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  • AERODYNAMICS  (1,175)
  • 1985-1989  (1,175)
  • 1950-1954
  • 1988  (546)
  • 1987  (629)
Collection
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  • 1985-1989  (1,175)
  • 1950-1954
Year
  • 101
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    Configuration trade and code validation study on a conical hypersonic vehicle (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A study has been conducted on a generic wing-cone transatmospheric vehicle at Mach numbers form 2.5 to 4.5. The objectives of the study were to experimentally define the aerodynamic characteristics of the vehicle and evaluate several computational aerodynamic prediction methods through comparison with the experimental results. The baseline wing-cone configuration fuselage consisted of a 5 deg half-angle cone forebody, cylindrical midbody, and 9 deg truncated cone afterbody. The 4-percent-thick diamond airfoil wing had an aspect ratio of 1. Several configuration variables were investigated to provide trade information on canard, wing-position and incidence, vertical tail, and nose bluntness effects. Results of the study show that wing-position and wing-incidence effects on the longitudinal aerodynamic characteristics can be significantly influenced by wing-body interference. The use of positive wing incidence to provide favorable forebody orientation for possible inlet performance improvement is accompanied by trim drag and lift-drag ratio penalties. The lateral-directional stability characteristics were strongly influenced by the location of the vertical tails. The higher-order full-potential method provided better estimates of the aerodynamic characteristics than either the linearized supersonic potential method or the tangent-cone/tangent-wedge/shock-expansion on method.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-4505
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    NASA TECHNICAL REPORTS
  • 102
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    Boundary layer crossflow stabilization of high subsonic speed LFC transport airplanes (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: With a view to the elaboration of the design of Mach 0.83 and 0.97 cruise-speed long-range aircraft employing LFC, a study is conducted of the laminar flow characteristics of supercritical airfoils of blunt leading-edge X88 type, for the case of lightly loaded wings that dispense with leading-edge flaps for low-speed operations. The boundary layer crossflow in the front acceleration zone of these airfoils' upper surface is optimally stabilized by suction in the upstream portion of the zone, yielding a crossflow that is neutrally stable.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0275
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  • 103
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    Transonic flow modes of an axisymmetric blunt body (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An experimental investigation of transonic flow past a plane-nosed circular cylinder with a plane-nosed circular probe extended coaxially ahead is reported. The possibilities of significant transonic drag reduction and the fluid mechanic phenomena which occur are examined. The probe length and diameter and the approaching flow Mach number are the independent variables. The relations which exist among the probe/cylinder geometry, Mach number, and flow field as revealed by measurements of the drag forces acting on the body are explored.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3536
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  • 104
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    Summary of low-speed wind tunnel results of several high-speed counterrotation propeller configurations (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The low speed aerodynamic performance characteristics of several advanced counterrotation pusher propeller configurations with cruise design Mach numbers of 0.72 and 0.80 were investigated in the NASA Low Speed Wind Tunnel. The tests were conducted at Mach numbers representative of the takeoff and landing flight regime. The investigation included: (1) the propeller performance characteristics over a range of blade angle settings and rotational speeds at a Mach number of 0.20; (2) the effect on the propeller performance of varying the axial rotor spacing and mismatching the power and rotational speeds on the propeller rotors; and (3) determining the reverse thrust performance characteristics at Mach numbers of 0.0, 0.10, 0.15 and 0.20. The results of the investigation indicated that the overall low speed performance of the counterrotation propeller configurations was reasonable.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3149
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  • 105
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    A block space-marching method for the Navier-Stokes equations (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A solution procedure is presented for predominantly supersonic viscous flows. The procedure approximately solves the Navier-Stokes equations by marching blocks of grid points in the streamwise direction and solving the fully elliptic equations within each block. In this manner elliptic effects of limited streamwise extent may be accurately calculated. Results are presented for calculations of a Mach 2 laminar flat-plate shock/boundary-layer interaction, and for a Mach 10 hypervelocity interceptor.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3199
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  • 106
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    A comparison of experimental and computational results for 5 and 10 degree cones at high Mach numbers (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Experiments were performed on 5 and 10 deg slender cones at a velocity of approximately 5 km/sec in the NASA-Ames ballistic ranges. The flowfields for the cones were computed using ideal-gas and chemical nonequilibrium-air parabolized Navier-Stokes codes. Experimentally determined drag coefficients and shock shapes are compared with the results of the computer codes. Both the flight-data analysis methods and the computational codes are examined to achieve the most meaningful comparison. Under the conditions of the experiments, skin-friction drag makes up approximately 50 percent of the total drag for the 5 deg cone and 30 percent of the total drag for the 10 deg cone. Computed drag coefficients of the 10 deg cone agree well with the experimental values; predictions fall below the experimental values for the 5 deg cone.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2705
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  • 107
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    A computational study of heat transfer for laminar hypervelocity flows (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Computational Fluid Dynamics (CFD) codes are routinely used to predict the flowfield and the heating environment around complex reentry configurations. At hypervelocities, where the velocity is greater than 3 km/sec, the AFWAL version of the blunt body code predicts the correct surface pressure distributions but underpredicts laminar wall heat fluxes. This study was performed to determine the reasons for the underprediction. The computer code chosen solves thin-layer Navier-Stokes equations in a time-asymptotic manner and assumes a constant isentropic exponent. Flowfields around a spherical configuration at various entry velocities are computed. The computed pressure distributions agree well with the tabulated, inviscid results of Lyubimov and Rusanov for entry velocities ranging from 0.6 to 5.92 km/sec. At hypervelocities, the calculated stagnation point heat transfer rates were lower by roughly fifty percent when compared to engineering correlations available in the literature. Good comparisons between heat transfer rates are obtained at hypervelocity entry conditions provided the CFD code is modified to include equilibrium air properties.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2666
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  • 108
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    Stagnation flow field analysis for an aeroassist flight experiment vehicle (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Continuum methods are used to analyze the stagnation flow field of the aeroassist flight experiment (AFE) vehicle. For the lower altitude portion of an AFE trajectory, the viscous shock-layer equations are employed. At higher altitudes, the full Navier-Stokes equations with chemical nonequilibrium and surface slip are used. Particular attention is given to the effect of surface catalyticity on surface heating, electron number density, and flow field structure.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2613
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  • 109
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    An implicit Navier-Stokes analysis of turbine rotor-stator interaction (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An implicit Navier-Stokes analysis using a single deforming mesh has been developed for the unsteady rotor-stator interaction problem. The technique has been used to simulate the flow through a turbine stator-rotor stage. Periodic two-dimensional solutions have been obtained using 1000 time steps per cycle without iteration at each time step. Computed surface pressure distributions compare favorably with experimental data available for this configuration.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3090
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  • 110
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    Artificial dissipation models for hypersonic external flow (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Four artificial dissipation models which augment central difference schemes were examined for hypersonic external flows. The models were a first and third order dissipation model, a directionally scaled first and third order dissipation model, a flux limited dissipation model, and a flux difference split dissipation model. Each model was implemented in the lower-upper symmetric-Gauss-Seidel (LU-SGS) algorithm to solve the full Navier-Stokes equations. The latter two models can be regarded as total variation diminishing (TVD) schemes. Test results for model problems showed that the flux limited dissipation model was robust enough to predict a high speed blunt body flow with strong shock and expansion waves. The flux difference split dissipation model was capable of shock capturing with higher resolution, but was less robust. First and third order dissipation models turned out to be neither accurate nor robust enough for high Mach number flow computations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3277
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  • 111
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    Turbulence modeling in hypersonic inlets (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A study is conducted to analyze the performance of different turbulence models when applied to flow through a Mach 7.4 hypersonic inlet. The analysis, which is two-dimensional, is done by comparing computational results from a Parabolized Navier-Stokes code and a full Navier-Stokes code, with experimental data. The McDonald-Camarata (MC) and Baldwin-Lomax (BL) models were the two zero-equation models used in the study. The Turbulent Kinetic Energy (TKE) model was chosen as a representative higher order model. The MC model, when run with transition of flow, provides a solution which compares excellently with the data. Transition has a first order effect on the overall solution provided by the code. The BL model predicts separation of flow in the inlet, which contradicts experimental findings. The TKE model does not perform any better than the MC and BL models, despite the fact that it is a higher order turbulence model. The BL and TKE models predict transition in the inlet at a location which is much earlier than observed in the experiment. This may be attributed to the empirical constants used to determine the point of transition.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2957
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  • 112
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    Hypersonic turbulent wall boundary layer computations (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The Baldwin-Lomax (1978) algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer-layer eddy-viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith (1974) model for a flat plate at hypersonic speed, the new values of the coefficients were obtained. The results show that the values of C(cp) and C(kleb) are functions of both Mach number and wall temperature ratio. The C(cp) and C(kleb) variations with Mach number and wall temperature were used for the calculations of both a 4-deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin-layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2829
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  • 113
