SEARCH · Engineering Papers

Results for “Stokes flow”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 1,009 records · Page 56

All sources NASA NTRS DOE OSTI

NASA NTRS1990

Development of a defect stream function, law of the wall/wake method for compressible turbulent boundary layers

The method presented is designed to improve the accuracy and computational efficiency of existing numerical methods for the solution of flows with compressible turbulent boundary layers. A compressible defect stream function formulation of the governing equations assuming an arbitrary turbulence model is derived. This formulation is advantageous because it has a constrained zero-order approximation with respect to the wall shear stress and the tangential momentum equation has a first integral. Previous problems with this type of formulation near the wall are eliminated by using empirically based analytic expressions to define the flow near the wall. The van Driest law of the wall for velocity and the modified Crocco temperature-velocity relationship are used. The associated compressible law of the wake is determined and it extends the valid range of the analytical expressions beyond the logarithmic region of the boundary layer. The need for an inner-region eddy viscosity model is completely avoided. The near-wall analytic expressions are patched to numerically computed outer region solutions at a point determined during the computation. A new boundary condition on the normal derivative of the tangential velocity at the surface is presented; this condition replaces the no-slip condition and enables numerical integration to the surface with a relatively coarse grid using only an outer region turbulence model. The method was evaluated for incompressible and compressible equilibrium flows and was implemented into an existing Navier-Stokes code using the assumption of local equilibrium flow with respect to the patching. The method has proven to be accurate and efficient.

Wahls, Richard A.↗

NASA NTRS1990

Effect of nose bluntness on boundary layer stability and transition

The effect of nose bluntness on boundary layer instability is studied theoretically for a Mach 8 flow past a 7 degree semivertex cone. The basic flow is computed by solving the parabolized Navier-Stokes equations. Linear stability analysis of the basic flow reveals that, with small amount of bluntness, the critical Reynolds number for the onset of instability increases by an order of magnitude compared to the sharp cone value. The computed second mode frequencies are also in reasonable agreement with the experimental results. The results are used to explain the effect of unit Reynolds number on transition present in the quiet aeroballistic range data.

Malik, M. R.↗

NASA NTRS1991

Calculating Viscous/Inviscid Interactions

Report describes some experiences in numerical simulation of compressible, viscous flows by finite-difference approximations of Navier-Stokes equations. Focuses upon viscous/inviscid-interaction approach, in which flow mathematically modeled with interacting zones, within each of which different model of viscous or inviscid flow applies. Reduces task of computation by taking advantage of simpler models and coarser grids in some of zones.

Van Dalsem, W. R.↗

NASA NTRS2015

An Investigation of Transonic Resonance in a Mach 2.2 Round Convergent-Divergent Nozzle

Hot-wire and acoustic measurements were taken for a round convergent nozzle and a round convergent-divergent (C-D) nozzle at a jet Mach number of 0.61. The C-D nozzle had a design Mach number of 2.2. Compared to the convergent nozzle jet flow, the Mach 2.2 nozzle jet flow produced excess broadband noise (EBBN). It also produced a transonic resonance tone at 1200 Herz. Computational simulations were performed for both nozzle flows. A steady Reynolds-Averaged Navier-Stokes simulation was performed for the convergent nozzle jet flow. For the Mach 2.2 nozzle flow, a steady RANS simulation, an unsteady RANS (URANS) simulation, and an unsteady Detached Eddy Simulation (DES) were performed. The RANS simulation of the convergent nozzle showed good agreement with the hot-wire velocity and turbulence measurements, though the decay of the potential core was over-predicted. The RANS simulation of the Mach 2.2 nozzle showed poor agreement with the experimental data, and more closely resembled an ideally-expanded jet. The URANS simulation also showed qualitative agreement with the hot-wire data, but predicted a transonic resonance at 1145 Herz. The DES showed good agreement with the hot-wire velocity and turbulence data. The DES also produced a transonic tone at 1135 Herz. The DES solution showed that the destabilization of the shock-induced separation region inside the nozzle produced increased levels of turbulence intensity. This is likely the source of the EBBN.

DES↗

NASA NTRS1982

A new numerical approach for compressible viscous flows

A numerical approach for computing unsteady compressible viscous flows was developed. This approach offers the capability of confining the region of computation to the viscous region of the flow. The viscous region is defined as the region where the vorticity is nonnegligible and the difference in dilatation between the potential flow and the real flow around the same geometry is also nonnegligible. The method was developed and tested. Also, an application of the procedure to the solution of the steady Navier-Stokes equations for incompressible internal flows is presented.

Wu, J. C.↗

NASA NTRS1988

Navier-Stokes calculations of scramjet-afterbody flowfields

Scramjet simulant-gas exhaust flows are computed using two-dimensional Navier-Stokes (NS) equations. The NS equations are solved using an implicit, upwind, finite-volume scheme, and the Reynolds stresses are modeled algebraically. The solutions are compared with experimental data where the freestream and the nozzle exhaust fluids are air. The NS equations for the flow, where the scramjet exhaust simulant gas is a mixture of Freon-12 and argon, are solved by an explicit, finite difference scheme. The results are compared with experimental data, and certain discrepancies are explained.

