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At least 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

A three-dimensional Navier-Stokes stage analysis of the flow through a compact radial turbine

A steady, three-dimensional Navier-Stokes average passage computer code is used to analyze the flow through a compact radial turbine stage. The code is based upon the average passage set of equations for turbomachinery, whereby the flow fields for all passages in a given blade row are assumed to be identical while retaining their three-dimensionality. A stage solution is achieved by alternating between stator and rotor calculations, while coupling the two solutions by means of a set of axisymmetric body forces which model the absent blade row. Results from the stage calculation are compared with experimental data and with results from an isolated rotor solution having axisymmetric inlet flow quantities upstream of the vacated stator space. Although the mass-averaged loss through the rotor is comparable for both solutions, the details of the loss distribution differ due to stator effects. The stage calculation predicts smaller spanwise variations in efficiency, in closer agreement with the data. The results of the study indicate that stage analyses hold promise for improved prediction of loss mechanisms in multi-blade row turbomachinery, which could lead to improved designs through the reduction of these losses.

Heidmann, James D.↗
NASA NTRS1996

Solutions to Three-Dimensional Thin-Layer Navier-Stokes Equations in Rotating Coordinates for Flow Through Turbomachinery

The viscous, Navier-Stokes solver for turbomachinery applications, MSUTC has been modified to include the rotating frame formulation. The three-dimensional thin-layer Navier-Stokes equations have been cast in a rotating Cartesian frame enabling the freezing of grid motion. This also allows the flow-field associated with an isolated rotor to be viewed as a steady-state problem. Consequently, local time stepping can be used to accelerate convergence. The formulation is validated by running NASA's Rotor 67 as the test case. results are compared between the rotating frame code and the absolute frame code. The use of the rotating frame approach greatly enhances the performance of the code with respect to savings in computing time, without degradation of the solution.

Ghosh, Amrit Raj↗
NASA NTRS1988

A streamwise upwind algorithm for the Euler and Navier-Stokes equations applied to transonic flows

A new algorithm was developed for the Euler and Navier-Stokes equations that uses upwind differencing based on the streawise direction. This algorithm is time accurate and can be used in codes for calculating unsteady transonic flows over wings. Such codes can be used for the flutter analysis of wings. In this algorithm, the coordinate system is locally rotated to align with the streamwise direction. For differencing the convective terms in the streamwise direction, a new form of flux splitting is employed, in which the biasing depends on the local Mach number. In the plane perpendicular to the stream direction, the new flux splitting uses the condition of no flow in that local plane. By using a locally rotated coordinate system, the convective flux vector biasing depends on the total Mach number. Hence, the switching of the flux vector biasing occurs across shock waves and the proper domain of dependence is used in supersonic regions. For comparison, many other upwind methods switch differencing based on Mach number of shock waves in multidimensional flows. The formulas for the convective flux vector differencing do not contain any user specified parameters. So, the amount of numerical dissipation is automatically determined.

Goorjian, Peter M.↗
NASA NTRS1988

A streamwise upwind algorithm for the Euler and Navier-Stokes equations applied to transonic flows

A new algorithm was developed for the Euler and Navier-Stokes equations that uses upwind differencing based on the streamwise direction. This algorithm is time accurate and can be used in codes for calculating unsteady transonic flows over wings. Such codes can be used for the flutter analysis of wings. In this algorithm, the coordinate system is locally rotated to align with the streamwise direction. For differencing the convective terms in the streamwise direction, a new form of flux splitting is employed, in which the biasing depends on the local Mach number. In the plane perpendicular to the stream direction, the new flux splitting uses the condition of no flow in that local plane. By using a locally rotated coordinate system, the convective flux vector biasing depends on the total Mach number. Hence, the switching of the flux vector biasing occurs across shock waves and the proper domain of dependence is used in supersonic regions. For comparison, many other upwind methods switch differencing based on Mach number of shock wave in multidimensional flows. The formulas for the convective flux vector differencing do not contain any user specified parameters. So, the amount of numerical dissipation is automatically determined.

Goorjian, P. M.↗
NASA NTRS1987

A new algorithm for the Navier-Stokes equations applied to transonic flows over wings

A new algorithm has been developed for the Navier-Stokes equations. For the convective terms, this algorithm employs flux vector splitting together with a locally rotated coordinate system that is aligned with the streamwise direction. The algorithm has been implemented into a Navier-Stokes code, the NASA Ames Research Center's TNS code, and several cases of steady flow have been calculated. The results show improvements over the original method in the code. First, in a case of separated flow, there is better agreement with the experimental results and also there is an absence of numerical oscillations that occur with the original method. Second, in a case with a strong supersonic to subsonic shock wave, comparisons show an improvement in the resolution of the shock wave and the reexpansion singularity.

Goorjian, Peter M.↗
NASA NTRS1991

An efficient and robust algorithm for two dimensional time dependent incompressible Navier-Stokes equations: High Reynolds number flows

An algorithm is presented for unsteady two-dimensional incompressible Navier-Stokes calculations. This algorithm is based on the fourth order partial differential equation for incompressible fluid flow which uses the streamfunction as the only dependent variable. The algorithm is second order accurate in both time and space. It uses a multigrid solver at each time step. It is extremely efficient with respect to the use of both CPU time and physical memory. It is extremely robust with respect to Reynolds number.