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    Navier-Stokes structure of merged layer flow on the spherical nose of a space vehicle (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Hypersonic merged layer flow on the forepart of a spherical surface of a space vehicle has been investigated on the basis of the full steady-state Navier-Stokes equations using slip and temperature jump boundary conditions at the surface and free-stream conditions far from the surface. The shockwave-like structure was determined as part of the computations. Using an equivalent body concept, computations were carried out under conditions that the Aeroassist Flight Experiment (AFE) Vehicle would encounter at 15 and 20 seconds in its flight path. Emphasis was placed on understanding the basic nature of the flow structure under low density conditions. Particular attention was paid to the understanding of the structure of the outer shockwave-like region as the fluid expands around the sphere. Plots were drawn for flow profiles and surface characteristics to understand the role of dissipation processes in the merged layer of the spherical nose of the vehicle.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2692
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  • 114
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    Numerical simulation of hypersonic flow over an aeroassist flight experiment vehicle (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Solutions of wind-tunnel and entry-flight flow around the vehicle are obtained from the Navier-Stokes equations coupled with the chemical species continuity equations if needed. The time-iterative method employs several techniques: shock fitting, chemistry-split ADI and an algebraic grid in conformal spherical-polar space. Sensitivities of the results to numerical parameters and to frozen, equilibrium and finite rate reactions are investigated in the forebody computation. Quantitative results are obtained for the shock layer and the near wake for the entire vehicle corresponding to both ground test and flight conditions. Complex flow characteristics are analyzed on the basis of the complete flowfield over the aerobrake and simplified afterbodies. The method is stable and cost effective, and has yielded shock locations and wall pressure distributions which are in good agreement with wind-tunnel data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2675
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  • 115
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    Porous wind tunnel corrections for counterrotation propeller testing (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Wind tunnel interference corrections have direct impact on measured propeller efficiency. A systematic series of wind tunnel tests was done in the porous-wall NASA Lewis 8- by 6-Foot Wind Tunnel to determine the wind tunnel interference corrections to the NASA Lewis counterrotation propeller test data. The test results were compared with calculations from a potential flow code to determine the interference corrections. At a Mach number of 0.8, the interference corrections resulted in a -0.008 Mach number correction which reduced the counterrotation propeller net efficiency data by 0.46 percent at the reduced Mach number. Additional wind tunnel tests were done to measure the effect of propeller thrust on wind tunnel wall interference. No wall interference corrections due to propeller thrust were found necessary for the high speed counterrotation propeller data obtained in the porous wall NASA Lewis 8- by 6-Foot Wind Tunnel.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2055
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  • 116
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    The computation of radiation from nonequilibrium hypersonic flows (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The results of the solution of the equations that describe a hypersonic ionized flow about an elliptically blunted cone are presented. The flow conditions correspond to those of the proposed Aeroassist Flight Experiment (AFE) vehicle at altitudes between the perigee at 78 km and the approximate limit of the continuum regime at 90 km. For the free-stream velocities of interest, about 9 km/sec, the flowfield is out of thermo-chemical equilibrium, electronically excited, ionized and radiating. The gas consists of eight-chemical species including free electrons. The thermal state of the gas is modeled with a translational-rotational temperature, four vibrational temperatures for the diatomic species and an electron-electronic temperature. The electronic excitation of molecules is included. The nonequilibrium air radiation from each fluid element is computed and the radiative heat flux at the body surface is determined. The stagnation point radiative heating result agrees with previous calculations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2678
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  • 117
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    Vectorization of a particle simulation method for hypersonic rarefied flow (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An efficient particle simulation technique for hypersonic rarefied flows is presented at an algorithmic and implementation level. The implementation is for a vector computer architecture, specifically the Cray-2. The method models an ideal diatomic Maxwell molecule with three translational and two rotational degrees of freedom. Algorithms are designed specifically for compatibility with fine grain parallelism by reducing the number of data dependencies in the computation. By insisting on this compatibility, the method is capable of performing simulation on a much larger scale than previously possible. A two-dimensional simulation of supersonic flow over a wedge is carried out for the near-continuum limit where the gas is in equilibrium and the ideal solution can be used as a check on the accuracy of the gas model employed in the method. Also, a three-dimensional, Mach 8, rarefied flow about a finite-span flat plate at a 45 degree angle of attack was simulated. It utilized over 10 to the 7th particles carried through 400 discrete time steps in less than one hour of Cray-2 CPU time. This problem was chosen to exhibit the capability of the method in handling a large number of particles and a true three-dimensional geometry.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2735
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  • 118
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    Viscous shock-layer solutions with nonequilibrium chemistry for hypersonic flows past slender bodies (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Laminar nonequilibrium heat transfer to slender vehicles is discussed, with heating-rate results presented as a ratio of the noncatalytic to the corresponding fully catalytic value to illustrate the maximum potential for a heating reduction in dissociated nonequilibrium flow at a given flight condition. Larger blunted cone half-angles are shown to produce the most significant nonequilibrium effects at distances beyond 100 nose radii, except in the fore-cone region. Increasing nose bluntness is found to produce large reductions in the ratio for the smaller cone angles at relatively large downstream surface lengths. It is noted that the nose radius and freestream density are not independent scaling parameters in nonequilibrium flow.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2709
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  • 119
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    Effects of surface pressures and streamline metrics on the calculation of laminar heating rates (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The effect of streamline geometry and pressure distributions on surface heating rates is examined for slender, spherically blunted cones. The modifications to the approximate aeroheating code include a curve fit of pressures computed by an Euler solution over a range of Mach numbers and cone angles. The streamline geometry is then found using the surface pressures and inviscid surface properties. Previously, streamlines were determined using the inviscid properties at the edge of the boundary layer when accounting for the effects of entropy-layer swallowing. Streamline calculations are now based on inviscid surface conditions rather than boundary-layer edge properties. However, the heating rates are calculated using inviscid properties at the edge of the boundary layer. Resulting heating rates compare favorably with solutions from the viscous-shock-layer equations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2708
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  • 120
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    Radiation heat transfer in the wake of a hypersonic vehicle (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Radiation heat transfer (RHT) from the wake of a hypersonic vehicle to its afterbody is evaluated from Gnoffo and Greene's (1987) calculated wake flowfield and the radiative properties of ionized high-temperature air with the calculated nonequilibrium composition. The 4.2-m aeroassisted flight experiment at an altitude of 75 km and velocity of 8900 m/s causes a 0.1-m-thick layer initially at T = 10,000 K and P = 1 kN/sq m to separate from the shoulder of the forebody heat shield and spread aft to form a wake at approximately T = 5000 K and P = 20 N/sq m. Gas in the separated flow region at approximately T = 3000 K and P = 10 N/sq m, recirculates about the afterbody. It is shown that the radiating layer, recirculating gases, and wake are optically thin for purposes of making engineering RHT calculations. Directional, spectral, and spatial variations of the radiation incident upon the afterbody are presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2634
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  • 121
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    An upwind parabolized Navier-Stokes code for chemically reacting flows (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A new upwind, parabolized Navier-Stokes (PNS) code has been developed to compute the hypersonic, viscous, chemically reacting flow around two-dimensional or axisymmetric bodies. The new code is an extension of the upwind (perfect gas) PNS code of Lawrence et al. (1986). The upwind algorithm is based on Roe's flux-difference splitting scheme which has been modified to account for real gas effects. The algorithm solves the gas dynamic and species continuity equations in a 'loosely' coupled manner. The new code has been validated by computing the laminar flow (at free stream Mach number 25) of chemically reacting air over a wedge and a cone. The results of these computations are compared with the results from a centrally-differenced, fully coupled, nonequilibrium PNS code. The agreement is excellent, except in the vicinity of the shock wave where the present code exhibits superior shock capturing capabilities.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2614
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  • 122
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    The numerical simulation of transonic separated flow about the complete F-16A (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The thin-layer, Reynolds-averaged, Navier-Stokes equations are used to simulate the transonic viscous flow about the complete F-16A fighter aircraft. These computations demonstrate how computational fluid dynamics (CFD) can be used to simulate turbulent viscous flow about realistic aircraft geometries. A zonal grid approach is used to provide adequate viscous grid clustering on all aircraft surfaces. Zonal grids extend inside the F-16A inlet and up to the compressor face while power on conditions are modeled by employing a zonal grid extending from the exhaust nozzle to the far field. A simple solution adaptive grid procedure is used on the wing surface and good agreement with experimental data is obtained. Computations for the F-16A in side slip are also presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2506
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  • 123
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    Recent advances in transonic computational aeroelasticity (1988)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A transonic unsteady aerodynamic and aeroelasticity code called CAP-TSD was developed for application to realistic aircraft configurations. The code permits the calculation of steady and unsteady flows about complete aircraft configurations for aeroelastic analysis in the flutter critical transonic speed range. The CAP-TSD code uses a time accurate approximate factorization algorithm for solution of the unsteady transonic small disturbance potential equation. An overview is given of the CAP-TSD code development effort and results are presented which demonstrate various capabilities of the code. Calculations are presented for several configurations including the General Dynamics 1/9 scale F-16 aircraft model and the ONERA M6 wing. Calculations are also presented from a flutter analysis of a 45 deg sweptback wing which agrees well with the experimental data. Descriptions are presented of the CAP-TSD code and algorithm details along with results and comparisons which demonstrate these recent developments in transonic computational aeroelasticity.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100663 , NAS 1.15:100663
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  • 124
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    The present status and the future of missile aerodynamics (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Some recent developments in the state of the art in missile aerodynamics are reviewed. Among the subjects covered are: (1) tri-service/NASA data base, (2) wing-body interference, (3) nonlinear controls, (4) hypersonic transition, (5) vortex interference, (6) airbreathers, supersonic inlets, (7) store separation problems, (8) correlation of missile data, (9) CFD codes for complete configurations, (10) engineering prediction methods, and (11) future configurations. Suggestions are made for future research and development to advance the state of the art of missile aerodynamics.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100063 , A-87289 , NAS 1.15:100063
