compressor with vaned diffuser (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A mathematical model is developed to analyze the suppression of rotating stall in an incompressible flow centrifugal compressor with a vaned diffuser, thereby addressing the important need for centrifugal compressor rotating stall and surge control. In this model, the precursor to to instability is a weak rotating potential velocity perturbation in the inlet flow field that eventually develops into a finite disturbance. To suppress the growth of this potential disturbance, a rotating control vortical velocity disturbance is introduced into the impeller inlet flow. The effectiveness of this control is analyzed by matching the perturbation pressure in the compressor inlet and exit flow fields with a model for the unsteady behavior of the compressor. To demonstrate instability control, this model is then used to predict the control effectiveness for centrifugal compressor geometries based on a low speed research centrifugal compressor. These results indicate that reductions of 10 to 15 percent in the mean inlet flow coefficient at instability are possible with control waveforms of half the magnitude of the total disturbance at the inlet.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1898
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    Euler calculations of unsteady transonic flow in cascades (1991)
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    Publication Date: 2019-06-28
    Description: In the present paper, Euler calculations of unsteady transonic flow in cascades are presented. A finite volume scheme is used to discretize the equations, which are implemented on a blade-fitted deformable mesh. The space-discretized equations are integrated forward in time using a multistage Runge-Kutta scheme. Adaptive dissipation terms of the type proposed by Jameson and Baker are added to capture shocks and to suppress nonphysical oscillations. Phase-shifted boundary conditions are used to reduce the computational domain to a single reference passage. No assumptions of small amplitudes or small flow deflections are made. Thus, the present code makes it possible to carry out aeroelastic calculations for cases where the shock strengths and oscillation amplitudes exceed the inherent limitations of potential flow codes.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1104
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  • 144
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    A shock-layer theory based on thirteen-moment equations and DSMC calculations of rarefied hypersonic flows (1991)
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    Publication Date: 2019-06-28
    Description: Grad's thirteen-moment equations are applied to the flow behind a bow shock under the formalism of a thin shock layer. Comparison of this version of the theory with Direct Simulation Monte Carlo calculations of flows about a flat plate at finite attack angle has lent support to the approach as a useful extension of the continuum model for studying translational nonequilibrium in the shock layer. This paper reassesses the physical basis and limitations of the development with additional calculations and comparisons. The streamline correlation principle, which allows transformation of the 13-moment based system to one based on the Navier-Stokes equations, is extended to a three-dimensional formulation. The development yields a strip theory for planar lifting surfaces at finite incidences. Examples reveal that the lift-to-drag ratio is little influenced by planform geometry and varies with altitudes according to a 'bridging function' determined by correlated two-dimensional calculations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0783
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  • 145
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    High alpha aerodynamic control by tangential fuselage blowing (1991)
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    Publication Date: 2019-06-28
    Description: This work explores the effect of tangential blowing on the vortical structures that develop around a tangent-ogive cylinder configuration at high angle of attack. A lateral force results if blowing is applied asymmetrically. The study is conducted numerically by solving the three-dimensional, compressible-flow Navier-Stokes equations. The computation was done for a Reynolds number of 52,000, Mach number of 0.2, blowing momentum coefficients of 0.0, 0.1, 0.2, and 0.4, and angle of attack of 10, 30, and 45 deg. Only asymmetrical blowing was considered. Surface streamlines, helicity contours and pressure distributions are obtained for each case.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0620
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    The symmetric turbulent plane wake downstream of a sharp trailing edge (1991)
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    Publication Date: 2019-06-28
    Description: The analysis and modeling of the symmetric turbulent plane wake downstream of a sharp trailing edge is addressed. A compact description of the flow near the trailing edge is formulated using the results of a previous asymptotic analysis. The new description retains the two-layered structure identified in the previous work, and it clarifies the principal dynamics of the flow in the near-wake outer layer, away from the wake centerline. For zero-pressure-gradient flow, the near-wake outer layer is shown to be represented to leading order by the similarity solution that governs the outer region of the surface boundary layer. The leading perturbation in the outer layer due to the developing near-wake inner-layer flow is identified, and this is shown to be asymptotically smaller than undetermined higher-order terms associated with the surface boundary-layer flow. Results of the new near-wake analysis are used to formulate an algebraic eddy viscosity model for wake flow predictions at arbitrary distances from the trailing edge. The model is used in a numerical solution of the boundary layer equations, and computed velocity and Reynolds stress profiles are shown to compare well with experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0612
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  • 147
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    Recent improvements in the nonequilibrium VSL scheme for hypersonic blunt-body flows (1991)
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    Publication Date: 2019-06-28
    Description: The nonequilibrium viscous shock-layer (VSL) solution scheme is revisited to improve its solution accuracy in the stagnation point region and also to minimize and control the errors in the conservation of elemental mass. The stagnation-point solution is improved by using a second-order expansion for the normal velocity and the elemental mass conservation is improved by directly imposing the element conservation equations as solution constraints. These modifications are such that the general structure and computational efficiency of the nonequilibrium VSL scheme is not affected. This revised nonequilibrium VSL scheme is used to study the Mach 20 flow over a 7-deg sphere-cone vehicle under zero and 20-deg angle-of-attack conditons. Comparisons are made with the corresponding predictions of Navier-Stokes and Parabolized Navier-Stokes solution schemes. The results of these tests show that the nonequilibrium blunt-body VSL scheme is indeed an accurate, fast and extremely efficient means for generating the blunt-body flowfield over spherical nosetips at small-to-large angles of attack.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0469
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    Chemical-kinetic problems of future NASA missions (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Thermochemical nonequilibrium in the shock layer surrounding vehicles entering the atmospheres of earth and Mars at superescape velocities is studied, deriving reaction rate coefficients that reproduce experimental data obtained in shock tubes. Thermodynamic properties and emitted radiation intensities are obtained for shock tube flow and flow in a shock layer over a blunt body. The results indicate that the viscous layer of the ablation product over an ablating heat shield is likely to be in chemical nonequilbrium. For earth entry flight, the thickness of the nonequilbrium region is between and 2 cm at the expected peak radiation point in the aerobraking trajectory, For Martian entry flight it is between 8 and 23 cm. For the earth entry case, nonequilibrium phenomena reduce radiative heating rate, while the opposite occurs for the Martian case. The radiative heat transfer rates are significant for the Mars entry conditions at entry velocities equal to or greater than 7 km/s.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0464