Goodrich, John W.↗
NASA NTRS1991

An efficient and robust algorithm for two dimensional time dependent incompressible Navier-Stokes equations - High Reynolds number flows

An algorithm is presented for unsteady two-dimensional incompressible Navier-Stokes calculations. This algorithm is based on the fourth order partial differential equation for incompressible fluid flow which uses the streamfunction as the only dependent variable. The algorithm is second order accurate in both time and space. It uses a multigrid solver at each time step. It is extremely efficient with respect to the use of both CPU time and physical memory. It is extremely robust with respect to Reynolds number.

Goodrich, John W.↗
NASA NTRS1984

A compressible Navier-Stokes code for turbulent flow modeling

An implicit, finite volume code for solving two dimensional, compressible turbulent flows is described. Second order upwind differencing of the inviscid terms of the equations is used to enhance stability and accuracy. A diagonal form of the implicit algorithm is used to improve efficiency. Several zero and two equation turbulence models are incorporated to study their impact on overall flow modeling accuracy. Applications to external and internal flows are discussed.

Coakley, T. J.↗
NASA NTRS1987

Navier-Stokes computations of vortical flows over low aspect ratio wings

An upwind-biased finite-volume algorithm is applied to the low-speed flow over a low aspect ratio delta wing from zero to forty degrees angle of attack. The differencing is second-order accurate spatially, and a multigrid algorithm is used to promote convergence to the steady state. The results compare well with the detailed experiments of Hummel (1983) and others for a Re(L) of 0.95 x 10 to the 6th. The predicted maximum lift coefficient of 1.10 at thirty-five degrees angle of attack agrees closely with the measured maximum lift of 1.06 at thirty-three degrees. At forty degrees angle of attack, a bubble type of vortex breakdown is evident in the computations, extending from 0.6 of the root chord to just downstream of the trailing edge.

Thomas, J. L.↗
NASA NTRS1987

Incompressible Navier-Stokes computations of vortical flows over double-delta wings

An implicit flux-difference splitting scheme is used to compute incompressible laminar vortical flows about a thin round-edged double-delta wing with 80 deg and 60 deg leading-edged sweep for the strake and the main wing respectively. The numerical scheme combines approximate factorization in crossflow planes with a symmetric planar Gauss-Seidel relaxation in the remaining spatial direction. It is second-order accurate spatially by applying TVD-like upwind discretization to the inviscid fluxes and central differencing to the viscous shear fluxes. The interaction between the two primary vortices emanating from the apex and kink leading-edges is successfully simulated. The computed trajectories of vortical cores compare well with experimental data.

Hsu, Chung-Hao↗
NASA NTRS1988

Navier-Stokes calculations of transonic flows past open and transitional cavities

Computational simulations of deep and transitional cavities at transonic regimes were performed. The three-dimensional unsteady separation on the cavity floor was analyzed, and computational flow visualization inside the cavity was done. A parametric study was performed to investigate the effects of varying Mach number and the ratio delta/D, where delta is the instantaneous boundary layer thickness at the upstream lip and D is the depth. Both time-averaged and instantaneous solutions were obtained. Comparison of the solutions with experimental data were generally favorable; some discrepancies are explained.

Srinivasan, S.↗
NASA NTRS1989

Development and validation of CNS (compressible Navier-Stokes) for hypersonic external flows

CNS, a new computational fluid dynamics procedure, has been developed to aid in hypersonic vehicle design. The code can be used to model the entire external flow around hypersonic vehicle shapes, from the captured shock at the nose to the beginning of the wake. Unlike space-marching codes, the technique allows axially separated flow regions to be modeled. Validation trials using sphere-cone data reveal good solution accuracy for the surface pressure and flowfield temperature.

Flores, Jolen↗
NASA NTRS1989

Euler and Navier-Stokes solutions for hypersonic flows

An upwind finite-element technique that uses cell-centered quantities and implicit and/or explicit time marching has been developed for computing hypersonic laminar viscous flows using adaptive unstructured grids in two and three dimensions. A perfect gas model as well as an equilibrium air model is implemented for solving high-speed flows. A first-order basic scheme and a higher-order flux-corrected transport (FCT) scheme have been implemented. This technique has been used to predict 'Type III and IV' shock interactions on a cylinder in two dimensions and a swept cylinder in three dimensions, with a view to determine the pressure and heating rate augmentation caused by an impinging shock on the leading edge of a cowl lip of an engine inlet. The predictions of wall pressure and heating rates compare very well with experimental data. The flow features are very distinctly captured with a sequence of adaptively-generated grids. Three-dimensional corner flow, typically encountered in engine inlets due to compression of the flow by ramps in the walls, is also modeled. This procedure is the first step in developing an integrated fluid, thermal, structural analysis capability for hypersonic flight vehicles like the National Aero-Space Plane.

Thareja, Rajiv R.↗