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  • 125
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    Three-dimensional Navier-Stokes simulations of turbine rotor-stator interaction (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Fluid flows within turbomachinery tend to be extremely complex in nature. Understanding such flows is crucial to improving current designs of turbomachinery. The computational approach can be used to great advantage in understanding flows in turbomachinery. A finite difference, unsteady, thin layer, Navier-Stokes approach to calculating the flow within an axial turbine stage is presented. The relative motion between the stator and rotor airfoils is made possible with the use of patched grids that move relative to each other. The calculation includes endwall and tip leakage effects. An introduction to the rotor-stator problem and sample results in the form of time averaged surface pressures are presented. The numerical data are compared with experimental data and the agreement between the two is found to be good.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100081 , A-88106 , NAS 1.15:100081
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  • 126
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    Analysis and Optimization of Truncated Scarf Nozzles Subject to External Flow Conditions (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Results of a calculation of an optimized truncated scarfed nozzle were compared. The truncated scarfed nozzle was designed for an exit Mach number of 6.0, i.e., the Mach number at the last nozzle characteristic is 6.0, with an external flow Mach number of 5.0. The nozzle was designed by the Rao method for optimum thrust nozzles modified for 2-D flow and truncated scarfed nozzle applications. This design was analyzed using a shock-fitting method for 2-D supersonic flows. Excellent agreement was achieved between the design and analysis. Truncation of the lower nozzle wall (cowl) revealed that there is an optimum length for truncating the cowl without degrading the nozzle performance. Truncation of the nozzle cowl past this optimal length should be analyzed in trade-off studies for thrust loss versus gross vehicle weight. Plots of the oblique shock wave equations were also identified which will allow computation of slip line angle, dynamic pressure coefficient, or ambient Mach number for various specific heat ratios.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100955 , E-4146 , NAS 1.15:100955
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  • 127
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    Adaptive mesh generation for viscous flows using Delaunay triangulation (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A method for generating an unstructured triangular mesh in two dimensions, suitable for computing high Reynolds number flows over arbitrary configurations is presented. The method is based on a Delaunay triangulation, which is performed in a locally stretched space, in order to obtain very high aspect ratio triangles in the boundary layer and the wake regions. It is shown how the method can be coupled with an unstructured Navier-Stokes solver to produce a solution adaptive mesh generation procedure for viscous flows.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181699 , ICASE-88-47 , NAS 1.26:181699
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    Space shuttle phase B wind tunnel model and test information. Volume 1: Booster configuration (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA Centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include the booster, the orbiter and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configurations include booster and orbiter components in various stacked and tandem combinations. This is Volume 1 (Part 1) of the report -- Booster Configuration.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178414-VOL-1-PT-1 , NAS 1.26:178414-VOL-1-PT-1 , DMS-DB-02-VOL-1-PT-1
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    Accurate solutions, parameter studies and comparisons for the Euler and potential flow equations (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Parameter studies are conducted using the Euler and potential flow equation models for unsteady and steady flows in both two and three dimensions. The Euler code is an implicit, upwind, finite volume code which uses the Van Leer method of flux-vector-splitting which has been recently extended for use on dynamic meshes and maintain all the properties of the original splitting. The potential flow code is an implicit, finite difference method for solving the transonic small disturbance equations and incorporates both entropy and vorticity corrections into the solution procedures thereby extending its applicability into regimes where shock strength normally precludes its use. Parameter studies resulting in benchmark type calculations include the effects of spatial and temporal refinement, spatial order of accuracy, far field boundary conditions for steady flow, frequency of oscillation, and the use of subiterations at each time step to reduce linearization and factorization errors. Comparisons between Euler and potential flows results are made as well as with experimental data where available.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100664 , NAS 1.15:100664
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  • 130
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    Space shuttle phase B wind tunnel model and test information. Volume 2: Orbiter configuration (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Archived wind tunnel test data are available for flyback booster or other alternate recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquiredin the Phase B development have been compiled into a database and are available for applying to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include the booster, the orbiter, and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide, and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configration types include booster and orbiter components in various stacked and tandom combinations. The digital database consists of 220 files of data containing basic tunnel recorded data.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178415-VOL-2-PT-1 , NAS 1.26:178415-VOL-2-PT-1 , DMS-DB-02-VOL-2-PT-1
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  • 131
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    Numerical computation of space shuttle orbiter flow field (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A new parabolized Navier-Stokes (PNS) code has been developed to compute the hypersonic, viscous chemically reacting flow fields around 3-D bodies. The flow medium is assumed to be a multicomponent mixture of thermally perfect but calorically imperfect gases. The new PNS code solves the gas dynamic and species conservation equations in a coupled manner using a noniterative, implicit, approximately factored, finite difference algorithm. The space-marching method is made well-posed by special treatment of the streamwise pressure gradient term. The code has been used to compute hypersonic laminar flow of chemically reacting air over cones at angle of attack. The results of the computations are compared with the results of reacting boundary-layer computations and show excellent agreement.
    Keywords: AERODYNAMICS
    Type: NASA-CR-183193 , NAS 1.26:183193 , ISU-ERI-AMES-89403 , CFD-19
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  • 132
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    Program user's manual for an unsteady helicopter rotor-fuselage aerodynamic analysis (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The Rotor-Fuselage Analysis is a method of calculating the aerodynamic reaction between a helicopter rotor and fuselage. This manual describes the structure and operation of the computer programs that make up the Rotor-Fuselage Analysis, programs which prepare the input and programs which display the output.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181701 , NAS 1.26:181701 , UTRC/R88-956977
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  • 133
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    Three-dimensional zonal grids about arbitrary shapes by Poisson's equation (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A method for generating 3-D finite difference grids about or within arbitrary shapes is presented. The 3-D Poisson equations are solved numerically, with values for the inhomogeneous terms found automatically by the algorithm. Those inhomogeneous terms have the effect near boundaries of reducing cell skewness and imposing arbitrary cell height. The method allows the region of interest to be divided into zones (blocks), allowing the method to be applicable to almost any physical domain. A FORTRAN program called 3DGRAPE has been written to implement the algorithm. Lastly, a method for redistributing grid points along lines normal to boundaries will be described.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101018 , A-88258 , NAS 1.15:101018
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  • 134
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    An improved version of NCOREL: A computer program for 3-D nonlinear supersonic potential flow computations (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A computer code called NCOREL (for Nonconical Relaxation) has been developed to solve for supersonic full potential flows over complex geometries. The method first solves for the conical at the apex and then marches downstream in a spherical coordinate system. Implicit relaxation techniques are used to numerically solve the full potential equation at each subsequent crossflow plane. Many improvements have been made to the original code including more reliable numerics for computing wing-body flows with multiple embedded shocks, inlet flow through simulation, wake model and entropy corrections. Line relaxation or approximate factorization schemes are optionally available. Improved internal grid generation using analytic conformal mappings, supported by a simple geometric Harris wave drag input that was originally developed for panel methods and internal geometry package are some of the new features.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4165 , NAS 1.26:4165 , RE-744
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  • 135
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    Euler equation computations for the flow over a hovering helicopter rotor (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A numerical solution technique is developed for computing the flow field around an isolated helicopter rotor in hover. The flow is governed by the compressible Euler equations which are integrated using a finite volume approach. The Euler equations are coupled to a free wake model of the rotary wing vortical wake. This wake model is incorporated into the finite volume solver using a prescribed flow, or perturbation, technique which eliminates the numerical diffusion of vorticity due to the artificial viscosity of the scheme. The work is divided into three major parts: (1) comparisons of Euler solutions to experimental data for the flow around isolated wings show good agreement with the surface pressures, but poor agreement with the vortical wake structure; (2) the perturbation method is developed and used to compute the interaction of a streamwise vortex with a semispan wing. The rapid diffusion of the vortex when only the basic Euler solver is used is illustrated, and excellent agreement with experimental section lift coefficients is demonstrated when using the perturbation approach; and (3) the free wake solution technique is described and the coupling of the wake to the Euler solver for an isolated rotor is presented. Comparisons with experimental blade load data for several cases show good agreement, with discrepancies largely attributable to the neglect of viscous effects. The computed wake geometries agree less well with experiment, the primary difference being that too rapid a wake contraction is predicted for all the cases.
    Keywords: AERODYNAMICS
    Type: NASA-CR-177493 , NAS 1.26:177493
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  • 136
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    Dual adaptive control: Design principles and applications (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The design of an actively adaptive dual controller based on an approximation of the stochastic dynamic programming equation for a multi-step horizon is presented. A dual controller that can enhance identification of the system while controlling it at the same time is derived for multi-dimensional problems. This dual controller uses sensitivity functions of the expected future cost with respect to the parameter uncertainties. A passively adaptive cautious controller and the actively adaptive dual controller are examined. In many instances, the cautious controller is seen to turn off while the latter avoids the turn-off of the control and the slow convergence of the parameter estimates, characteristic of the cautious controller. The algorithms have been applied to a multi-variable static model which represents a simplified linear version of the relationship between the vibration output and the higher harmonic control input for a helicopter. Monte Carlo comparisons based on parametric and nonparametric statistical analysis indicate the superiority of the dual controller over the baseline controller.
