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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.

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At least 469 records · Page 26
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Numerical investigation of separated transonic turbulent flows with a multiple-time-scale turbulence model

A numerical investigation of transonic turbulent flows separated by curvature and shock wave - boundary layer interaction is presented. The free stream Mach numbers considered are 0.4, 0.5, 0.6, 0.7, 0.8, 0.825, 0.85, 0.875, 0.90, and 0.925. In the numerical method, the conservation of mass equation is replaced by a pressure correction equation for compressible flows and thus incremental pressure is solved for instead of density. The turbulence is described by a multiple-time-scale turbulence model supplemented with a near-wall turbulence model. The present numerical results show that there exists a reversed flow region at all free stream Mach numbers considered whereas various k-epsilon turbulence models fail to predict such a reversed flow region at low free stream Mach numbers. The numerical results also show that the size of the reversed flow region grows extensively due to the shock wave - turbulent boundary layer interaction as the free stream Mach number is increased. These numerical results show that the turbulence model can resolve the turbulence field subjected to extra strains caused by the curvature and the shock wave - turbulent boundary layer interaction and that the numerical method yields a significantly accurate solution for the complex compressible turbulent flow.

Kim, S.-W.↗
NASA NTRS1985

On the use of wall functions as boundary conditions for two-dimensional separated compressible flows

A new and improved wall function method for compressible turbulent flows has been developed and tested. This method is applicable to attached and separated flows, to both high- and low-Reynolds number flows, and to flows with adiabatic and nonadiabatic surfaces. This wall function method has been applied to the Launder-Spalding k-epsilon two-equation model of turbulence. The tests consist of comparisons of calculated and experimental results for: (1) an axisymmetrical transonic shock-wave/boundary-wave interaction flow at low Reynolds number in an adiabatic tube, (2) an axisymmetrical high-Reynolds number transonic flow over a nonadiabatic bump, and (3) a two-dimensional supersonic high-Reynolds number flow on a nonadiabatic deflected flap. Each of these experiments had significant regions of flow separation. The calculations are performed with an implicit algorithm that solves the Reynolds-averaged Navier-Stokes equations. It is shown that the results obtained agree very well with the data for the complex compressible flows tested.

Viegas, J. R.↗
NASA NTRS1987

Confined separated-swirling flows in diffusing ducts

The objective is to establish a detailed experimental data base for evaluation of Navier-Stokes codes for confined separated flows in diffusing s-ducts. The computational thrusts include the following: (1) extension and validation of the LeRC parabolized Navier-Stokes solver, PEPSIG, into the separated flow regime using 'flare' type approximations; (2) evaluation and extensions of state-of-the-art turbulence models for confined separated flow with and without swirl; and (3) evaluation and validation of LeRC time marching 3-D Navier-Stokes code, PROTEUS, into confined separate flow regime. Various aspects of the study are presented in viewgraph form.

Anderson, Bernhard H.↗
NASA NTRS1967

Liquid propulsion

Expulsion bladders for liquid propellant systems, resonant combustion injector, and propellant flow separation

FLOW SEPARATION↗
NASA NTRS1985

Structure of the wall pressure fluctuations in a shock-induced separated turbulent flow

This paper presents the results of an experimental study of the unsteady nature of a shock wave/turbulent boundary layer interaction. The interaction was generated using an unswept compression ramp. The incoming freestream Mach number was 2.9 and the flow was separated at the corner. An array of flush mounted miniature high frequency pressure transducers was used to make multi-channel measurements of the fluctuating wall pressure within the interaction. From the present results, an overall picture of the instantaneous structure of the unsteady shock system (as inferred from the wall pressure signals) can be constructed. The flow ahead of the corner can be considered as composed of two regions, namely the 'intermittent' region where there is essentially a single leading shock which exhibits significant streamwise 'flapping' and spanwise 'rippling', and the separated region where the flow experiences continuous compression.

Muck, K. C.↗
NASA NTRS1978

Thin-layer approximation and algebraic model for separated turbulent flows

An algebraic turbulence model for two- and three-dimensional separated flows is specified that avoids the necessity for finding the edge of the boundary layer. Properties of the model are determined and comparisons made with experiment for an incident shock on a flat plate, separated flow over a compression corner, and transonic flow over an airfoil. Separation and reattachment points from numerical Navier-Stokes solutions agree with experiment within one boundary-layer thickness. Use of law-of-the-wall boundary conditions does not alter the predictions significantly. Applications of the model to other cases are contained in companion papers.

Baldwin, B.↗
NASA NTRS1992

Maneuvering technology for advanced fighter aircraft

The need for increased maneuverability has its genesis from the first aerial combat engagement when two adversaries entangled themselves in a deadly aerial dance trying to gain the advantage over the other. It has only been in the past two decades that technologies have been investigated to increase aircraft control at maneuver attitudes that are typically dominated by highly separated flows. These separated flow regions are aggravated by advanced fighter aircraft shapes required to defeat an electronic enemy. This paper discusses passive and active devices that can be used to enhance the maneuverability of advanced fighter aircraft through vortex flow control, boundary layer control, and innovative flow manipulation.

