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At least 1,621 records · Page 90

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NASA NTRS1975

A fast semidirect method for computing transonic aerodynamic flows

A fast, semidirect, iterative computational method, previously introduced for finite-difference solution of subsonic and slightly supercritical flow over airfoils, is extended both to apply to strongly supercritical conditions and to include full second-order accuracy in computing inviscid flows over airfoils. The nonlinear small-disturbance equations are solved iteratively by a direct, linear, elliptic solver. General, fully conservative, type-dependent difference equations are formulated, including parabolic- and shock-point transition operators that provide consistency with the integral conservation laws. These equations specialize to either first-order or to fully second-order-accurate equations. Various free parameters are evaluated for rapid convergence of the first-order scheme. Resulting pressure distributions and computing times are compared with the improved Murman-Cole line-relaxation method.

Martin, E. D.↗

NASA NTRS1975

Supercritical wing sections II: A handbook

The numerical aspect of theoretical work on transonics and supercritical wing sections are compiled. A model of the trailing edge is introduced which eliminates the loss of 15 to 20 percent experienced with heavily aft-loaded models, and it is indicated how drag creep can be reduced at off-design conditions. A rotated finite difference scheme is presented which can handle supersonic as well as subsonic free stream Mach numbers and leads to an effective three-dimensional program for the computation of transonic flow past an oblique wing. In the case of two-dimensional flow, the method is extended to take into account the displacement thickness computed by a semiempirical turbulent boundary layer correction. A series of supercritical wing sections is discussed together with comparisons between experimental and theoretical data. Computer programs and a brief manual for their operation are listed. It is shown that the programs furnish a physically adequate computer simulation of the compressible flows that arise in problems of transonic aerodynamics.

Bauer, F.↗

NASA NTRS1975

Inviscid supersonic/hypersonic body flowfields and aerodynamics from shock-capturing technique calculations

A shock-capturing, finite-difference computational procedure has been used to predict the body flowfields over a wide range of initial conditions, and these results have been compared with wind tunnel data. Two different studies have been performed. In the first, details of the shock layer surrounding space shuttle orbiter configurations have been calculated. In the second, the aerodynamic forces and moments on a parametrically varied set of blunt biconic bodies of revolution have been evaluated for angles of attack up to 20 degrees and Mach numbers from 5 to 20.

Chaussee, D. S.↗

NASA NTRS1975

Aerothermodynamic environment for Jovian entry conditions

Solutions are presented for the aerothermal heating environments about blunt probes entering the atmosphere of Jupiter. The analysis employs a novel prediction procedure in which finite differences versions of the thin layer fluid dynamic equations are solved simultaneously with, and coupled to, a fully detailed radiation flux model. Predictions are presented and compared with those of earlier investigators. Radiative blowing corrections are presented for on- and off-stagnation point conditions. The use of a blowing correlation based on the retained mass was found to be useful for off-stagnation point conditions.

Nicolet, W. E.↗

NASA NTRS1975

Application of a general boundary layer analysis to turbulent boundary layers subjected to strong favorable pressure gradients

Theoretical predictions of turbulent boundary layer development under the influence of strong favorable pressure gradients made using a finite-difference calculation procedure are compared to experimental data. Comparisons are presented for low speed flows with and without wall heat transfer as well as for supersonic flows with adiabatic walls. The turbulence model used is governed by an integral form of the turbulence kinetic energy equation and the results are compared with predictions made using a conventional equilibrium turbulence model based upon Prandtl's mixing length, a Clauser-type eddy viscosity model used by Cebecci and Mosinskis, and a two-equation turbulence energy model of Launder and Jones.

Kreskovsky, J. P.↗

NASA NTRS1975

Application of boundary integral equations to elastoplastic problems

The application of the boundary integral equation method (BIE) to the elastoplastic torsion problem is considered. It is found that the BIE is very suitable for the elastoplastic analysis of the torsion of prismatic bars. A comparison of the BIE with the finite difference method shows savings for the BIE concerning the number of unknowns which have to be determined and also a much faster convergence rate. Attention is given to the problem of an edge-notched beam in pure bending, taking into account a biharmonic formulation and a displacement formulation.