Baysal, O.↗

NASA NTRS1995

Computational Study of Flow Establishment in a Ram Accelerator

The temporal evolution of the combustion process established during projectile transition from the launch tube into the ram accelerator section containing an explosive hydrogen-oxygen-argon gas mixture is studied. The Navier-Stokes equations for chemically reacting flow are solved in a fully coupled manner, using an implicit, time accurate algorithm. The solution procedure is based on a spatially second order total variation diminishing scheme and a temporally second order, variable-step, backward differentiation formula method. The hydrogen-oxygen chemistry is modeled with a 9-species, 19-step mechanism. The accuracy of the solution method is first demonstrated by several benchmark calculations. Numerical simulations of two ram accelerator configurations are then presented. In particular, the temporal developments of shock-induced combustion and thrust forces are followed. Positive thrust is established in both cases; however, in one of the ram accelerator configurations studied, combustion in the boundary layer enhances its separation, ultimately resulting in unstart.

Yungster, S.↗

NASA NTRS1992

Numerical simulation of steady and unsteady asymmetric vortical flow

The unsteady, compressible, thin-layer, Navier-Stokes (NS) equations are solved to simulate steady and unsteady, asymmetric, vortical laminar flow around cones at high incidences and supersonic Mach numbers. The equations are solved by using an implicit, upwind, flux-difference splitting (FDS), finite-volume scheme. The locally conical flow assumption is used and the solutions are obtained by forcing the conserved components of the flowfield vector to be equal at two axial stations located at 0.95 and 1.0. Computational examples cover steady and unsteady asymmetric flows around a circular cone and its control using side strakes. The unsteady asymmetric flow solution around the circular cone has also been validated using the upwind, flux-vector splitting (FVS) scheme with the thin-layer NS equations and the upwind FDS with the full NS equations. The results are in excellent agreement with each other. Unsteady asymmetric flows are also presented for elliptic- and diamond-section cones, which model asymmetric vortex shedding around round- and sharp-edged delta winds.

Kandil, Osama A.↗

NASA NTRS1997

Heat Transfer Computations of Internal Duct Flows With Combined Hydraulic and Thermal Developing Length

This study investigated the Navier-Stokes computations of the surface heat transfer coefficients of a transition duct flow. A transition duct from an axisymmetric cross section to a non-axisymmetric cross section, is usually used to connect the turbine exit to the nozzle. As the gas turbine inlet temperature increases, the transition duct is subjected to the high temperature at the gas turbine exit. The transition duct flow has combined development of hydraulic and thermal entry length. The design of the transition duct required accurate surface heat transfer coefficients. The Navier-Stokes computational method could be used to predict the surface heat transfer coefficients of a transition duct flow. The Proteus three-dimensional Navier-Stokes numerical computational code was used in this study. The code was first studied for the computations of the turbulent developing flow properties within a circular duct and a square duct. The code was then used to compute the turbulent flow properties of a transition duct flow. The computational results of the surface pressure, the skin friction factor, and the surface heat transfer coefficient were described and compared with their values obtained from theoretical analyses or experiments. The comparison showed that the Navier-Stokes computation could predict approximately the surface heat transfer coefficients of a transition duct flow.

Wang, C. R.↗

NASA NTRS1980

Numerical simulation of flow over wavy walls

Spectral methods are used to solve the Navier-Stokes equations numerically for two-dimensional incompressible flow over wavy surfaces. Results obtained for high Reynolds number laminar flows and turbulent flows are compared to experiment and analytical theory. Overall, good agreement has been achieved.

Balasubramanian, R.↗

NASA NTRS1986

Numerical simulation of vortical flow over an elliptical-body missile at high angles of attack

Numerical solutions to the Reynolds-averaged Navier-Stokes equations are given for the flow about an elliptical body missile (3:1 ellipse) at a Mach number of 2.5 and a unit Reynolds number of 6.56 x 10 to the 6th/m. At high angles of attack, the flow is dominated by large-scale free vortices which occur in the lee-side flow field due to crossflow boundary-layer separation. Emphasis is focused on the accurate prediction of the lee-side vortical flow. Solutions are presented for both symmetric and asymmetric (body rolled 45 deg) configurations at 10 deg and 20 deg angle of attack. The computed results are compared with experimental surface pressure coefficients and vapor-screen photographs. Excellent agreement is obtained in all cases.