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  • 149
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    Numerical algorithm comparison for the accurate and efficient computation of high-incidence vortical flow (1991)
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    Publication Date: 2019-06-28
    Description: Computations from two Navier-Stokes codes, NSS and F3D, are presented for a tangent-ogive-cylinder body at high angle of attack. Features of this steady flow include a pair of primary vortices on the leeward side of the body as well as secondary vortices. The topological and physical plausibility of this vortical structure is discussed. The accuracy of these codes are assessed by comparison of the numerical solutions with experimental data. The effects of turbulence model, numerical dissipation, and grid refinement are presented. The overall efficiency of these codes are also assessed by examining their convergence rates, computational time per time step, and maximum allowable time step for time-accurate computations. Overall, the numerical results from both codes compared equally well with experimental data, however, the NSS code was found to be significantly more efficient than the F3D code.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0175
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    Effects of sweep angle and passive relaminarization devices on a supersonic swept-cylinder boundary layer (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The effects of passive leading-edge devices on the boundary layer of a swept cylinder in supersonic flow with very low free-stream disturbance levels are investigated. Tests are conducted at Mach 3.5 over a free-stream Reynolds number based on model diameter of 15,000 to 150,000, and sweep angles of 60 deg and 76 deg are used to evaluate possible sweep-angle effects. The devices tested include a sawtoothed leading edge, square device, and a fence. It is observed that at a sweep angle of 76 deg, relaminarization of the supersonic attachment-line flow is achieved by several of the devices, while no devices are successful at a sweep angle of 60 deg.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0066
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    Nonlinear disturbances in a hypersonic laminar boundary layer (1991)
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    Publication Date: 2019-06-28
    Description: Three separate but similar experiments on the growth of naturally-occurring instability waves in the laminar boundary-layers on sharp cones at hypersonic Mach numbers have been conducted. Each provided clear evidence that the theoretically-predicted second mode of instability was responsible for high-amplitude wave trains observed to prevail upstream of boundary-layer transition to turbulence. However, each also seemed to reveal the presence of an additional instability not accounted for by the linear theory. Here, examination is made of the tape-recorded hot-wire anemometer signals of one experiment on a sharp cone at Mach 8 for evidence of nonlinearity, the finding of which would explain the presence of the additional mode as a consequence of harmonic generation. Several approaches for identification of the residual effects of nonlinearity are described and utilized. Also, a simplified model describing certain fluctuation characteristics has been developed. Altogether, the evidence of nonlinear wave development is found to be strong. Quantitative comparisons of linear theory to experiments must be made with caution when nonlinearity is present in the experiment.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0320
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  • 152
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    Wind tunnel blockage study of a generic three-dimensional sidewall compression scramjet inlet at Mach 10 (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A large scale model of a generic three-dimensional sidewall compression scramjet inlet has been designed based on the results of a computational parametric study for testing in the 31-inch Mach 10 Hypersonic Wind Tunnel at the NASA Langley Research Center. In order to increase the instrumentation density in interaction regions for a highly instrumented model, it is desirable to make the model as large as possible. When the cross-sectional area of a model becomes large relative to the inviscid core size of the tunnel, the effects of blockage must be considered. In order to assess these effects, a blockage model (an inexpensive, much less densely instrumented version of the configuration) was fabricated for preliminary testing. Since it was desired to determine both the effect of the model on the performance of the wind tunnel and also to determine if the inlet would start, the model possessed a total of 32 static pressure orifices distributed on the forebody plane and sidewalls; seventeen static pressure orifices on the tunnel wall and 3 pitot probes on the model monitored the tunnel performance. This paper presents the design considerations in the development of the wind tunnel model and the blockage aspects of the effects of contraction ratio, cowl location, Reynolds number, and angle of attack.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0294
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    Development of a solution adaptive unstructured scheme for quasi-3D inviscid flows through advanced turbomachinery cascades (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The objective of the present research is to develop a general solution adaptive scheme for the accurate prediction of inviscid quasi-three-dimensional flow in advanced compressor and turbine designs. The adaptive solution scheme combines an explicit finite-volume time-marching scheme for unstructured triangular meshes and an advancing front triangular mesh scheme with a remeshing procedure for adapting the mesh as the solution evolves. The unstructured flow solver has been tested on a series of two-dimensional airfoil configurations including a three-element analytic test case presented here. Mesh adapted quasi-three-dimensional Euler solutions are presented for three spanwise stations of the NASA rotor 67 transonic fan. Computed solutions are compared with available experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0132
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  • 154
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    Navier-Stokes simulation of a close-coupled canard-wing-body configuration (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The thin-layer Navier-Stokes equations are solved for the flow about a coplanar close-coupled canard-wing-body configuration at a transonic Mach number of 0.90 and at angles of attack ranging from 0 to 12 degrees. The influence of the canard on the wing flowfield, including canard-wing vortex interaction and wing vortex breakdown, is investigated. A study of canard downwash and canard leading-edge vortex effects, which are the primary mechanisms of the canard-wing interaction, is emphasized. Comparisons between the computations and experimental measurements of surface pressure coefficients, lift, drag and pitching moment data are favorable. A grid refinement study for configurations with and without canard shows that accurate results are obtained using a refined grid for angles of attack where vortex burst is present. At an angle of attack of approximately 12 deg, favorable canard-wing interaction which delays wing vortex breakdown is indicated by the computations and is in good agreement with experimental findings.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0070
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  • 155