    Keywords: AERODYNAMICS
    Type: NASA-CR-177485 , NAS 1.26:177485
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  • 137
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    Investigation of oscillating cascade aerodynamics by an experimental influence coefficient technique (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Fundamental experiments are performed in the NASA Lewis Transonic Oscillating Cascade Facility to investigate the torsion mode unsteady aerodynamics of a biconvex airfoil cascade at realistic values of the reduced frequency for all interblade phase angles at a specified mean flow condition. In particular, an unsteady aerodynamic influence coefficient technique is developed and utilized in which only one airfoil in the cascade is oscillated at a time and the resulting airfoil surface unsteady pressure distribution measured on one dynamically instrumented airfoil. The unsteady aerodynamics of an equivalent cascade with all airfoils oscillating at a specified interblade phase angle are then determined through a vector summation of these data. These influence coefficient determined oscillation cascade data are correlated with data obtained in this cascade with all airfoils oscillating at several interblade phase angle values. The influence coefficients are then utilized to determine the unsteady aerodynamics of the cascade for all interblade phase angles, with these unique data subsequently correlated with predictions from a linearized unsteady cascade model.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101313 , E-4308 , NAS 1.15:101313 , AIAA PAPER 88-2815
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  • 138
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    Method to predict external store carriage characteristics at transonic speeds (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Development of a computational method for prediction of external store carriage characteristics at transonic speeds is described. The geometric flexibility required for treatment of pylon-mounted stores is achieved by computing finite difference solutions on a five-level embedded grid arrangement. A completely automated grid generation procedure facilitates applications. Store modeling capability consists of bodies of revolution with multiple fore and aft fins. A body-conforming grid improves the accuracy of the computed store body flow field. A nonlinear relaxation scheme developed specifically for modified transonic small disturbance flow equations enhances the method's numerical stability and accuracy. As a result, treatment of lower aspect ratio, more highly swept and tapered wings is possible. A limited supersonic freestream capability is also provided. Pressure, load distribution, and force/moment correlations show good agreement with experimental data for several test cases. A detailed computer program description for the Transonic Store Carriage Loads Prediction (TSCLP) Code is included.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4170 , NAS 1.26:4170
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  • 139
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    VSAERO analysis of tip planforms for the free-tip rotor (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The results of a numerical analysis of two interacting lifting surfaces separated in the spanwise direction by a narrow gap are presented. The configuration consists of a semispan wing with the last 32 percent of the span structurally separated from the inboard section. The angle of attack of the outboard section is set independently from that of the inboard section. In the present study, the three-dimensional panel code VSAERO is used to perform the analysis. Computed values of tip surface lift and pitching moment coefficients are correlated with experimental data to determine the proper approach to model the gap region between the surfaces. Pitching moment data for various tip planforms are also presented to show how the variation of tip pitching moment with angle of attack may be increased easily in incompressible flow. Calculated three-dimensional characteristics in compressible flow at Mach numbers of 0.5 and 0.7 are presented for new tip planform designs. An analysis of sectional aerodynamic center shift as a function of Mach number is also included for a representative tip planform. It is also shown that the induced drag of the tip surface is reduced for negative incidence angles relative to the inboard section. The results indicate that this local drag reduction overcomes the associated increase in wing induced drag at high wing lift coefficients.
    Keywords: AERODYNAMICS
    Type: NASA-CR-177487 , NAS 1.26:177487
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  • 140
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    A computational analysis of under-expanded jets in the hypersonic regime (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Underexpanded axisymmetric jets are studied numerically using a full Navier-Stokes solver. Emphasis has been given to supersonic and hypersonic jets in supersonic and hypersonic ambient flows, a phenomenon previously overlooked. It is demonstrated that the shear layers and shock patterns in a jet plume can be captured without complicated viscous/inviscid and subsonic/supersonic coupling schemes. In addition, a supersonic pressure relief effect has been identified for underexpanded jets in supersonic ambient flows. While it is well known that an underexpanded jet in a quiescent ambience (or subsonic ambience) contains multiple shock cells, the present study shows that because of the supersonic pressure relief effect, an underexpanded jet in a supersonic or hypersonic ambience contains only one major shock cell.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101319 , E-4317 , NAS 1.15:101319
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  • 141
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    Space shuttle phase B wind tunnel model and test information. Volume 2: Orbiter configuration (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a data base and are available for applying to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Data Base is structured by vehicle component and configuration type. Basic components include the booster, the orbiter, and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retro-glide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configuration types include booster and orbiter components in various stacked and tandem combinations.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178415-VOL-2-PT-2 , NAS 1.26:178415-VOL-2-PT-2 , DMS-DB-02-VOL-2
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  • 142
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    The determination of the topological structure of skin friction lines on a rectangular wing-body combination (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A short tutorial in the application of topological ideas to the intepretation of oil flow patterns is presented. Topological concepts such as critical points, phase portraits, topological stability, and indexing are discussed. These concepts are used in an ordered procedure to construct phase portraits of skin friction lines with oil flow patterns for a wing-body combination and two angles of attack. The relationship between the skin friction phase portrait and planar cuts of the velocity field is also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4168 , NAS 1.26:4168
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  • 143
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    Vortex methods for separated flows (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The numerical solution of the Euler or Navier-Stokes equations by Lagrangian vortex methods is discussed. The mathematical background is presented in an elementary fashion and includes the relationship with traditional point-vortex studies, the convergence to smooth solutions of the Euler equations, and the essential differences between two- and three-dimensional cases. The difficulties in extending the method to viscous or compressible flows are explained. The overlap with the excellent review articles available is kept to a minimum and more emphasis is placed on the area of expertise, namely two-dimensional flows around bluff bodies. When solid walls are present, complete mathematical models are not available and a more heuristic attitude must be adopted. The imposition of inviscid and viscous boundary conditions without conformal mappings or image vortices and the creation of vorticity along solid walls are examined in detail. Methods for boundary-layer treatment and the question of the Kutta condition are discussed. Practical aspects and tips helpful in creating a method that really works are explained. The topics include the robustness of the method and the assessment of accuracy, vortex-core profiles, timemarching schemes, numerical dissipation, and efficient programming. Calculations of flows past streamlined or bluff bodies are used as examples when appropriate.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100068 , A-88097 , NAS 1.15:100068
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    A diagonally inverted LU implicit multigrid scheme (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A new Diagonally Inverted LU Implicit scheme is developed within the framework of the multigrid method for the 3-D unsteady Euler equations. The matrix systems that are to be inverted in the LU scheme are treated by local diagonalizing transformations that decouple them into systems of scalar equations. Unlike the Diagonalized ADI method, the time accuracy of the LU scheme is not reduced since the diagonalization procedure does not destroy time conservation. Even more importantly, this diagonalization significantly reduces the computational effort required to solve the LU approximation and therefore transforms it into a more efficient method of numerically solving the 3-D Euler equations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100911 , E-4163 , NAS 1.15:100911 , AIAA PAPER 88-3567
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  • 145
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    High speed inlet calculations with real gas effects (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A 2-D steady-state Navier-Stokes solver has been upgraded to include the effects of frozen and equilibrium air chemistry for applications to high speed flight vehicles. To provide a computationally economical first order approximation to the high temperature physics, variable thermodynamic data is used for the chemically frozen mode to allow for a variation with temperature of the air specific heats and enthalpy. For calculations involving air in chemical equilibrium, a specially modified version of the NASA Lewis Chemical Equilibrium Code, CEC, is used to compute the chemical composition and resultant thermochemical properties. The upgraded solver is demonstrated by comparing results from calorically perfect (C sub p=constant), thermally perfect (frozen) and equilibrium air calculations for a variety of geometries, and flight Mach numbers.
    Keywords: AERODYNAMICS
    Type: NASA-CR-182167 , E-4287 , NAS 1.26:182167 , AIAA PAPER 88-3076
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  • 146
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    Numerical study of delta wing leading edge blowing (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Spanwise and tangential leading edge blowing as a means of controlling the position and strength of the leading edge vortices are studied by numerical solution of the three-dimensional Navier-Stokes equations. The leading edge jet is simulated by defining a permeable boundary, corresponding to the jet slot, where suitable boundary conditions are implemented. Numerical results are shown to compare favorably with experimental measurements. It is found that the use of spanwise leading edge blowing at moderate angle of attack magnifies the size and strength of the leading edge vortices, and moves the vortex cores outboard and upward. The increase in lift primarily comes from the greater nonlinear vortex lift. However, spanwise blowing causes earlier vortex breakdown, thus decreasing the stall angle. The effects of tangential blowing at low to moderate angles of attack tend to reduce the pressure peaks associated with leading edge vortices and to increase the suction peak around the leading edge, so that the integrated value of the surface pressure remains about the same. Tangential leading edge blowing in post-stall conditions is shown to re-establish vortical flow and delay vortex bursting, thus increasing C sub L sub max and stall angle.
    Keywords: AERODYNAMICS
    Type: NASA-CR-183101 , NAS 1.26:183101 , JIAA-TR-86
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  • 147
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    Improved numerical methods for turbulent viscous recirculating flows (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The performance of discrete methods for the prediction of fluid flows can be enhanced by improving the convergence rate of solvers and by increasing the accuracy of the discrete representation of the equations of motion. This report evaluates the gains in solver performance that are available when various acceleration methods are applied. Various discretizations are also examined and two are recommended because of their accuracy and robustness. Insertion of the improved discretization and solver accelerator into a TEACH mode, that has been widely applied to combustor flows, illustrates the substantial gains to be achieved.