Newsome, R. N.↗

NASA NTRS1989

Numerical simulation of high-incidence flow over the F-18 fuselage forebody

As part of the NASA High Alpha Technology Program, fine-grid Navier-Stokes solutions have been obtained for flow over the fuselage forebody and wing leading-edge extension of the F/A-18 High Alpha Research Vehicle at large incidence. The resulting flows are complex and exhibit cross-flow separation from the sides of the forebody and from the leading-edge extension. A well-defined vortex pattern is observed in the leeward-side flow. Results obtained for laminar flow show good agreement with flow visualizations obtained in ground-based experiments. Further, turbulent flows computed at high-Reynolds-number flight-test conditions show good agreement with surface and off-surface visualizations obtained in flight.

Schiff, Lewis B.↗

NASA NTRS2003

Numerical Simulation of Turbulent MHD Flows Using an Iterative PNS Algorithm

A new parabolized Navier-Stokes (PNS) algorithm has been developed to efficiently compute magnetohydrodynamic (MHD) flows in the low magnetic Reynolds number regime. In this regime, the electrical conductivity is low and the induced magnetic field is negligible compared to the applied magnetic field. The MHD effects are modeled by introducing source terms into the PNS equation which can then be solved in a very efficient manner. To account for upstream (elliptic) effects, the flowfields are computed using multiple streamwise sweeps with an iterated PNS algorithm. Turbulence has been included by modifying the Baldwin-Lomax turbulence model to account for MHD effects. The new algorithm has been used to compute both laminar and turbulent, supersonic, MHD flows over flat plates and supersonic viscous flows in a rectangular MHD accelerator. The present results are in excellent agreement with previous complete Navier-Stokes calculations.

Kato, Hiromasa↗

NASA NTRS1978

Semidirect calculation of steady two- and three-dimensional flows

This paper describes a semidirect method for rapidly solving steady-state viscous flows described by the complete Navier-Stokes equations. The current results are for two-dimensional incompressible flows in general channels at arbitrary Reynolds numbers, but work in progress on compressible and three-dimensional flows is also described. The basic concept of semidirect methods is to use the recently developed fast (direct, or noniterative) linear solvers to solve linearized equations, which are then iterated to solve the nonlinearity. The method used here is an extension of the Split NOS method (Roache, 1975)

Roache, P. J.↗

NASA NTRS1981

Numerical solution of supersonic viscous flow over blunt delta wings

A general parabolized Navier-Stokes code has been developed to compute the steady supersonic viscous flow around arbitrary body shapes at high angles of attack. A nonorthogonal three-dimensional coordinate frame permits the code to march with solution surfaces which are the most appropriate. The code has been used to calculate the laminar flow over a slab delta wing with 70 deg sweep at angles of attack up to 41.5 deg and Mach numbers of 6.8 and 9.6. The computed shock shapes, surface pressures and heat transfer coefficients are compared with experiment and show agreement.

Tannehill, J. C.↗

NASA NTRS1979

Supersonic viscous flow over cones at incidence

The conically symmetric Navier-Stokes equations are solved by MacCormack's method for the supersonic flow past sharp cones at incidence. To provide closure for the case of turbulent flows, a scalar eddy-viscosity model based on mixing length hypotheses is used. The results are compared with available experimental data.

Mcrae, D. S.↗

NASA NTRS1989

Numerical simulation of separated and vortical flows on bodies at large angles of attack

Navier-Stokes computations were carried out for high-angle-of-attack flow about bodies of revolution in supersonic and subsonic flow. Numerical and physical factors affecting flow predictions are discussed. Accurate prediction of the complex three-dimensional separated flow is found to depend on using grids sufficiently fine to resolve the details of both the viscous boundary layer and the off-surface separated flow structures, and, for turbulent flow cases, use of an eddy-viscosity turbulence model which accounts for the leeward vortical flow structures. An example of high-incidence flow computed about a more complex geometry, the F-18 fighter fuselage forebody and leading edge extension, is also presented.

Schiff, Lewis B.↗

NASA NTRS2005

Numerical Prediction of Chevron Nozzle Noise Reduction using Wind-MGBK Methodology

Numerical predictions for single-stream chevron nozzle flow performance and farfield noise production are presented. Reynolds Averaged Navier Stokes (RANS) solutions, produced via the WIND flow solver, are provided as input to the MGBK code for prediction of farfield noise distributions. This methodology is applied to a set of sensitivity cases involving varying degrees of chevron inward bend angle relative to the core flow, for both cold and hot exhaust conditions. The sensitivity study results illustrate the effect of increased chevron bend angle and exhaust temperature on enhancement of fine-scale mixing, initiation of core breakdown, nozzle performance, and noise reduction. Direct comparisons with experimental data, including stagnation pressure and temperature rake data, PIV turbulent kinetic energy fields, and 90 degree observer farfield microphone data are provided. Although some deficiencies in the numerical predictions are evident, the correct farfield noise spectra trends are captured by the WIND-MGBK method, including the noise reduction benefit of chevrons. Implications of these results to future chevron design efforts are addressed.

Engblom, W.A.↗