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    A study of dynamic stall using real time interferometry (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Dynamic stall over an oscillating airfoil in compressible flow was studied using a real-time interferometry technique. Instantaneous flow field data was obtained for various unsteady as well as steady flow conditions. Comparison of steady flow interferograms with those taken in unsteady flow reveal a significant delay in the development of leading edge suction peaks in the unsteady case. The interferograms permit detailed analysis of the leading edge pressure field; as many as 13 pressure values have been obtained around the leading edge in the first 1 percent of the airfoil chord. The results offer a significant new insight into the character of the dynamic stall vortex, and the stall delay that is observed during dynamic motions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0007
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    Aerodynamic control using forebody strakes (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Aerodynamic control using a rotatable, miniature nose-tip strake system was investigated in a water tunnel with a forebody model. Flow visualization and yawing moment measurements were performed. The results show that the system was highly effective in controlling the forebody vortices and producing controlled yawing moments at moderate-to-high angles of attack. In comparison with the forebody strakes used in several previous studies, the system can potentially be substantially smaller in size, simpler in operation, and effective over wide ranges of angles of attack and sideslip.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0618
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  • 157
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    Thin-layer and full Navier-Stokes, locally-conical and three-dimensional asymmetric solutions (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The unsteady, compressible thin-layer and full Navier-Stokes equations are used to numerically simulate steady and unsteady asymmetric, supersonic, locally-conical flows around a 5-deg-semiapex angle circular cone. The main computational scheme used for the present computations is the implicit, upwind, flux-difference splitting, finite-volume scheme. Comparisons of the solutions using the two sets of equations are presented for the flow asymmetry and its control. Computational studies are also presented to investigate the effects of the freestream Reynolds number and the locally-scaled Reynolds number on the flow asymmetry. These studies are carried out using the full Navier-Stokes equations. Three-dimensional, asymmetric flow solutions are also presented for a 5-deg-semiapex angle cone of unit length and a cone-cylinder configuration. The three-dimensional solutions are obtained by using the thin-layer equations and short-duration transient side-slip disturbances along with a very fine grid.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0547
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  • 158
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    Aerothermal study in an axial compression corner with shock impingement at Mach 6 (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An experimental study of an axial-compression interior-corner region with shock impingement was conducted in the NASA Langley 20-inch Mach 6 Tunnel to obtain aerodynamic heating rate distributions. The model, a generic hypersonic engine inlet, had a sharp leading edge forebody, cowl, and interchangeable struts with sharp and blunt leading edges. Experiments were conducted at free-stream unit Reynolds numbers of 1.68 x 10 to the 6th and 3.35 x 10 to the 6th per foot for the blunt and sharp-strut configurations, respectively. For the sharp-strut configuration, with and without trips, the peak heating occurred on the cowl, near the compression corner, downstream of the forebody oblique shock impingement region. The peak heating levels, for both the tripped and untripped boundary-layer flows were about 45 and 50 times the reference laminar heating levels, respectively. The peak heating for the blunt leading edge strut model, which was 30 times the reference value, occurred on the cowl, away from the corner, where the bow shock from the strut leading edge intersected the cowl surface.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0527
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  • 159
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    An examination of coherent structures in a lobed mixer using multifractal measures in conjunction with the proper orthogonal decomposition (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A 'lobed mixer' device that enhances mixing through secondary flows and streamwise vorticity is presently studied within the framework of multifractal-measures theory, in order to deepen understanding of velocity time trace data gathered on its operation. Proper orthogonal decomposition-based knowledge of coherent structures has been applied to obtain the generalized fractal dimensions and multifractal spectrum of several proper eigenmodes for data samples of the velocity time traces; this constitutes a marked departure from previous multifractal theory applications to self-similar cascades. In certain cases, a single dimension may suffice to capture the entire spectrum of scaling exponents for the velocity time trace.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0521
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  • 160
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    An experimental investigation of a supersonic vortical flow (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: Although much research has been done on subsonic vortical flow, the current understanding of these flows remains limited. The effect is characterized of adding swirl to a supersonic jet. The motive is to study the enhancement of supersonic mixing in order to provide more efficient fuel injectors for supersonic combustion (scramjet) engines. The vortical flow was created by tangential injection into a swirl chamber ahead of a converging and/or diverging nozzle. The amount of swirl was varied by changing the number of tangential injection holes and with the removal of the end piece, the jet could be run without swirl. Shadowgraphy, conventional schlieren, and focusing schlieren were used to obtain a qualitative understanding of the jet flow structure. It was determined that an increase in swirl produced an increase in the shear layer growth. Pressure and temperature probes were used to obtain more flow data. The probe data compared favorably with the theoretical calculations, except in the viscous core where viscous effects were not considered negligible. These results verified that a supersonic vortical flow was being created with a maximum helix angle of 33 degs.
    Keywords: AERODYNAMICS
    Type: NASA-TM-105102 , NAS 1.15:105102
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  • 161
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    Experimental investigation of turbulent flow through a circular-to-rectangular transition duct (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: Steady, incompressible, turbulent, swirl-free flow through a circular-to-rectangular transition duck was studied experimentally. The cross-sectional area remains the same at the exit as at the inlet, but varies through the transition section to a maximum value approximately 15 percent above the inlet value. The cross-sectional geometry everywhere along the duct is defined by the equation of a superellipse. Mean and turbulence data were accumulated utilizing pressure and hot-wire instrumentation at five stations along the test section. Data are presented for operating bulk Reynolds numbers of 88,000 and 390,000. Measured quantities include total and static pressure, the three components of the mean velocity vector, and the six components of the Reynolds stress tensor. In addition to the transition duct measurements, a hot-wire technique which relies on the sequential use of single rotatable normal and slant-wire probes was proposed. The technique is applicable for measurement of the total mean velocity vector and the complete Reynolds stress tensor when the primary flow is arbitrarily skewed relative to a plane which lies normal to the probe axis of rotation.