Mendelson, A.↗

NASA NTRS1975

High Reynolds number transonic flow simulation

A code has been developed for simulating high Reynolds number transonic flow fields of arbitrary configuration. An explicit finite-difference method with time splitting is used to solve the time-dependent equations for compressible turbulent flow. A nonorthogonal computational mesh of arbitrary configuration facilitates the description of the flow field. The code is applied to simulate the flow over a two-dimensional 18 percent thick circular-arc biconvex airfoil at zero angle of attack for several different Reynolds numbers and a free-stream Mach number of 0.775.

Deiwert, G. S.↗

NASA NTRS1975

Solution of the three-dimensional compressible, laminar, and turbulent boundary-layer equations with comparisons to experimental data

A technique for solving three-dimensional boundary-layer flows is presented. The governing equations are transformed to obtain a nondimensional form and a similarity-type transform is defined for the normal coordinate and velocity. The resulting equation is solved in an iterative mode using a finite difference technique suggested by Dwyer (1968). A simple mean field turbulence model is discussed and numerical data obtained in a test case are compared with experimental data.-

Harris, J. E.↗

NASA NTRS1975

A generalized hyperbolic marching technique for three-dimensional supersonic flow with shocks

The current paper introduces a generalized, nonorthogonal coordinate system for application to supersonic flows with such large angles between streamlines of the initial Cauchy data and the marching direction that the velocity component in the marching direction is actually subsonic. Finite-difference operators are developed to solve the three-dimensional, steady, inviscid conservation equations of fluid flow referenced to this general frame. Moreover, a new method of aligning the difference mesh to the bow shock wave is presented which eliminates both the necessity for differencing the free-stream flow properties and the spurious fluctuations that often arise in conservative variables when they are differenced across the discontinuity. Results are given for the case of flow past a conical surface.

Rizzi, A. W.↗

NASA NTRS1975

An experimental and analytical investigation of axisymmetric diffusers

A finite difference computer program for turbulent compressible flow was used to establish the performance of several diffuser shapes for experimental testing. The diffusers were designed to have a linear change in Mach number, a linear change in pressure or a curvature fitted by a quadratic equation. Testing was performed with M = 0.1 to 0.9 with and without boundary layer bleed. Above M = 0.6, data were obtained with a normal shock upstream of the diffuser entrance. Peak static pressure recovery occurred with a diffuser inlet M-0.75. The quadratic diffuser yielded the highest total pressure recovery.

Povinelli, L. A.↗

NASA NTRS1975

Application of a transitional boundary-layer theory in the low hypersonic Mach number regime

An investigation is made to assess the capability of a finite-difference boundary-layer procedure to predict the mean profile development across a transition from laminar to turbulent flow in the low hypersonic Mach-number regime. The boundary-layer procedure uses an integral form of the turbulence kinetic-energy equation to govern the development of the Reynolds apparent shear stress. The present investigation shows the ability of this procedure to predict Stanton number, velocity profiles, and density profiles through the transition region and, in addition, to predict the effect of wall cooling and Mach number on transition Reynolds number. The contribution of the pressure-dilatation term to the energy balance is examined and it is suggested that transition can be initiated by the direct absorption of acoustic energy even if only a small amount (1 per cent) of the incident acoustic energy is absorbed.

Shamroth, S. J.↗

NASA NTRS1975

Scanning thermal plumes

Over a three-year period 800 thermal line scans of power plant plumes were made by an airborne scanner, with ground truth measured concurrently at the plants. Computations using centered finite differences in the thermal scanning imagery show a lower bound in the horizontal temperature gradient in excess of 1.6 C/m. Gradients persist to 3 m below the surface. Vector plots of the velocity of thermal fronts are constructed by tracing the front motion in successive thermal images. A procedure is outlined for the two-point ground calibration of a thermal scanner from an equation describing the scanner signal and the voltage for two known temperatures. The modulation transfer function is then calculated by input of a thermal step function and application of digital time analysis techniques using Fast Fourier Transforms to the voltage output. Field calibration tests are discussed. Data accuracy is limited by the level of ground truth effort chosen.