    Keywords: AERODYNAMICS
    Type: NASA-CR-180852 , NAS 1.26:180852
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  • 148
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    Parametric flutter studies of an arrow-wing configuration: Some early results (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Some early experimental results from a combined experimental and analytical study being conducted at NASA-Langley of the transonic flutter characterisitics of a generic arrow wing configuration are presented. The planned study includes the parametric variation of a variety of structural and geometric characteristics. Presented here are flutter results of the basic arrow wing, for the basic wing with the addition of two simulated lower-surface-mounted engine nacelles, and for the basic wing with the addition of both the fin and the engine nacelles.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100608 , NAS 1.15:100608
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  • 149
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    Calibration of averaging total pressure flight wake rake and natural-laminar-flow airfoil drag certification (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An averaging total pressure wake rake used by the Cessna Aircraft Company in flight tests of a modified 210 airplane with a laminar flow wing was calibrated in wind tunnel tests against a five-tube pressure probe. The model generating the wake was a full-scale model of the Cessna airplane wing. Indications of drag trends were the same for both instruments.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181630 , NAS 1.26:181630 , AR-87-1
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    Empty test section streamlining of the transonic self-streamlining wind tunnel fitted with new walls (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The original flexible top and bottom walls of the Transonic Self-Streamlining Wind Tunnel (TSWT), at the University of Southampton, have been replaced with new walls featuring a larger number of static pressure tappings and detailed mechanical improvements. This report describes the streamling method, results, and conclusions of a series of tests aimed at defining sets of aerodynamically straight wall contours for the new flexible walls. This procedure is a necessary prelude to model testing. The quality of data obtained compares favorably with the aerodynamically straight data obtained with the old walls. No operational difficulties were experienced with the new walls.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181680 , NAS 1.26:181680 , AASU-MEMO-86/10
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  • 151
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    A low speed wind tunnel investigation of Reynolds number effects on a 60-deg swept wing configuration with leading and trailing edge flaps (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A low-speed wind tunnel test was performed to investigate Reynolds number effects on the aerodynamic characteristics of a supersonic cruise wing concept model with a 60-deg swept wing incorporating leading-edge and trailing-edge flap deflections. The Reynolds number ranged from 0.3 to 1.6 x 10 to the 6th, and corresponding Mach numbers from .05 to 0.3. The objective was to define a threshold Reynolds number above which the flap aerodynamics basically remained unchanged, and also to generate a data base useful for validating theoretical predictions for the Reynolds number effects on flap performance. This report documents the test procedures used and the basic data acquired in the investigation.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181684 , NAS 1.26:181684
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  • 152
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    The experimental verification of wall movement influence coefficients for an adaptive walled test section (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Flexible walled wind tunnels have for some time been used to reduce wall interference effects at the model. A necessary part of the 3-D wall adjustment strategy being developed for the Transonic Self-Streamlining Wind Tunnel (TSWT) of Southampton University is the use of influence coefficients. The influence of a wall bump on the centerline flow in TSWT has been calculated theoretically using a streamline curvature program. This report details the experimental verification of these influence coefficients and concludes that it is valid to use the theoretically determined values in 3-D model testing.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181681 , NAS 1.26:181681 , AASU-87/4
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  • 153
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    Correlations of velocity and temperature fluctuations in the stagnation-point flow of circular cylinder in turbulent flow (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Boundary layer flow and turbulence transport analyses to study the influence of the free-stream turbulence on the surface heat transfer rate and the skin friction around the stagnation point of a circular cylinder in a turbulent flow are presented. The analyses are formulated with the turbulent boundary layer equations, the Reynolds stress transport equations and the k - epsilon two-equation turbulence modeling. The analyses are used to calculate the time-averaged turbulence double correlations, the mean flow properties, the surface heat transfer rate and the skin friction with an isotropic turbulence in the freestream. The analytical results are described and compared with the existing experimental measurements. Depending on the free-stream turbulence properties, the turbulence kinetic energy can increase or decrease as the flow moves toward the surface. However, the turbulence kinetic energy induces large Reynolds normal stresses at the boundary layer edge. The Reynolds normal stresses change the boundary layer profiles of the time-averaged double correlations of the velocity and temperature fluctuations, the surface heat transfer rate and the skin friction. The free-stream turbulence dissipation rate can affect the stagnation-point heat transfer rate but the influence of the free-stream temperature fluctuation on the heat transfer rate is insignificant.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100930 , E-4201 , NAS 1.15:100930
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  • 154
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    On the validation of a code and a turbulence model appropriate to circulation control airfoils (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A computer code for calculating flow about a circulation control airfoil within a wind tunnel test section has been developed. This code is being validated for eventual use as an aid to design such airfoils. The concept of code validation being used is explained. The initial stages of the process have been accomplished. The present code has been applied to a low-subsonic, 2-D flow about a circulation control airfoil for which extensive data exist. Two basic turbulence models and variants thereof have been successfully introduced into the algorithm, the Baldwin-Lomax algebraic and the Jones-Launder two-equation models of turbulence. The variants include adding a history of the jet development for the algebraic model and adding streamwise curvature effects for both models. Numerical difficulties and difficulties in the validation process are discussed. Turbulence model and code improvements to proceed with the validation process are also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100090 , A-88127 , NAS 1.15:100090
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  • 155
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    Aerofoil testing in a self-streamlining flexible walled wind tunnel (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Two-dimensional self-streamlining flexible walled test sections eliminate, as far as experimentally possible, the top and bottom wall interference effects in transonic airfoil testing. The test section sidewalls are rigid, while the impervious top and bottom walls are flexible and contoured to streamline shapes by a system of jacks, without reference to the airfoil model. The concept of wall contouring to eliminate or minimize test section boundary interference in 2-D testing was first demonstrated by NPL in England during the early 40's. The transonic streamlining strategy proposed, developed and used by NPL has been compared with several modern strategies. The NPL strategy has proved to be surprisingly good at providing a wall interference-free test environment, giving model performance indistinguishable from that obtained using the modern strategies over a wide range of test conditions. In all previous investigations the achievement of wall streamlining in flexible walled test sections has been limited to test sections up to those resulting in the model's shock just extending to a streamlined wall. This work however, has also successfully demonstrated the feasibility of 2-D wall streamlining at test conditions where both model shocks have reached and penetrated through their respective flexible walls. Appropriate streamlining procedures have been established and are uncomplicated, enabling flexible walled test sections to cope easily with these high transonic flows.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4128 , NAS 1.26:4128
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  • 156
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    Inflow measurements made with a laser velocimeter on a helicopter model in forward flight. Volume 4: Tapered planform blades at an advance ratio of 0.15 (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted in the 14- by 22-Foot Subsonic Tunnel at NASA Langley Research Center to measure the inflow into the scale model helicopter rotor in forward flight (mu sub infinity = 0.15). The measurements were made with a two-component Laser Velocimeter (LV) one chord above the plane formed by the path of the rotor tips (tip path plane). A conditional sampling technique was employed to determine the position of the rotor at the time that each velocity measurement was made so that the azimuthal fluctuations in velocity could be determined. Measurements were made at a total of 146 separate locations in order to clearly define the inflow character. This data is presented herein without analysis. In order to increase the availability of the resulting data, both the mean and azimuthally dependent values are included as part of this report on two 5.25 inch floppy disks in MS-DOS format.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100544 , AVSCOM-TM-88-B-007 , NAS 1.15:100544
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  • 157
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    Laser velocimeter measurements in a wing-fuselage type juncture (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A single axis, five beam, three component laser velocimeter system was used in a juncture flow experiment. A description of the seeding system developed for and used in this experiment is given. The performanace of the LV system was evaluated, and some of the problems associated with it were identified. Satisfactory results were obtained in the juncture flow experiments using this LV system.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100588 , NAS 1.15:100588
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  • 158
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    Inflow measurement made with a laser velocimeter on a helicopter model in forward flight. Volume 3: Rectangular planform blades at an advance ratio of 0.30 (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted in the 14- by 22-Foot Subsonic Tunnel at NASA Langley Research Center to measure the inflow into a scale model helicopter rotor in forward flight (micron sub infinity = 0.30). The measurements were made with a two component Laser Velocimeter (LV) one chord above the plane formed by the path of the rotor tips (tip path plane). A conditional sampling technique was employed to determine the azimuthal position of the rotor at the time that each velocity measurement was made so that the azimuthal fluctuations in velocity could be determined. Measurements were made at a total of 180 separate locations in order to clearly define the inflow character. These data are presented without analysis.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100543 , NAS 1.15:100543 , AVSCOM-TM-88-B-006
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  • 159
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    Application of Hybrid Laminar Flow Control to Global Range Military Transport Aircraft (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A study was conducted to evaluate the application of hybrid laminar flow control (HLFC) to global range military transport aircraft. The global mission included the capability to transport 132,500 pounds of payload 6500 nautical miles, land and deliver the payload and without refueling return 6500 nautical miles to a friendly airbase. The preliminary design studies show significant performance benefits obtained for the HLFC aircraft as compared to counterpart turbulent flow aircraft. The study results at M=0.77 show that the largest benefits of HLFC are obtained with a high wing with engines on the wing configuration. As compared with the turbulent flow baseline aircraft, the high wing HLFC aircraft shows 17 percent reduction in fuel burned, 19.2 percent increase in lift-to-drag ratio, an insignificant increase in operating weight, and a 7.4 percent reduction in gross weight.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181638 , NAS 1.26:181638
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  • 160
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    Hot gas ingestion testing of an advanced STOVL concept in the NASA Lewis 9- by 15-foot Low Speed Wind Tunnel with flow visualization (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Advanced Short Takeoff/Vertical Landing (STOVL) aircraft capable of operating from remote sites, damaged runways, and small air capable ships are being pursued for deployment around the turn of the century. To achieve this goal, it is important that the technologies critical to this unique class of aircraft be developed. Recognizing this need, NASA Lewis Research Center, McDonnell Douglas Aircraft, and DARPA defined a cooperative program for testing in the NASA Lewis 9- by 15-foot Low Speed Wind Tunnel (LSWT) to establish a database for hot gas ingestion, one of the technologies critical to STOVL. Results from a test program are presented along with a discussion of the facility modifications allowing this type of testing at modal scale. These modifications to the tunnel include a novel ground plane, an elaborate model support which included 4 degrees of freedom, heated high pressure air for nozzle flow, a suction system exhaust for inlet flow, and tunnel sidewall modifications. Several flow visualization techniques were employed including water mist in the nozzle flows and tufts on the ground plane. Headwind (free-stream) velocity was varied from 8 to 23 knots.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3025