    Keywords: AERODYNAMICS
    Type: NASA-TM-105210 , E-6522 , NAS 1.15:105210
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  • 162
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    A Computational and Experimental Investigation of a Three-Dimensional Hypersonic Scramjet Inlet Flow Field (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A combined computational and experimental parametric study of the internal aerodynamics of a generic three dimensional sidewall compression scramjet inlet configuration was performed. The study was designed to demonstrate the utility of computational fluid dynamics as a design tool in hypersonic inlet flow fields, to provide a detailed account of the nature and structure of the internal flow interactions, and to provide a comprehensive surface property and flow field database to determine the effects of contraction ratio, cowl position, and Reynolds number on the performance of a hypersonic scramjet inlet configuration.
    Keywords: AERODYNAMICS
    Type: NASA-CR-188788 , NAS 1.26:188788
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  • 163
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    On the linear stability of compressible plane Couette flow (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: The linear stability of compressible plane Couette flow is investigated. The correct and proper basic velocity and temperature distributions are perturbed by a small amplitude normal mode disturbance. The full small amplitude disturbance equations are solved numerically at finite Reynolds numbers, and the inviscid limit of these equations is then investigated in some detail. It is found that instability can occur, although the stability characteristics of the flow are quite different from unbounded flows. The effects of viscosity are also calculated, asymptotically, and shown to have a stabilizing role in all the cases investigated. Exceptional regimes to the problem occur when the wavespeed of the disturbances approaches the velocity of either of the walls, and these regimes are also analyzed in some detail. Finally, the effect of imposing radiation-type boundary conditions on the upper (moving) wall (in place of impermeability) is investigated, and shown to yield results common to both bounded and unbounded flows.
    Keywords: AERODYNAMICS
    Type: NASA-CR-187610 , NAS 1.26:187610 , ICASE-91-64 , AD-A240688
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  • 164
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    An effective multigrid method for high-speed flows (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: The use is considered of a multigrid method with central differencing to solve the Navier-Stokes equations for high speed flows. The time dependent form of the equations is integrated with a Runge-Kutta scheme accelerated by local time stepping and variable coefficient implicit residual smoothing. Of particular importance are the details of the numerical dissipation formulation, especially the switch between the second and fourth difference terms. Solutions are given for 2-D laminar flow over a circular cylinder and a 15 deg compression ramp.
    Keywords: AERODYNAMICS
    Type: NASA-CR-187602 , NAS 1.26:187602 , ICASE-91-56 , AD-A240568
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  • 165
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    Multigrid for hypersonic viscous two- and three-dimensional flows (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: The use of a multigrid method with central differencing to solve the Navier-Stokes equations for hypersonic flows is considered. The time dependent form of the equations is integrated with an explicit Runge-Kutta scheme accelerated by local time stepping and implicit residual smoothing. Variable coefficients are developed for the implicit process that removes the diffusion limit on the time step, producing significant improvement in convergence. A numerical dissipation formulation that provides good shock capturing capability for hypersonic flows is presented. This formulation is shown to be a crucial aspect of the multigrid method. Solutions are given for two-dimensional viscous flow over a NACA 0012 airfoil and three-dimensional flow over a blunt biconic.
    Keywords: AERODYNAMICS
    Type: NASA-CR-187603 , NAS 1.26:187603 , ICASE-91-57 , AD-A240707
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  • 166
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    Extension of multigrid methodology to supersonic/hypersonic 3-D viscous flows (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A multigrid acceleration technique developed for solving 3-D Navier-Stokes equations for subsonic/transonic flows was extended to supersonic/hypersonic flows. An explicit multistage Runge-Kutta type of time stepping scheme is used as the basic algorithm in conjunction with the multigrid scheme. Solutions were obtained for a blunt conical frustum at Mach 6 to demonstrate the applicability of the multigrid scheme to high speed flows. Computations were performed for a generic High Speed Civil Transport configuration designed to cruise at Mach 3. These solutions show both the efficiency and accuracy of the present scheme for computing high speed viscous flows over configurations of practical interest.
    Keywords: AERODYNAMICS
    Type: NASA-CR-187612 , NAS 1.26:187612 , ICASE-91-66 , AD-A240396
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  • 167
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    Testing of transition-region models: Test cases and data (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: Mean flow quantities in the laminar turbulent transition region and in the fully turbulent region are predicted with different models incorporated into a 3-D boundary layer code. The predicted quantities are compared with experimental data for a large number of different flows and the suitability of the models for each flow is evaluated.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4371 , NAS 1.26:4371
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  • 168
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    Investigation of Advanced Counterrotation Blade Configuration Concepts for High Speed Turboprop Systems. Task 2: Unsteady Ducted Propfan Analysis (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The primary objective was the development of a time dependent 3-D Euler/Navier-Stokes aerodynamic analysis to predict unsteady compressible transonic flows about ducted and unducted propfan propulsion systems at angle of attack. The resulting computer codes are referred to as Advanced Ducted Propfan Analysis Codes (ADPAC). A computer program user's manual is presented for the ADPAC. Aerodynamic calculations were based on a four stage Runge-Kutta time marching finite volume solution technique with added numerical dissipation. A time accurate implicit residual smoothing operator was used for unsteady flow predictions. For unducted propfans, a single H-type grid was used to discretize each blade passage of the complete propeller. For ducted propfans, a coupled system of five grid blocks utilizing an embedded C grid about the cowl leading edge was used to discretize each blade passage. Grid systems were generated by a combined algebraic/elliptic algorithm developed specifically for ducted propfans. Numerical calculations were compared with experimental data for both ducted and unducted flows.
