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At least 163 records · Page 9

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NASA NTRS1995

Effect of turbulence models on criticality conditions in swirling flows

The critical state of vortex cores downstream of vortex breakdown has been studied. Base vortical flows were computed using the Reynolds-averaged, axisymmetric Navier-Stokes equations. Standard K - epsilon, RNG and second-order Reynolds stress models were employed. Results indicate that the return to supercriticality is highly dependent on the turbulence model. The K - epsilon model predicted a rapid return of the vortex to supercritical conditions, the location of which showed little sensitivity to changes in the swirl ratio. The Reynolds stress model predicted that the vortex remains subcritical to the end of the domain for each of the swirl ratios employed, and provided results in qualitative agreement with experimental work. The RNG model produced intermediate results, with a downstream movement in the critical location with increasing swirl. Calculations for which area reductions were introduced at the exit in a subcritical flow were also performed using the Reynolds stress model. The structure of the resulting recirculation zone was altered significantly. However, when area reductions were employed within supercritical flows as predicted using the two-equation models, no significant influence on the recirculation zone was noted.

Gatski, Thomas B.↗

NASA NTRS1986

Thermodynamic evaluation of transonic compressor rotors using the finite volume approach

A method was developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite volume, time marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier Stokes equations. The entire calculation is performed in the physical domain. The method is currently limited to the calculation of attached flows. The features of the current method can be summarized as follows. Control volumes are chosen so that smoothing of flow properties, typically required for stability, is now needed. Different time steps are used in the different governing equations to improve the convergence speed of the viscous calculations. A new pressure interpolation scheme is introduced which improves the shock capturing ability of the method. A multi-volume method for pressure changes in the boundary layer allows calculations which use very long and thin control volumes. A special discretization technique is also used to stabilize these calculations. A special formulation of the energy equation is used to provide improved transient behavior of solutions which use the full energy equation. The method is then compared with a wide variety of test cases. The freestream Mach numbers range from 0.075 to 2.8 in the calculations. Transonic viscous flow in a converging diverging nozzle is calculated with the method; the Mach number upstream of the shock is approximately 1.25. The agreement between the calculated and measured shock strength and total pressure losses is good. Essentially incompressible turbulent boundary layer flow in a adverse pressure gradient is calculated and the computed distribution of mean velocity and shear stress are in good agreement with the measurements. At the other end of the Mach number range, a flat plate turbulent boundary layer with a freestream Mach number of 2.8 is calculated using the full energy equation; the computed total temperature distribution and recovery factor agree well with the measurements when a variable Prandtl number is used through the boundary layer.

Nicholson, S.↗

NASA NTRS1989

Numerical simulation and comparison of symmetrical/supercritical airfoils for the near tip region of a helicopter in forward flight

Aerodynamic loads on a multi-bladed helicopter rotor in forward flight at transonic tip conditions are calculated. The unsteady, three-dimensional, time-accurate compressible Reynolds-averaged thin layer Navier-Stokes equations are solved in a rotating coordinate system on a body-conformed, curvilinear grid of C-H topology. Detailed boundary layer and global numerical comparisons of NACA-0012 symmetrical and CAST7-158 supercritical airfoils are made under identical forward flight conditions. The rotor wake effects are modeled by applying a correction to the geometric angle of attack of the blade. This correction is obtained by computing the local induced downwash velocity with a free wake analysis program. The calculations are performed on the Numerical Aerodynamic Simulation Cray 2 and the VPS32 (a derivative of a Cyber 205 at the Langley Research Center) for a model helicopter rotor in forward flight.

Badavi, F. F.↗

NASA NTRS1990

Implicit solution of three-dimensional internal turbulent flows

The scalar form of the approximate factorization method was used to develop a new code for the solution of three-dimensional internal laminar and turbulent compressible flows. The Navier-Stokes equations in their Reynolds-averaged form are iterated in time until a steady solution is reached. Evidence is given to the implicit and explicit artificial damping schemes that proved to be particularly efficient in speeding up convergence and enhancing the algorithm robustness. A conservative treatment of these terms at domain boundaries is proposed in order to avoid undesired mass and/or momentum artificial fluxes. Turbulence effects are accounted for by the zero-equation Baldwin-Lomax turbulence model and the q-omega two-equation model. For the first, an investigation on the model behavior in case of multiple boundaries is performed. The flow in a developing S-duct is then solved in the laminar regime at Reynolds number (Re) 790 and in the turbulent regime at Re=40,000 using the Baldwin-Lomax model . The Stanitz elbow is then solved using an inviscid version of the same code at M(sub inlet)=0.4. Grid dependence and convergence rate are investigated showing that for this solver the implicit damping scheme may play a critical role for convergence characteristics. The same flow at Re=2.5x10(exp 6) is solved with the Baldwin-Lomax and the q-omega models. Both approaches showed satisfactory agreement with experiments, although the q-omega model is slightly more accurate.