Scarpace, F. L.↗

NASA NTRS1975

Calculation of three-dimensional compressible laminar and turbulent boundary layers. Calculation of three-dimensional compressible boundary layers on arbitrary wings

A very general method for calculating compressible three-dimensional laminar and turbulent boundary layers on arbitrary wings is described. The method utilizes a nonorthogonal coordinate system for the boundary-layer calculations and includes a geometry package that represents the wing analytically. In the calculations all the geometric parameters of the coordinate system are accounted for. The Reynolds shear-stress terms are modeled by an eddy-viscosity formulation developed by Cebeci. The governing equations are solved by a very efficient two-point finite-difference method used earlier by Keller and Cebeci for two-dimensional flows and later by Cebeci for three-dimensional flows.

Cebeci, T.↗

NASA NTRS1975

Subgrid or Reynolds stress-modeling for three-dimensional turbulence computations

A review is given of recent advances in two distinct computational methods for evaluating turbulence fields, namely, statistical Reynolds stress modeling and turbulence simulation, where large eddies are followed in time. It is shown that evaluation of the mean Reynolds stresses, rather than use of a scalar eddy viscosity, permits an explanation of streamline curvature effects found in several experiments. Turbulence simulation, with a new volume averaging technique and third-order accurate finite-difference computing is shown to predict the decay of isotropic turbulence in incompressible flow with rather modest computer storage requirements, even at Reynolds numbers of aerodynamic interest.

Rubesin, M. W.↗

NASA NTRS1975

Numerical computation of viscous flows on the lee side of blunt shapes flying at supersonic speeds

A numerical method for solving the parabolic approximation to the steady-state compressible Navier-Stokes equations is examined. The approximation neglects only the streamwise gradients of shear stress. An implicit finite difference method is used which advances the solution downstream from an initial data surface and determines the complete viscous-inviscid flow between the body and bow shock wave. It is necessary that the inviscid portion of the flow field be supersonic. Crossflow separation is determined as part of the solution. The method is applied to a 15 deg sphere-cone at 15 deg angle of attack, and the results are compared with an inviscid method-of-characteristics calculation.

Rakich, J. V.↗

NASA NTRS1975

Time-dependent transonic flow solutions for axial turbomachinery

Three-dimensional unsteady transonic flow through an axial turbomachine stage is described in terms of a pair of two-dimensional formulations pertaining to orthogonal surfaces, namely, a blade-to-blade surface and a hub-to-casing surface. The resulting systems of nonlinear, inviscid, compressible equations of motion are solved by an explicit finite-difference technique. The blade-to-blade program includes the periodic interaction between rotor and stator blade rows. Treatment of the boundary conditions and of the blade slipstream motion by a characteristic type procedure is discussed in detail. Harmonic analysis of the acoustic far field produced by the blade row interaction, including an arbitrary initial transient, is outlined. Results from the blade-to-blade program are compared with experimental measurements of the rotating pressure field at the tip of a high-speed fan. The hub-to-casing program determines circumferentially averaged flow properties on a meridional plane. Blade row interactions are neglected in this formulation, but the force distributions over the entire blade surface for both the rotor and stator are obtained. Results from the hub-to-casing program are compared with a relaxation method solution for a subsonic rotor. Results are also presented for a quiet fan stage which includes transonic flow in both the rotor and stator and a normal shock in the stator.

Erdos, J.↗

NASA NTRS1975

A comparison of a shock-capturing technique with experimental data for three-dimensional internal flows

Shock-capturing solutions for an axisymmetric supersonic inlet at small angles of attack are obtained. Good overall agreement between the shock-capturing solutions and experimental data is shown except in regions of strong viscous effects or boundary-layer removal. Although the results indicate a strong potential for the use of shock-capturing or finite-difference solutions for internal flows, improvement in the ability to handle the reflection of strong shockwaves having downstream Mach numbers near 1 is needed.

Presley, L. L.↗

NASA NTRS1975

Internal and external axial corner flows

The inviscid, internal, and external axial corner flows generated by two intersecting wedges traveling supersonically are obtained by use of a second-order shock-capturing, finite-difference approach. The governing equations are solved iteratively in conical coordinates to yield the complicated wave structure of the internal corner and the simple peripheral shock of the external corner. The numerical results for the internal flows compare favorably with existing experimental data.

Kutler, P.↗