    Keywords: AERODYNAMICS
    Type: NASA-CR-187106 , NAS 1.26:187106
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  • 169
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    Three dimensional unstructured multigrid for the Euler equations (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: The three dimensional Euler equations are solved on unstructured tetrahedral meshes using a multigrid strategy. The driving algorithm consists of an explicit vertex-based finite element scheme, which employs an edge-based data structure to assemble the residuals. The multigrid approach employs a sequence of independently generated coarse and fine meshes to accelerate the convergence to steady-state of the fine grid solution. Variables, residuals and corrections are passed back and forth between the various grids of the sequence using linear interpolation. The addresses and weights for interpolation are determined in a preprocessing stage using linear interpolation. The addresses and weights for interpolation are determined in a preprocessing stage using an efficient graph traversal algorithm. The preprocessing operation is shown to require a negligible fraction of the CPU time required by the overall solution procedure, while gains in overall solution efficiencies greater than an order of magnitude are demonstrated on meshes containing up to 350,000 vertices. Solutions using globally regenerated fine meshes as well as adaptively refined meshes are given.
    Keywords: AERODYNAMICS
    Type: NASA-CR-187565 , NAS 1.26:187565 , ICASE-91-41 , AD-A237201
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  • 170
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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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  • 171
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    Multigrid solution of compressible turbulent flow on unstructured meshes using a two-equation model (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The system of equations consisting of the full Navier-Stokes equations and two turbulence equations was solved for in the steady state using a multigrid strategy on unstructured meshes. The flow equations and turbulence equations are solved in a loosely coupled manner. The flow equations are advanced in time using a multistage Runge-Kutta time stepping scheme with a stability bound local time step, while the turbulence equations are advanced in a point-implicit scheme with a time step which guarantees stability and positively. Low Reynolds number modifications to the original two equation model are incorporated in a manner which results in well behaved equations for arbitrarily small wall distances. A variety of aerodynamic flows are solved for, initializing all quantities with uniform freestream values, and resulting in rapid and uniform convergence rates for the flow and turbulence equations.
    Keywords: AERODYNAMICS
    Type: NASA-CR-187513 , NAS 1.26:187513 , ICASE-91-11 , AD-A233443
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  • 172
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    In-flight flow visualization characteristics of the NASA F-18 high alpha research vehicle at high angles of attack (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Surface and off-surface flow visualization techniques were used to visualize the 3-D separated flows on the NASA F-18 high alpha research vehicle at high angles of attack. Results near the alpha = 25 to 26 deg and alpha = 45 to 49 deg are presented. Both the forebody and leading edge extension (LEX) vortex cores and breakdown locations were visualized using smoke. Forebody and LEX vortex separation lines on the surface were defined using an emitted fluid technique. A laminar separation bubble was also detected on the nose cone using the emitted fluid technique and was similar to that observed in the wind tunnel test, but not as extensive. Regions of attached, separated, and vortical flow were noted on the wing and the leading edge flap using tufts and flow cones, and compared well with limited wind tunnel results.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4193 , H-1576 , NAS 1.15:4193
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  • 173
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    Rotorcraft application of advanced computational aerodynamics (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: The objective was to develop the capability to compute the unsteady viscous flow around rotor-body combinations. In the interest of tractability, the problem was divided into subprograms for: (1) computing the flow around a rotor blade in isolation; (2) computing the flow around a fuselage in isolation, and (3) integrating the pieces. Considerable progress has already been made by others toward computing the rotor in isolation (Srinivasen) and this work focused on the remaining tasks. These tasks required formulating a multi-block strategy for combining rotating blades and nonrotating components (i.e., a fuselage). Then an appropriate configuration was chosen for which suitable rotor body interference test data exists. Next, surface and volume grids were generated and state-of-the-art CFD codes were modified and applied to the problem.
    Keywords: AERODYNAMICS
    Type: NASA-CR-187767 , NAS 1.26:187767 , MCAT-91-001
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  • 174
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    Free stream capturing in fluid conservation law for moving coordinates in three dimensions (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The free-stream capturing technique for both the finite-volume (FV) and finite-difference (FD) framework is summarized. For an arbitrary motion of the grid, the FV analysis shows that volumes swept by all six surfaces of the cell have to be computed correctly. This means that the free-stream capturing time-metric terms should be calculated not only from a surface vector of a cell at a single time level, but also from a volume swept by the cell surface in space and time. The FV free-stream capturing formulation is applicable to the FD formulation by proper translation from an FV cell to an FD mesh.
    Keywords: AERODYNAMICS
    Type: NASA-CR-177572 , A-91022 , NAS 1.26:177572
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  • 175
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    Static internal performance of an axisymmetric nozzle with multiaxis thrust-vectoring capability (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the static test facility of the Langley 16 Foot Transonic Tunnel in order to determine the internal performance characteristics of a multiaxis thrust vectoring axisymmetric nozzle. Thrust vectoring for this nozzle was achieved by deflection of only the divergent section of this nozzle. The effects of nozzle power setting and divergent flap length were studied at nozzle deflection angles of 0 to 30 at nozzle pressure ratios up to 8.0.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4237 , L-16809 , NAS 1.15:4237
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  • 176
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    Internal performance of a hybrid axisymmetric/nonaxisymmetric convergent-divergent nozzle (1991)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: An investigation was conducted in the static test facility of the Langley 16-foot transonic tunnel to determine the internal performance of a hybrid axisymmetric/nonaxisymmetric nozzle in forward-thrust mode. Nozzle cross-sections in the spherical convergent section were axisymmetric whereas cross-sections in the divergent flap area nonaxisymmetric (two-dimensional). Nozzle concepts simulating dry and afterburning power settings were investigated. Both subsonic cruise and supersonic cruise expansion ratios were tested for the dry power nozzle concepts. Afterburning power configurations were tested at an expansion ratio typical for subsonic acceleration. The spherical convergent flaps were designed in such a way that the transition from axisymmetric to nonaxisymmetric cross-section occurred in the region of the nozzle throat. Three different nozzle throat geometries were tested for each nozzle power setting. High-pressure air was used to simulate jet exhaust at nozzle pressure ratios up to 12.0.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4230 , L-16816 , NAS 1.15:4230
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    Aerothermodynamic environments of aerobraking vehicles for manned Mars missions (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The aerothermodynamic environments of manned spacraft aerobraking in the Martian and earth atmospheres are evaluated. Thermal performance of aerobrake concepts are examined for current cryogenic-aerobrake and advanced propulsion missions entailing three different modes of aerobraking: (1) aerocapture into an orbit about Mars, (2) descent and landing at Mars, and (3) Mars return direct entry at earth. Analyses for these vehicles and modes included both radiative and convective heating, where radiative heating is shown to be a significant portion of the total stagnation point heating induced on the vehicle. A comprehensive parametric study of the effects of ballistic coefficient, nose radius, entry velocity, and L/D on stagnation point heating is described. Optimal nose radii for ranges of ballistic coefficient and entry velocity are determined. The peak heating rates are shown to be 83 W/sq cm and 90 W/sq cm for a low and high L/D Mars transfer vehicle configuration, respectively. Heating profiles for these vehicles using boundary layer techniques show that a high L/D shape will result in a smaller high-temperature region provided the flow is laminar. An examination of a crew return vehicle for a Mars return direct entry trajectory shows that the thermal protection for this aerobrake will require an ablative material for heat rejection due to the large heating rates (about 1 kW/sq cm).