Michelassi, V.↗

NASA NTRS1986

Unsteady separated boundary layer in a transonic diffuser flow with self-excited oscillations

A numerical investigation of two-dimensional unsteady boundary layer in a transonic diffuser flow with self-excited oscillations and strong flow separation by solving the compressible, Reynolds-averaged, thin-layer Navier-Stokes equations with two-equations turbulence model is described. Three different meshes with constant streamwise mesh distribution and varying vertical mesh distribution were used. Results obtained indicate that a refinement of mesh studied here has minimal effect on the mean boundary layer flow but significantly increases the amplitude of oscillation of all flow variables. Comparisons of unsteady wall pressure, velocity profile, terminal shock, and separation pocket among computations and with experiment are presented.

Hsieh, T.↗

NASA NTRS1986

An explicit finite-volume time-marching procedure for turbulent flow calculations

A method was developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite-volume, time-marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier-Stokes equations. The entire calculation is performed in the physical domain. Control volumes are chosen so that smoothing of flow properties, typically required for stability, is not required. Different time steps are used in the different governing equations. A new pressure interpolation scheme is introduced which improves the shock capturing ability of the method. A multi-volume method for pressure changes in the boundary layer allows calculations which use very long and thin control volumes (length/height - 1000). The method is compared with two test cases. Essentially incompressible turbulent boundary layer flow in an adverse pressure gradient is calculated and the computed distributions of mean velocity and shear are in good agreement with the measurements. Transonic viscous flow in a converging diverging nozzle is calculated; the Mach number upstream of the shock is approximately 1.25. The agreement between the calculated and measured shock strength and total pressure losses is good.

Nicholson, Stephen↗

NASA NTRS1986

Explicit finite-volume time-marching calculations of total temperature distributions in turbulent flow

A method was developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite volume, time-marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier-Stokes equations. The entire calculation is performed in the physical domain. This paper investigates the introduction of a new formulation of the energy equation which gives improved transient behavior as the calculation converges. The effect of variable Prandtl number on the temperature distribution through the boundary layer is also investigated. A turbulent boundary layer in an adverse pressure gradient (M = 0.55) is used to demonstrate the improved transient temperature distribution obtained when the new formulation of the energy equation is used. A flat plate turbulent boundary layer with a supersonic free-stream Mach number of 2.8 is used to investigate the effect of Prandtl number on the distribution of properties through the boundary layer. The computed total temperature distribution and recovery factor agree well with the measurements when a variable Prandtl number is used through the boundary layer.

Nicholson, Stephen↗

NASA NTRS1987

Downstream boundary effects on the frequency of self-excited oscillations in transonic diffuser flows

An investigation of downstream boundary effects on the frequency of self-excited oscillations in two-dimensional, separated transonic diffuser flows has been conducted numerically by solving the compressible, Reynolds-averaged, thin-layer Navier-Stokes equation with a two-equation turbulence model. It was found that the unsteady diffuser flowfields are very sensitive to the location of the downstream boundary. Extension of the diffuser downstream boundary significantly reduces the frequency and amplitude of oscillations for pressure, velocity and shock. Computational results suggest that the mechanism causing the self-excited oscillation changes from viscous convective wave dominated oscillations to inviscid acoustic wave dominated oscillations when the location of downstream boundary varies from 8.66 to 134.7 throat height. The existence of a suction slot in the experimental setup obscures the physical downstream boundary and, therefore, presents a difficulty for quantitative comparisons between computation and experiment.

Hsieh, T.↗

NASA NTRS1987

Numerical simulation of viscous flows over transonic aircraft configurations

Numerical simulation of compressible viscous flow fields is performed for a transonic transport configuration. A single structured grid system is constructed using analytical transformations such as conformal mapping, shearing/twisting/rotating/clustering/stretching transformations. The Reynolds-averaged, thin-layer Navier-Stokes equations are solved on a supercomputer, FACOM VP-400, using the LU-ADI factorization method.

Takanashi, Susumu↗

NASA NTRS1987

Explicit finite-volume time-marching calculations of total temperature distributions in turbulent flow

Nicholson, Stephen↗

NASA NTRS1987

An explicit finite-volume time-marching procedure for turbulent flow calculations

Nicholson, Stephen↗

NASA NTRS1990

Implicit and explicit computations of flows past cavities with and without yaw

Implicit and explicit computations are presented to study the supersonic, turbulent flows past three-dimensional, rectangular cavities at 0 and 45 deg yaw angle. The Reynolds-averaged, unsteady, compressible, complete Navier-Stokes equations are solved time-accurately, including the Reynolds stresses. The number of major vortices inside and the separation characteristics on the cavity floor are different for cavities with length-to-depth ratios of 3.0 and 6.7. Due to the oscillations of the shear layer, which randomly exposes the cavity flow to the external flow, vortices are shed into the mainstream. When the upstream flow approaches a cavity at yaw, the width-to-depth ratio also becomes an important parameter. The dominant rotational planes of the major vortices incline at an angle to the vertical plane due to yaw. The execution of the implicit scheme is about three times faster, but it requires about four times more computer memory.