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  • 161
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    Euler analysis of a swirl recovery vane design for use with an advanced single-rotation propfan (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Recent work has demonstrated the propulsive efficiency improvement available from single- and counter-rotation propfans as compared with current technology high bypass ratio turbofans. The concept known as swirl recovery vanes (SRV) is examined through the use of a 3-D Euler code. At high speed cruise conditions, the SRV can improve the efficiency level of a single-rotation propfan, but a concern is to have adequate hub choke margin. The SRV was designed with 2-D methods and was predicted to have hub choking at Mach 0.8 cruise. The 3-D Euler analysis properly accounts for sweep effects and 3-D relief, and predicts that at cruise the SRV will recover roughly 5 percent of the 10 percent efficiency loss due to swirl and have a good hub choke margin.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3152
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  • 162
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    Three dimensional simulation of an underexpanded jet interacting with a supersonic cross flow (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A new three-dimensional computational fluid dynamics (CFD) code has been developed to simulate the flow fields of an underexpanded jet transversely injected into supersonic air stream inside the combustors of ramjets and scramjets. The code employs an implicit finite volume, lower-upper (LU) time marching method to solve the complete three-dimensional Navier-Stokes equations in a fully-coupled and very efficient manner. Results clearly depict the flow characteristics, including the shock structure, separated flow regions around the injector, and the wake flow in the lee of the injector.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3181
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  • 163
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    High speed inlet calculations with real gas effects (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A 2-D steady-state Navier-Stokes solver has been upgraded to include the effects of frozen and equilibrium air chemistry for applications to high speed flight vehicles. To provide a computationally economical first order approximation to the high temperature physics, variable thermodynamic data is used for the chemically frozen mode to allow for a variation with temperature of the air specific heats and enthalpy. For calculations involving air in chemical equilibrium, a specially modified version of the NASA Lewis Chemical Equilibrium Code, CEC, is used to compute the chemical composition and resultant thermochemical properties. The upgraded solver is demonstrated by comparing results from calorically perfect (C sub p=constant), thermally perfect (frozen) and equilibrium air calculations for a variety of geometries, and flight Mach numbers.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3076
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  • 164
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    PNS calculations for 3-D hypersonic corner flow with two turbulence models (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A three-dimensional parabolized Navier-Stokes code has been used as a testbed to investigate two turbulence models, the McDonald Camarata and Bushnell Beckwith model, in the hypersonic regime. The Bushnell Beckwith form factor correction to the McDonald Camarata mixing length model has been extended to three-dimensional flow by use of an inverse averaging of the resultant length scale contributions from each wall. Two-dimensional calculations are compared with experiment for Mach 18 helium flow over a 4-deg wedge. Corner flow calculations have been performed at Mach 11.8 for a Reynolds number of .67 x 10 to the 6th, based on the duct half-width, and a freestream stagnation temperature of 1750-deg Rankine.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2958
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  • 165
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    Investigation of oscillating cascade aerodynamics by an experimental influence coefficient technique (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Fundamental experiments are performed in the NASA Lewis Transonic Oscillating Cascade Facility to investigate the torsion mode unsteady aerodynamics of a biconvex airfoil cascade at realistic values of the reduced frequency for all interblade phase angles at a specified mean flow condition. In particular, an unsteady aerodynamic influence coefficient technique is developed and utilized in which only one airfoil in the cascade is oscillated at a time and the resulting airfoil surface unsteady pressure distribution measured on one dynamically instrumented airfoil. The unsteady aerodynamics of an equivalent cascade with all airfoils oscillating at a specified interblade phase angle are then determined through a vector summation of these data. These influence coefficient determined oscillating cascade data are correlated with data obtained in this cascade with all airfoils oscillating at several interblade phase angle values. The influence coefficients are then utilized to determine the unsteady aerodynamics of the cascade for all interblade phase angles, with these unique data subsequently correlated with predictions from a linearized unsteady cascade model.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2815
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  • 166
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    Implementation of a hypersonic rarefied flow particle simulation on the Connection Machine (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A very efficient direct particle simulation algorithm for hypersonic rarefied flows is presented and its implmentation on a Connection Machine is described. The implementation simulates ideal diatomic Maxwell molecules with three translational and two rotational degrees of freedom. Results for a 2-D simulation of supersonic flow over a 30 deg wedge are presented and used for validation.
    Keywords: AERODYNAMICS
    Type: NASA-CR-185428 , NAS 1.26:185428 , RIACS-TR-88.46
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  • 167
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    Evaluation and application of a new interferometry technique for compressible flow research (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A new method for obtaining large scale interferograms of flow fields in real time was investigated. The method was based on the point diffraction interferometry technique. The method was modified to accommodate the higher laser power required in recording transonic and supersonic flow fields. Basic tests were conducted in unsteady flows and flows about circulation control airfoils at transonic speeds. It was found that vibration was not a significant factor in the application of the system. In the case of the circulation control airfoils, the real-time viewing allowed the identification of the Coanda jet interaction with the external flow and the shedding of large scale vortices. The method proved to be very sensitive to the optical quality of the wind tunnel windows. The results obtained were compared with earlier interferograms obtained using interferometry. These results were in qualitative agreement.
    Keywords: AERODYNAMICS
    Type: NASA-CR-177467 , NAS 1.26:177467
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  • 168
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    Simulation of 2-dimensional viscous flow through cascades using a semi-elliptic analysis and hybrid C-H grids (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A semi-elliptic formulation, termed the interacting parabolized Navier-Stokes (IPNS) formulation, is developed for the analysis of a class of subsonic viscous flows for which streamwise diffusion is neglible but which are significantly influenced by upstream interactions. The IPNS equations are obtained from the Navier-Stokes equations by dropping the streamwise viscous-diffusion terms but retaining upstream influence via the streamwise pressure-gradient. A two-step alternating-direction-explicit numerical scheme is developed to solve these equations. The quasi-linearization and discretization of the equations are carefully examined so that no artificial viscosity is added externally to the scheme. Also, solutions to compressible as well as nearly compressible flows are obtained without any modification either in the analysis or in the solution process. The procedure is applied to constricted channels and cascade passages formed by airfoils of various shapes. These geometries are represented using numerically generated curilinear boundary-oriented coordinates forming an H-grid. A hybrid C-H grid, more appropriate for cascade of airfoils with rounded leading edges, was also developed. Satisfactory results are obtained for flows through cascades of Joukowski airfoils.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4180 , REPT-86-9-71 , E-4286 , NAS 1.26:4180
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  • 169
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    Three-dimensional self-adaptive grid method for complex flows (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A self-adaptive grid procedure for efficient computation of three-dimensional complex flow fields is described. The method is based on variational principles to minimize the energy of a spring system analogy which redistributes the grid points. Grid control parameters are determined by specifying maximum and minimum grid spacing. Multidirectional adaptation is achieved by splitting the procedure into a sequence of successive applications of a unidirectional adaptation. One-sided, two-directional constraints for orthogonality and smoothness are used to enhance the efficiency of the method. Feasibility of the scheme is demonstrated by application to a multinozzle, afterbody, plume flow field. Application of the algorithm for initial grid generation is illustrated by constructing a three-dimensional grid about a bump-like geometry.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101027 , A-88277 , NAS 1.15:101027
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  • 170
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    Adaptive wall technology for minimization of wall interferences in transonic wind tunnels (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Modern experimental techniques to improve free air simulations in transonic wind tunnels by use of adaptive wall technology are reviewed. Considered are the significant advantages of adaptive wall testing techniques with respect to wall interferences, Reynolds number, tunnel drive power, and flow quality. The application of these testing techniques relies on making the test section boundaries adjustable and using a rapid wall adjustment procedure. A historical overview shows how the disjointed development of these testing techniques, since 1938, is closely linked to available computer support. An overview of Adaptive Wall Test Section (AWTS) designs shows a preference for use of relatively simple designs with solid adaptive walls in 2- and 3-D testing. Operational aspects of AWTS's are discussed with regard to production type operation where adaptive wall adjustments need to be quick. Both 2- and 3-D data are presented to illustrate the quality of AWTS data over the transonic speed range. Adaptive wall technology is available for general use in 2-D testing, even in cryogenic wind tunnels. In 3-D testing, more refinement of the adaptive wall testing techniques is required before more widespread use can be planned.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4191 , NAS 1.26:4191
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  • 171
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    Experimental results for the Eppler 387 airfoil at low Reynolds numbers in the Langley low-turbulence pressure tunnel (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experimental results were obtained for an Eppler 387 airfoil in the Langley Low Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.03 to 0.13 and a chord Reynolds number range for 60,000 to 460,000. Lift and pitching moment data were obtained from airfoil surface pressure measurements and drag data for wake surveys. Oil flow visualization was used to determine laminar separation and turbulent reattachment locations. Comparisons of these results with data on the Eppler 387 airfoil from two other facilities as well as the Eppler airfoil code are included.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4062 , L-16430 , NAS 1.15:4062
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  • 172
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    Turbofan forced mixer lobe flow modeling. Part 3: Application to augment engines (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Military engines frequently need large quantities of thrust for short periods of time. The addition of an augmentor can provide such thrust increases but with a penalty of increased duct length and engine weight. The addition of a forced mixer to the augmentor improves performance and reduces the penalty, as well as providing a method for siting the required flame holders. In this report two augmentor concepts are investigated: a swirl-mixer augmentor and a mixer-flameholder augmentor. Several designs for each concept are included and an experimental assessment of one of the swirl-mixer augmentors is presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4147-PT-3 , E-4085 , NAS 1.26:4147-PT-3
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  • 173
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    Weak-wave analysis of shock interaction with a slipstream (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A weak wave analysis of shock interaction with a slipstream is presented. The theory is compared to that for the acoustic case and to the exact nonlinear analysis. Sample calculations indicate that the weak wave theory yields a good approximation to the exact solution when the shock waves are sufficiently weak that the associated entropy increase is negligible. A qualitative discussion of the case of counterflowing streams is also included.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2848 , L-16469 , NAS 1.60:2848
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  • 174