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-2872
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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
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    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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    The structure and development of streamwise vortex arrays embedded in a turbulent boundary layer (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation of the structure and development of streamwise vortices embedded in a turbulent boundary layer was conducted. The vortices were generated by a single spanwise row of rectangular vortex generator blades. A single embedded vortex was examined, as well as arrays of embedded counter rotating vortices produced by equally spaced vortex generators. Measurements of the secondary velocity field in the crossplane provided the basis for characterization of vortex structure. Vortex structure was characterized by four descriptors. The center of each vortex core was located at the spanwise and normal position of peak streamwise vorticity. Vortex concentration was characterized by the magnitude of the peak streamwise vorticity, and the vortex strength by its circulation. Measurements of the secondary velocity field were conducted at two crossplane locations to examine the streamwise development of the vortex arrays. Large initial spacings of the vortex generators produced pairs of strong vortices which tended to move away from the wall region while smaller spacings produced tight arrays of weak vortices close to the wall. A model of vortex interaction and development is constructed using the experimental results. The model is based on the structure of the Oseen Vortex. Vortex trajectories are modelled by including the convective effects of neighbors.
    Keywords: AERODYNAMICS
    Type: NASA-TM-105211 , E-6523 , NAS 1.15:105211
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    Transonic and supersonic Euler computations of vortex-dominated flow fields about a generic fighter (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Flow fields about a generic flighter model were computed using FL057, a 3-D, finite volume Euler code. Computed pressure coefficients, forces, and moments at several Mach numbers (0.6, 0.8, 1.2, 1.4, and 1.6) are compared with wind tunnel data over a wide range of angles of attack in order to determine the applicability of the code for the analysis of fighter configurations. Two configurations were studied, a wing-body and a wing-body-chine. FL057 predicted pressure distributions, forces, and moments well at low angles of attack, at which the flow was fully attached. The FL057 predictions were also accurate for some test conditions once the leading edge vortex became well established. At the subsonic speeds, FL057 predicted vortex breakdown earlier than that seen in the experimental results. Placing the chine on the forebody delayed the onset of bursting and improved the correlation between numerical and experimental data at the subsonic conditions.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3156 , A-90161 , NAS 1.60:3156
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    Measurements of forces, moments, and pressures on a generic store separating from a box cavity at supersonic speeds (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted to measure the forces, moments, and pressure distributions on the generic store separating from a rectangular box cavity contained in a flat plate surface at supersonic speeds. Pressure distributions inside the cavity and oil flow and vapor-screen photographs of the cavity flow field were also obtained. The measurements were obtained for the store separating from a flat plate surface, from two shallow cavities having length to depth ratios (L/h) of 16.778 and 12.073, and from a deep cavity having L/h = 6.730. Measurements for the shallow cavities were obtained both with and without rectangular doors attached to sides of the cavities. The tests were conducted at free stream Mach numbers of 1.69, 2.00 and 2.65 for a free stream Reynolds number per foot of 2 x 10(exp 6). Presented here are a discussion of the results, a complete tabulation of the pressure data, figures of both the pressure and force and moment data, and representative oil flow and vapor screen photographs.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3110 , L-16866 , NAS 1.60:3110
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  • 182
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    Laser velocimetry measurements of oscillating airfoil dynamic stall flow field (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Ensemble-averaged two-component velocity measurements over an airfoil experiencing oscillatory dynamic stall under compressibility conditions were obtained. The measurements show the formation of a separation bubble over the airfoil that persists till angles of attack close to when the dynamic stall vortex forms and convects. The fluid attains mean velocities as large as 1.6 times the free stream velocity with instantaneous values of 1.8 times the free stream velocity. The airfoil motion induces these large velocities in regions that are far removed from the surface.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1799
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  • 183
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    Computational studies of a fluid spike as a leading edge protection device for shock-shock interference heating (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The effectiveness of a fluid spike as a device to protect leading edges of hypersonic atmospheric flight vehicles from high aerothermal loads produced by complex shock-shock interference is studied. The two-dimensional Navier-Stokes equations are solved using an unstructured cell-centered, fully implicit, flux-difference split algorithm. Adaptively generated unstructured meshes are employed. A type IV shock-shock interference for Mach 8 flow on a cylindrical leading edge with and without a small contraflow supersonic jet (fluid spike) placed at two different locations on the body is solved. A typical flow past a blunt body with a type IV shock-shock interference produces very high pressures and heat fluxes on the leading edge. Present results indicate that a fluid spike displaces the bow shock further in front of the body and modifies the shock-shock interference pattern. This leads to reduced peak pressures and heat fluxes on the body.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1734
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  • 184
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    A powered lift experiment for CFD validation (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The transitional flight characteristics of a geometrically simplified STOVL aircraft configuration were measured in the NASA Ames 7- by 10-Foot Wind Tunnel. The experiment is the first in a sequence of tests designed to provide detailed data for evaluating the capability of computational fluid dynamics methods to predict the important flow parameters for powered lift. The model consists of a 60-deg delta wing planform with two circular high-pressure air jets located in a blended fuselage. The measured flows have a maximum freestream Mach number of 0.2. The flow is sonic and at ambient temperature in both jets. The data presented include forces and moments measured using an internal balance, pressures measured at the 281 surface pressure ports, and jet pressures and temperatures. Measurements of the flow are also made in the tunnel test section upstream and downstream of the model and at the jet exists to provide boundary conditions for the planned computations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1731
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  • 185