Baysal, Oktay↗

NASA NTRS1989

A numerical analysis of the British Experimental Rotor Program blade

Two Computational Fluid Dynamic codes which solve the compressible full-potential and the Reynolds-Averaged Thin-Layer Navier-Stokes equations were used to analyze the nonrotating aerodynamic characteristics of the British Experimental Rotor Program (BERP) helicopter blade at three flow regimes: low angle of attack, high angle of attack and transonic. Excellent agreement was found between the numerical results and experiment. In the low angle of attack regime, the BERP had less induced drag than a comparable aspect ratio rectangular planform wing. At high angle of attack, the blade attained high-lift by maintaining attached flow at the outermost spanwise locations. In the transonic regime, the BERP design reduces the shock strength at the outer spanwise locations which affects wave drag and shock-induced separation. Overall, the BERP blade exhibited many favorable aerodynamic characteristics in comparison to conventional helicopter rotor blades.

Duque, Earl P. N.↗

NASA NTRS1981

Laser velocimetry applied to transonic flow past airfoils

Measurements of turbulent Reynolds stresses and mean velocities obtained by laser velocimetry for cases of transonic shock-induced separation are presented. These results are compared with solutions of the Reynolds-averaged time-independent compressible Navier-Stokes equations. Insufficient agreement between predictions and experimental results is attributed to the deficiencies in the models for the turbulence Reynolds stresses.

Johnson, D. A.↗

NASA NTRS1990

Computing interference flows past cylinder-fin-sting assembly in and near a cavity

The internal carriage of stores by the military aircraft is an option, particularly at high speeds, for a possible reduction in the aerodynamic drag and in the radar observability. Trade studies of this option, which include the safe separation of stores from their bays and aerodynamic complications of a cavity flow, require studying the aircraft and the stores together. With this motivation, the computer code, VISCC, is developed to solve the three-dimensional, Reynolds-averaged, unsteady, compressible, and complete Navier-Stokes equations. The store considered here for simulation is a missile configuration with four tail fins and an L-shaped offset sting. The missile is placed first directly above the opening of a rectangular box cavity, then submerged completely inside the cavity. The freestream Mach and the unit Reynolds numbers are 2.75 and 2.97 million per foot, respectively. The results of the time-accurate computations depict these complex flows and help understanding the interference effects. The time-averaged surface pressures compare favorably with the available wind tunnel data.

Baysal, O.↗

NASA NTRS1986

Transonic airfoil calculations including wind tunnel wall-interference effects

The results of Reynolds-averaged time-dependent inviscid and turbulent compressible Navier-Stokes computations using the implicit finite-difference approach of Steger (1978), modified by incorporating a pressure boundary condition, (PBC) to account for wall interference are compared with experimental data on a NACA 64A010 airfoil (Johnson and Bachalo, 1980) in graphs and briefly characterized. The computational approach is the same as that used by King and Johnson (1980), but a 137 x 50 mesh is used instead of a 97 x 35 mesh, and special care is taken in resolving the nose, shock, and trailing-edge regions. Imposition of PBC is shown to improve significantly the accuracy of the computations for the flowfield on the upper surface of the airfoil, shifting the shock forward to its experimentally measured position in the case of turbulent flow. The failure of the method, even with PBC, to match the experimental shock location in the case of a flow with a separation bubble is attributed to inadequacies in the algebraic turbulence model employed (Baldwin and Lomax, 1978).

King, L. S.↗

NASA NTRS1990

Multi-processing on supercomputers for computational aerodynamics

The MIMD concept is applied, through multitasking, with relatively minor modifications to an existing code for a single processor. This approach maps the available memory to multiple processors, exploiting the C-FORTRAN-Unix interface. An existing single processor algorithm is mapped without the need for developing a new algorithm. The procedure of designing a code utilizing this approach is automated with the Unix stream editor. A Multiple Processor Multiple Grid (MPMG) code is developed as a demonstration of this approach. This code solves the three-dimensional, Reynolds-averaged, thin-layer and slender-layer Navier-Stokes equations with an implicit, approximately factored and diagonalized method. This solver is applied to a generic, oblique-wing aircraft problem on a four-processor computer using one process for data management and nonparallel computations and three processes for pseudotime advance on three different grid systems.

Yarrow, Maurice↗

NASA NTRS1979

Developments in the computation of turbulent boundary layers

Computational techniques applicable to turbulent boundary layers are classified into solutions of Reynolds-averaged equations, in which all the effects of the turbulence are modelled, and solutions of three-dimensional, time dependent Navier-Stokes equations, in which the large eddies are calculated and only the turbulence at scales smaller than the computational mesh spacings has to be modelled. Current computation costs place engineering computations in the first of these categories; large eddy simulations are appropriate currently for special studies of the dynamical processes of turbulence in idealized flow fields. It is shown that the two methods are interrelated and that each can gain from advances in the other. The degree of success of a pair of increasingly complex Reynolds stress models to broaden their range of applicability is examined through comparisons with experimental data for a variety of flow conditions. An example of a large-eddy simulation is presented, compared with experimental results, and used to evaluate the models for pressure rate-of-strain correlations and dissipation in the Reynolds-averaged equations.

Rubesin, M. W.↗