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    Design point variation of 3-D loss and deviation for axial compressor middle stages (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The available data on middle-stage research compressors operating near design point are used to derive simple empirical models for the spanwise variation of three-dimensional viscous loss coefficients for middle-stage axial compressor blading. The models make it possible to quickly estimate the total loss and deviation across the blade span when the three-dimensional distribution is superimposed on the two-dimensional variation calculated for each blade element. It is noted that extrapolated estimates should be used with caution since the correlations have been derived from a limited data base.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 88-GT-57
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  • 175
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    Suction laminarization of highly swept supersonic laminar flow control wings (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An evaluation is made of a suction-based method for the laminarization of highly-swept supersonic wings at cruise Mach numbers in the 2.0-2.5 range, in the interest of the reduction of wave drag due to lift. The laminar boundary layer development, as well as Tollmien-Schlichting and crossflow instabilities, have been analyzed for the case of an X66 supercritical airfoil at 60 and 72 deg sweep, for Mach numbers of 1.56 and 2.52, respectively. Strong suction is found to be needed at the front part of the upper surface and both the upper and lower rear pressure-rise areas.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-4471
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  • 176
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    Flutter of a fan blade in supersonic axial flow (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An application of a simple aeroelastic model to an advanced supersonic axial flow fan is presented. Lane's cascade theory is used to determine the unsteady aerodynamic loads. Parametric studies are performed to determine the effects of mode coupling, Mach number, damping, pitching axis location, solidity, stagger angle, and mistuning. The results show that supersonic axial flow fan and compressor blades are susceptible to a strong torsional mode flutter having critical reduced velocities which can be less than one.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 88-GT-78
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  • 177
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    Viscous induced drag (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A fundamentally new approach to the aircraft minimum induced drag problem is presented. The method, a 'viscous lifting line', is based on the minimum entropy production principle and does not require the planar wake assumption. An approximate, closed form solution is obtained and compared with several classical results. In addition, the problem of optimizing in-plane wing sweep with constant wing root bending moment is considered. Like the classical lifting line theory, this theory predicts that induced drag is proportional to the square of the lift coefficient and inversely proportional to the wing aspect ratio. Unlike the classical theory, it predicts that in-plane wing sweep may significantly reduce induced drag, that induced drag is Reynolds number dependent, and that the optimum spanwise circulation distribution is non-elliptic.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2550
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  • 178
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    Numerical study of chemically reacting flows using an LU scheme (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A new computational fluid dynamic code has been developed for the study of mixing and chemical reactions in the flow fields of ramjets and scramjets. The code employs an implicit finite volume, lower-upper symmetric successive overrelaxation scheme for solving the complete two-dimensional Navier-Stokes equations and species transport equations in a fully-coupled and very efficient manner. The combustion processes are modeled by an 8-species, 14-step finite rate chemistry model whereas turbulence is simulated by a Baldwin-Lomax algebraic model. The validity of the code is demonstrated by comparing the numerical calculations with both experimental data and previous calculations of a cold flow helium injection into a straight channel and premixed hydrogen-air reacting flows in a ramped duct. The code is then used to calculate the mixing and chemical reactions of a hydrogen jet transversely injected into a supersonic airstream. Results are presented describing the flow field, the recirculation regions in front and behind the injector, and the chemical reactions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0436
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  • 179
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    A longitudinally-patched grid approach with applications to high speed flows (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The combined use of an implicit, upwind finite-volume scheme with an 'equivalent' gamma formulation (for real gas effects) and with a three-dimensional conservative patched grid formulation is discussed. Results are presented over a wide Mach number range on both single and patched grids.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0715
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  • 180
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    Application of a nonisentropic full potential method to AGARD standard airfoils (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: An entropy-correction method for the unsteady full potential equation is presented. The unsteady potential equation is modified to model the entropy jumps across shock waves. The conservative form of the modified equation is solved in generalized coordinates using an implicit, approximate factorization method. A flux-biasing differencing method, which generates the proper amounts of artificial viscosity in supersonic regions, is used to discretize the flow equations in space. Calculated results are presented for the NLR 7301, NACA 0012, and NACA 64A010A airfoils. Comparisons of the present method and solutions of the Euler equations are presented for the NLR 7301 airfoil, and comparisons of the present method and experimental data are presented for all three airfoils. The comparisons show that the present method more accurately models solutions of the Euler equations and experiment than does the isentropic potential formulation. In addition, it is shown that modeling shock-generated entropy extends the range of validity of the full potential method.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0710
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  • 181
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    Aerodynamic interaction between propellers and wings (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A combined computational/experimental investigation has been conducted to determine the time-averaged interactive performance of a propeller and wing in tractor configuration at Mach 0.1 and Re=470,000, based on a wind tunnel model wing chord of 8 in. Wing angle-of-attack was varied from 0 to +13 deg, and propeller advance ratio ranged from 2.4 (windmilling) to 1.1 (maximum power). Both a semiempirical model and a vortex lattice simulation were used in the computational analysis. Good agreement has been obtained between theory and experiment.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0665
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  • 182
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    A multi-element vortex lattice method for calculating the geometry and effects of a helicopter rotor wake in forward flight (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A method is described for the analysis of the unsteady, incompressible potential flow associated with a helicopter rotor and it's wake in forward flight. This method is particularly useful in low advance ratio flight due to the major contribution, in the near field, of the deformed wake. The rotor geometry is prescribed and the unsteady wake geometry is computed from the local flow perturbation velocities. The wake is modeled as a full vortex lattice. The rotor geometry is arbitrary and several rotor blades can be represented. The unsteady airloads on the rotor blades are computed in the presence of the deformed rotor wake by a time-stepping technique. Solution for the load distribution on the blade surfaces is found by prescribing boundary conditions in a reference system which rotates with the blade tips. Transformation tensors are used to describe the contribution of the wake in the inertial system to the rotor in the rotating reference system. The effects of blade cyclic pitch variation are computed using a rotation tensor. The deformation of the wake is computed in the inertial frame. The wake is started impulsively from rest, allowing a natural convection of the wake with time.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0664
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  • 183
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    Effects of numerical dissipation on finite-volume solutions of compressible flow problems (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The effects of numerical dissipation upon solutions to the Euler equations are considered, and results for transonic flows past airfoils are presented to demonstrate the effects of the dissipative terms. The equations are approximated using a finite-volume spatial approximation with added dissipation provided by an adaptive mixture of second and fourth differences. The resulting difference equations are solved using either an explicit multistage Runge-Kutta method or a diagonalized implicit method. It is found that errors in surface values can be introduced by the averaging required to calculate derived quantities of interest.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0621
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  • 184
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    Numerical investigation of three-dimensional flow separation using the boundary layer equations (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The steady, incompressible, three-dimensional laminar and turbulent boundary-layer equations are solved in a streamline coordinate system and in a self-adaptive grid system using Matsuno's finite difference method. Techniques are described for calculating laminar and turbulent separation using the boundary-layer equations. Any type (bubble type or free vortex-layer type) of major separation line can be calculated at an angle of attack on ellipsoids of revolution by this boundary layer code. Results are presented for ellipsoids of revolution at angles of attack up to 45 degrees. Agreements with other numerical and experimental results are very good for laminar flows. Turbulent flows are also investigated with algebraic turbulence models proposed by Rotta and Cebeci and Smith. Good agreement with experimental results was obtained at a small angle of attack (10 degrees) but only qualitative agreement was obtained at a high angle of attack (30 degrees) for turbulent flow on a 6:1 ellipsoid of revolution.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0617
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  • 185
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    Flow control in a wing/fuselage-type juncture (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The laminar flow around a juncture formed by an unswept wing and a flat plate has been studied using a combination of smoke flow visualization, and velocity and pressure measurements. The effectiveness of swept leading-edge fillets in controlling the juncture flow field has been evaluated. Flow separation upstream of the wing leading edge is confined to a small region near the plate. This separation results in periodic shedding of horseshoe type vortices. The pressure gradient measured upstream of the leading edge in this laminar juncture is steeper than that of the turbulent flow case. The use of fillets eliminates the leading-edge flow separation and reduces the size of juncture wake, as observed from flow visualization. For one of the filleted cases, there is a significant increase in the extent of laminar flow in the juncture region, and a sizable reduction in the juncture drag.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0614
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  • 186
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    Turbulent separated flow in the vicinity of a single-slotted airfoil flap (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Detailed measurements of pressure and velocity characteristics are presented and analyzed for flow over and downstream of a NACA 4412 airfoil equipped with a NACA 4415 single-slotted flap at high angle of attack and close to maximum lift. The flow remained attached over the main element while a large region of recirculating flow occurred over the aft 61 percent of the flap. The airfoil configuration was tested at a Mach number of 0.09 and chord based Reynolds number of 1.8 x 10 to the 6th in the NASA Ames Research Center 7- by 10-Foot Wind Tunnel. Measurements of mean and fluctuating velocities were obtained in the region of recirculation and high turbulence intensity using three-dimensional laser velocimetry. In regions where the flow had a preferred direction and relatively low turbulence intensity hot-wire anemometry was used. Emphasis was placed on obtaining flow characteristics in the confluent boundary layer, the region of recirculating flow and in the downstream wake. Surface pressure measurements were made on the main airfoil, flap, wind tunnel roof, and wind tunnel floor. In addition to the presentation of pressure and velocity characteristics, the near wall results inside the separated region are analyzed as are the relative importance of terms in the momentum and turbulence kinetic energy equations in the confluent separated boundary layer and the recirculating region of the near wake.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0613
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  • 187
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    Performance characteristics from wind-tunnel tests of a low-Reynolds-number airfoil (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Wind tunnel lift and pitching-moment data have been obtained from pressure measurements, and drag data from wake surveys, for an Eppler 387 low Reynolds number airfoil over the Re range of 60,000 to 460,000; oil flow visualizations were also used to determine laminar separation and turbulent reattachment locations. Airfoil performance is found to be dominated by laminar separation bubbles below Re 200,000, and two flow regimes, namely laminar separations with and without turbulent reattachment, were observed at the same angle-of-attack for an Re of 60,000.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0607
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  • 188