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    Computation of saddle point of attachment (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Low-speed flows over a cylinder mounted on a flat plate are studied numerically in order to confirm the existence of a saddle point of attachment in the flow before an obstacle, to analyze the flow characteristics near the saddle point theoretically, and to address the significance of the saddle point of attachment to the construction of external flow structures, the interpretation of experimental surface oil-flow patterns, and the theoretical definition of three-dimensional flow separation. Two numerical codes, one for an incompressible flow and another for a compressible flow, are used for various Mach numbers, Reynolds numbers, grid sizes, and numbers of grid points. It is pointed out that the potential presence of a saddle point of attachment means that a line of 'oil accumulation' from both sides of a skin-friction line emanating outward from a saddle point can be either a line of separation or a line of attachment.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1713
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  • 186
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    Numerical performance estimates for a generic hypersonic forebody (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The computation of a hypersonic flow past the forebody of a hypersonic vehicle is analyzed for two sets of flow conditions: the first represents flow conditions of an experimental test of the model in the Calspan 96-inch shock tunnel, and the second set is chosen to represent actual flight conditions. Solutions are derived for sharp- and blunt-nose versions of the geometry in order to understand the effects of the entropy layer on the forebody pressure and heat transfer rates. Some sensitivity of predicted heating rates to grid refinement is observed, but it is found to be small compared to the effects of bluntness. Real-gas effects are studied for the blunt-nose version at flight-like conditions, and these effects are found to have a significant effect on inlet-face performance measures such as mass capture and kinetic energy efficiency.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1695
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  • 187
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    An efficient method for estimating neighboring steady-state numerical solutions to the Euler equations (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The paper concentrates on a linear approximation method for predicting the changes occurring in steady-state numerical solutions of the Euler equations as a consequence of small changes in the independent variables which control the problem. The importance of proper boundary-condition treatment and other issues concerning the problem are covered along with the importance of proper algorithm selection for a fully supersonic inviscid flow. The method is applied to a subsonic nozzle involving variation of the pressure on the outflow boundary and to a supersonic inlet involving variation of the inflow Mach number. In the subsonic test case, the comparisons between the predicted and conventional numerical solutions are shown to be good, while in the supersonic test case, the agreement between the approximation method and conventional numerical solution starts out well but rapidly degenerates at some point in the flowfield as the perturbation of the boundary conditions is increased.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1680
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  • 188
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    Waves produced from a harmonic point source in a supersonic boundary layer (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The disturbance wave pattern produced by a harmonic point source in a compressible flat-plate boundary layer is computed using linear stability theory and direct numerical integration approach. Receptivity coefficients are computed for the spectrum of spanwise modes generated at the source. The effect of boundary layer growth on the development of linear waves is determined by using the method of multiple scales. Results are presented for Mach numbers of 0, 2, 4.5, and 7. It is found that disturbances spread in wedge-shaped regions behind the source and the wedge angle decreases with Mach number. The lateral spreading angle for the instability waves turns out to be quite close to the angle found experimentally for turbulence lateral contamination.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1646
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  • 189
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    Potential flow applications to complex configurations (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Recent advances in CFD methods have enabled the analytic calculation of the carriage loads for stores mounted on complex aircraft. The latest results have demonstrated excellent agreement with test data for the F-15 at M = 0.98. However, in a preliminary design environment, the necessity of generating and validating a Euler grid to fit the aircraft and store arrangement may not be feasible, particularly when effects of configuration changes are considered. For that reason alternative approaches which require less time to arrive at an answer deserve consideration. The paper presents the results of a study to determine if potential flow solutions can give acceptable estimates of store carriage loads at transonic speeds in a timely manner.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1634
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  • 190
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    Compressible stability of growing boundary layers using parabolized stability equations (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The parabolized stability equation (PSE) approach is employed to study linear and nonlinear compressible stability with an eye to providing a capability for boundary-layer transition prediction in both 'quiet' and 'disturbed' environments. The governing compressible stability equations are solved by a rational parabolizing approximation in the streamwise direction. Nonparallel flow effects are studied for both the first- and second-mode disturbances. For oblique waves of the first-mode type, the departure from the parallel results is more pronounced as compared to that for the two-dimensional waves. Results for the Mach 4.5 case show that flow nonparallelism has more influence on the first mode than on the second. The disturbance growth rate is shown to be a strong function of the wall-normal distance due to either flow nonparallelism or nonlinear interactions. The subharmonic and fundamental types of breakdown are found to be similar to the ones in incompressible boundary layers.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1636
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  • 191
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    A parametric study of the leading edge attachment line for the F-16XL (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A three-dimensional Navier-Stokes code is used to computationally simulate the flow about a modified F-16XL. Transition mechanisms (e.g. attachment line location and crossflow instability) near the swept wing leading edge are analyzed in detail. Flow visualization is used to study the influence of angle-of-attack on the aforementioned transition mechanisms. Validation of the code is accomplished by comparison of numerically generated surface pressures with that obtained by in-flight experiments.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1621
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  • 192
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    Flight test and numerical simulation of transonic flow around YAV-8B Harrier II wing (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A computational fluid dynamics (CFD) method is used to study the aerodynamics of the YAV-8B Harrier II wing in the transonic region. A numerical procedure is developed to compute the flow field around the complicated wing-pylon-fairing geometry. The surface definition of the wing and pylons were obtained from direct measurement using theodolite triangulation. A thin-layer Navier-Stokes code with the Chimera technique is used to compute flow solutions. The computed pressure distributions at several span stations are compared with flight test data and show good agreement. Computed results are correlated with flight test data that show the flow is severely separated in the vicinity of the wing-pylon junction. Analysis shows that shock waves are induced by pylon swaybrace fairings, that the flow separation is much stronger at the outboard pylon and that the separation is caused mainly by the crossflow passing the geometry of wing-pylon junction.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1628