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    Numerical study of turbulent internal shear layer flow in an axisymmetric U-duct (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper presents a simple algebraic turbulence model developed for internal flows which is based on the vorticity length and magnitude. The length scale is defined using the distance between the maximum and the minimum points of the absolute vorticity. This model is numerically tested in a turbulent internal layer flow through an axisymmetric U-duct with very sharp curvature; the ratio of the boundary layer thickness to the radius of curvature of the duct is of order 1. In this U-duct flow, strong adverse and favorable pressure gradients coexist and interact with each other. Satisfactoy agreement with experimental results is obtained.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0596
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  • 189
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    Navier-Stokes cascade analysis with a stiff k-epsilon turbulence solver (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The two dimensional, compressible, thin layer Navier-Stokes equations with the Baldwin-Lomax turbulence model and the kinetic energy-energy dissipation (k-epsilon) model are solved numerically to simulate the flow through a cascade. The governing equations are solved for the entire flow domain, without the boundary layer assumptions. The stiffness of the k-epsilon equations is discussed. A semi-implicit, Runge-Kutta, time-marching scheme is developed to solve the k-epsilon equations. The impact of the k-epsilon solver on the explicit Runge-Kutta Navier-Stokes solver is discussed. Numerical solutions are presented for two dimensional turbulent flow over a flat plate and a double circular arc cascade and compared with experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0594
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  • 190
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    The effect of spatial wind gradients on airplane aerodynamics (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An account is given of the results of recent studies of the effect of a spatially sheared wind field on airplane aerodynamics; the wind shear was computed by a modified vortex-lattice computer program, and characterized through the formulation of wind shear aerodynamic coefficients. The magnitude of the aerodynamic effect was demonstrated by computing the change in conventional wing/tail configuration aerodynamics for a fixed flight path through a simulated microburst. A substantial portion of the control authority of the aircraft may be required to counteract the wind shear-induced forces and moments in the microburst environment; both aperiodic and oscillatory instabilities may be generated by shear-dependent dynamic modes.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0579
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  • 191
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    Euler/Navier-Stokes zonal scheme with applications to separated and isolated-vortex flows (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An Euler/Navier-Stokes zonal scheme with boundary-layer compatibility conditions is developed to economically compute separated and vortex flows. The scheme is based on dividing the flow region into zones, where different levels of mathematical approximations of the governing equations are used in each zone. The scheme is applied to two specific problems: the two dimensional flow over a blunt leading-edge plate and the quasi-axisymmetric flow of an isolated vortex core. In the first problem, the computational domain is divided into inner and outer zones where the Navier-Stokes and Euler equations are used, respectively. On the downstream boundary of the computational domain, boundary-layer compatibility conditions are used. In the second problem, boundary-layer-like equations for slender, compressible, vortex flows are developed. A compatibility condition has been used to ensure consistency of the boundary and initial conditions. The outer boundary conditions of the flow are derived from Euler equations for a stream surface.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0507
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  • 192
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    Control of leading edge vortex breakdown by blowing (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An investigation into the effects of using a jet of air to control the vortex breakdown position on a 70 degree delta wing is presented. The specific objectives focused on optimizing the blowing positions in terms of maximum lift increments obtained for minimum blowing rates. The tests were conducted at chord Reynolds numbers of 150,000, 200,000, and 250,000 at angles of incidence of 30 and 35 degrees. Visualization and force data is presented to show the effect of the jet on the wing aerodynamic characteristics. The results indicate a jet position located at and aligned parallel to the leading edge to be the optimum. Nearness to the apex and tangency to the upper surface were also crucial factors. The influence of the jet on the leading edge vortex structure was examined using laser Doppler anemometry. Velocity surveys through the vortex showed that at high blowing rates the parallel velocity in the outer swirling region of the vortex increased and the normal velocity decreased. This resulted in a decrease in the swirling angle in the outer region. The peak core velocity was reduced and the vortex breakdown was delayed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0504
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  • 193
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    Skin friction measurements by laser interferometry in swept shock wave/turbulent boundary-layer interactions (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The laser interferometric skin friction meter was used to measure wall shear stress distributions in two interactions of fin-generated swept shock waves with turbulent boundary layers. The basic research configuration was an unswept sharp-leading-edge fin of variable angle mounted on a flatplate. The results indicate that such measurements are practical in high-speed interacting flows, and that a repeatability of + or - 6 percent or better is possible. Marked increases in wall shear were observed in both swept interactions tested.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0497
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  • 194
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    Computations of vortical supersonic flows over a blunt-nose-cylinder at high angles of attack (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Three-dimensional, viscous, and separated flows over a blunt-nose-cylinder at 20, 32, and 44-deg attack angles were computed. The approaching freestream was at a Mach number of 1.6 and a unit Reynolds number of 2 x 10 to the 6th/ft with a total temperature of 585 R. The cylinder used for the computations had a length-to-diameter ratio of 6.67 with a base diameter of 3 in. The flowfield was dominated by large-scale and multiple vortices generated by crossflow separation. The effect of turbulence on the flow structure of one case was modeled algebraically with modifications to correct the length and velocity scales in the regions of separation. The mass averaged Navier-Stokes equations were solved by an approximately factored, upwind-biased, implicit, finite volume scheme. The initialization of the flows was enhanced by a mesh sequencing strategy applied to the diagonalized form of the discretized equations. The convergence to steady-state was accelerated by a multigrid algorithm and using the block inversions for the discretized equations. Calculations were compared with experimental results.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0485
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  • 195
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    Effects of surface perturbations on the asymmetric vortex flow over a slender body (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An experimental investigation of the effects of surface perturbations on the asymmetric flow past a slender body has been conducted for laminar flow conditions. Beads with diameters ranging from 3/32 to 12/32 in. were attached near the apex of a cone/cylinder model having a base diameter of 3.5 in. and a cone semiapex angle of 9 deg at an angle of attack of 40 deg in an attempt to alter the sense of the asymmetric vortex flow pattern. Circumferential position as well as longitudinal location were varied to determine the most effective bead position. Whether or not the beads were effective in controlling the magnitude and direction of the vortex asymmetries was determined by 3 circumferential rows of pressure taps and by a helium-bubble flow visualization technique. The most effective circumferential position was found to be approximately 140 deg from the windward ray. While holding this circumferential position constant, the effect of bead size at three stations further along the body was also investigated. It was found that the size of the bead necessary to reverse the asymmetry increased more rapidly than the growth in cylinder radius. In general, these results indicate that discrete geometric imperfections on a body's surface can force asymmetry in a given direction if they are sufficiently large relative to the local radius.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0483
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  • 196
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    Evaluation of a three-dimensional empirically derived wing at supersonic speeds (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A novel wing design concept is introduced which takes advantage of the existence of conical flow at supersonic speeds. The present wing design concept is to create a near conical wing geometry by redistributing airfoils in a spanwise direction. In addition, a set of graphs which review the supersonic aerodynamics of delta wings have been employed to select a design wing sweep and Mach number. An iteration through the wing design logic resulted in the selection of a 65 deg swept delta wing and a design Mach number of 1.62. Theoretical analysis was performed with a nonlinear full-potential analysis method to assess the merits of the wing design approach. The analysis showed large reductions in drag due to lift compared to delta wings configured with traditional thickness and airfoil distributions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0481
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  • 197
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    Studies of aerothermal loads generated in regions of shock/shock interaction in hypersonic flow (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Detailed measurements of heat transfer and pressure distribution have been made in the present study of the aerothermal characteristics of regions of two-dimensional shock/shock interaction generated by the incidence of single and multiple shocks onto the bow shock ahead of a spanwise cylinder. For transitional flows, the measurements demonstrated a large increase in the aerothermal loads with increasing Mach number that cannot be predicted by simple phenomenological models. The studies with multiple incident shocks demonstrate that the largest aerothermal loads are generated on the cylinder when the shocks coalesce before they are incident on the bow shock.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0477
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  • 198
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    A full-potential theory analysis of the supersonic aerodynamics of a 60-deg delta wing-body configuration (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Experimental spanwise pressure distributions for a 60-deg delta wing/body of approximate fineness ratio 7.6 have been obtained and compared to predictions using full-potential theory. Analysis was performed at Mach 1.6 for angles of attack in the range 0.8 to 10 deg, and for Mach numbers ranging from 1.4 to 1.8 at lift coefficients 0.3 and 0.4. The intent of the study was to examine an attached flow approach for maneuver wing design in the presence of a fuselage. For the Mach number, angle-of-attack conditions considered, the full-potential theory accurately modeled the pressure distributions provided the flow remained attached. By combining the full-potential theory results with an empirical shock-induced separation criterion, it was found that the onset of shock-induced separation can be predicted. The investigation showed that, if an attached-flow approach is used with an empirical method of indicating shock-induced separation, the full-potential method is capable of being used as an effective tool for designing maneuver wings.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0480
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  • 199
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    Direct simulation of three-dimensional hypersonic flow about intersecting blunt wedges (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The general three-dimensional direct simulation Monte Carlo method is used to study the hypersonic flow around two blunt wedges that intersect at a 90 deg angle. Results are obtained for the transitional flow regimes found at 85 and 100 km altitude with a reentry velocity of 7.5 km/s. The disturbance field in front of the double-wedge body is found to be larger than that produced by a single wedge. Surface pressures and flow densities are higher near the wedge intersection, whereas surface heating and shear streses are greater at locations removed from the corner. Results also show that three-dimensional flow structure occurs only near the wedge corner, and that the flow monotonically approaches the limiting two-dimensional wedge flow case in the spanwise direction.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0463
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  • 200
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    Investigation of two-dimensional chemically reacting and radiating supersonic channel flows (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The two-dimensional time-dependent Navier-Stokes equations are solved in order to study supersonic flows with finite rate chemistry and radiation for hydrogen-air systems. The problem of the flow in a channel with a ten-degree compression-expansion ramp is solved using the finite volume technique of Jameson et al. (1981) and the unsplit finite difference scheme of MacCormack (1969). The problem of chemically reacting and radiating flows is considered for the flow of premixed hydrogen-air through a channel with parallel boundaries and a channel with a compression corner. Results suggest that radiative interaction can have a significant effect on the entire flowfield.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0462
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