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  • 193
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    A study of flow separation on an aspect ratio three flap at Mach number 2.4 (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Static pressure distributions and dark central ground interferometry are used to document the flow on a flap over a range of Reynolds numbers and deflection angles. The pressure profiles and interferograms are analyzed to distinguish laminar, transitional and turbulent flows in attached, incipient and separated cases. Sideplates are used to compare two and three dimensional effects. Data is also presented for extended flaps used to isolate the effect of the trailing edge expansion. The issues of unsteadiness and hysteresis are also addressed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1620
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  • 194
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    Enhancements to the high speed convective heating and viscous drag prediction techniques of the Aerodynamic Preliminary Analysis System (APAS) (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: APAS is an interactive computer program that allows a user to quickly estimate the aerodynamic characteristics of aerospace and aeronautical vehicles throughout the speed range using a single geometry definition. This report documents the major differences between the high speed analysis portion of the APAS and the Mark III version of the hypersonic arbitrary-body program from which it has evolved and compares convective heating and viscous drag results from the APAS with CFD results, experimental data, and flight data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1435
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  • 195
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    Viscous equilibrium computations using program LAURA (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) has been modified to compute viscous equilibrium flow. Periodic calls to the thermodynamic and transport property curve-fits enable solutions to be computed for small percentage increase in computer time when compared with perfect gas times. The code is used to compute the hypersonic flow over slender and blunt cones, and solutions are compared with other computational techniques and flight data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1389
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  • 196
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    Effect of non-equilibrium flow chemistry and surface catalysis on surface heating to AFE (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The effect of nonequilibrium flow chemistry on the surface temperature distribution over the forebody heat shield on the Aeroassisted Flight Experiment (AFE) vehicle was investigated using a reacting boundary-layer code. Computations were performed by using boundary-layer-edge properties determined from global iterations between the boundary-layer code and flow field solutions from a viscous shock layer (VSL) and a full Navier-Stokes solution. Surface temperature distribution over the AFE heat shield was calculated for two flight conditions during a nominal AFE trajectory. This study indicates that the surface temperature distribution is sensitive to the nonequilibrium chemistry in the shock layer. Heating distributions over the AFE forebody calculated using nonequilibrium edge properties were similar to values calculated using the VSL program.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1373
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  • 197
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    A parametric study of dissociation and ionization models at 12 km/sec (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Thermochemical nonequilibrium-solution dependence on available models for the chemical reaction rates is examined. Solutions from the Kang and Dunn (1973) reaction-rate set, the Park rate set of 1987, and the Park rate set of 1991 are compared. The blunt-nosed, axisymmetric geometry considered is a 60-deg sphere cone with nose radius of 1.07 m and cicular aft skirt. The nonequilibrium test case is 12 km/sec entry into the earth's atmosphere at 80 km altitude. The model variations are implemented into the Langley aerothermodynamics upwind relaxation algorithm code. While variations in the reaction rates have no effect on the surface pressure distribution and little effect on the convective heating, the effect on degree of ionization and radiative heating can be a factor of three.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1368
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  • 198
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    An approximate viscous shock layer approach to calculating hypersonic flows about blunt-nosed bodies (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An approximate axisymmetric method has been developed which can reliably calculate fully viscous hypersonic flows over blunt-nosed bodies. By substituting Maslen's second order pressure expression for the normal momentum equation, a simplified form of the viscous shock layer (VSL) equations is obtained. This approach can solve both the subsonic and supersonic regions of the shock layer without a starting solution for the shock shape. Since the method is fully viscous, the problems associated with coupling a boundary-layer solution with an inviscid-layer solution are avoided. This procedure is significantly faster than the parabolized Navier-Stokes (PNS) or VSL solvers and would be useful in a preliminary design environment. Problems associated with a previously developed approximate VSL technique are addressed. Surface heat transfer and pressure predictions are comparable to both VSL results and experimental data. The present technique generates its own shock shape as part of its solution, and therefore could be used to provide more accurate initial shock shapes for higher-order procedures which require starting solutions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1348
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  • 199
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    Energetics of gas-surface interactions in transitional flows at entry velocities (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The direct simulation Monte Carlo (DSMC) method has been used to calculate the molecular velocity and energy distributions of molecules striking a surface after traversing a shock layer in hypersonic transitional flow. The calculations were performed for a 1.6-m-diameter sphere at a nominal velocity for re-entry of 7.5 km/s over an altitude range of 130 to 90 km. Real gas effects and chemical reactions were included in the DSMC simulations. Results are presented for these conditions and the need for gas-surface interaction experiments is discussed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1338
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  • 200
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    Prediction of ice shapes and their effect on airfoil performance (1991)
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    Publication Date: 2019-06-28
    Description: Calculations of ice shapes and the resulting drag increases are presented for experimental data on an NACA 0012 airfoil. They were made with a combination of LEWICE and interactive boundary-layer codes for a wide range of conditions which include airspeed and temperature, the droplet size and liquid water content of the cloud, and the angle of attack of the airfoil. In all cases the calculated results account for the drag increase due to ice accretion and, in general, show good agreement with data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0264
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