Alexander, Michael G.↗
NASA NTRS1986

Computation of separation ahead of blunt fin in supersonic turbulent flow

Separation ahead of a flat-face blunt fin in a supersonic turbulent boundary layer was studied numerically. The following observations and conclusions were made: (1) the length of separation increases to about 5.2 D, compared with about 2.0 to 2.5 D for the typical hemi-cylindrical results, and this numerical result confirms experimental observation; (2) even though there is a kink in pressure in the present case, there is no secondary separation under the main horseshoe vortices and there are three vortices, leading to the conclusion that the number of vortices is not always an even number; and (3) for the case investigated the separation point is connected to the inner (first) horseshoe vortex, rather than the outer (second) one. The four layers of fluid entrain in the three vortices, respectively.

Hung, Ching-Mao↗
NASA NTRS1977

Computation of transonic boattail flow with separation

The relaxation procedure of South and Jameson for the full potential transonic flow equation was coupled to a modified Reshotko-Tucker integral boundary-layer technique with an empirical model for separated flow. The viscous and inviscid flows were solved iteratively until convergence was obtained. This iterative method was then applied to the subsonic and transonic flow over a series of axisymmetric circular-arc boattails with solid jet plume simulators. Comparisons of theoretical and experimental surface pressures and boattail drag are presented over a free-stream Mach numbers below 0.90. The qualitative variation of boattail drag with free-stream Mach number and boattail angle well into the region of transonic drag rise was correctly predicted; however, the absolute drag levels were significantly underpredicted. For separated flows, the empirical discriminating streamline model gives good results up to a free-stream Mach number of about 0.90 and allows reasonable predictions for shock-induced separation if the proper separation location and separation turning angle are known.

Wilmoth, R. G.↗
NASA NTRS1991

Examination of a vortex-ring interaction phenomenon in an axisymmetric flow

Unsteady separated flow in an axisymmetric configuration is investigated, with the objective of analyzing vortex-ring formation and break-up and the ensuing interactions believed to be leading to breakdown of laminar flow. The model problem considered is an idealized representation of a combustor with a centerbody. Direct solution of the time-dependent incompressible Navier-Stokes equations is employed, with central-differencing for all spatial derivatives, to determine the response of this flow for Re = 2000, 5000 and 10,000. The last case revealed a rather novel unsteady vortex-ring interaction phenomenon. This case has been further investigated using third-order accurate upwind differencing for the convective terms. The flow is interrogated carefully by examining the corresponding instantaneous vorticity, stream-function and velocity fields, as well as by tracking mass-less marker particles in some of the critical regions. Such detailed examination is deemed essential, prior to assigning fundamental physical relevance to the observed vortex-interaction phenomenon.

Ghia, U.↗
NASA NTRS1985

The structure of dilute combusting sprays

An experimental and theoretical study of drop processes in a turbulent flame is described. The experiments involved a monodisperse (105 and 180 micro m initial diameter) stream of methanol drops injected at the base of a turbulent methane-fueled diffusion flame burning in still air. The following measurements were made: mean and fluctuating phase velocities, mean drop number flux, drop-size distributions and mean gas-phase temperatures. Measurements were compared with predictions of two separated flow models: (1) deterministic separated flow, where drop-turbulence interactions are ignored; and (2) stochastic separated flow, where drop-turbulence interactions are considered using random-walk computations. The stochastic separated flow analysis yielded best agreement with measurements, since it provides for turbulent dispersion of drops which was important for present test conditions (and probably for most combusting sprays as well). Distinguishing the presence or absence of envelope flames around the drops, however, was relatively unimportant for present test conditions, since the drops spent most of their lifetime in fuel-rich regions of the flow where this distinction is irrelevant.

Shuen, J. S.↗
NASA NTRS1982

Analysis of separated boundary-layer flows

A method for predicting the strong interaction between the viscous and inviscid flows which occurs in flow separation is reviewed. An inverse boundary-layer procedure approximately accounts for normal pressure gradients that may be important in strongly interacting flows. Transformed boundary-layer equations are written in which the pressure gradient is set equal to the inviscid pressure gradient. As the boundary-layer edge is approached and the viscous shear and heat conduction terms vanish, the viscous flow solution is required to asymptotically approach the inviscid solution over the generalized displacement body. Attention is then focused on viscous-inviscid interacting flows with a first-order viscous formulation and constant pressure across the boundary layer. Results obtained with this procedure are presented for: (1) transitional separation bubbles near an airfoil leading edge, (2) subsonic boattail separated turbulent flow, and (3) transonic turbulent shock wave boundary-layer interaction on an axisymmetric bump configuration.

Carter, J. E.↗
NASA NTRS1986

Upwind Navier-Stokes solutions for separated periodic flows

The application of an upwind implicit approximate factorization Navier-Stokes algorithm to highly separated flow is described. Using both the thin-layer and complete forms of the Navier-Stokes equations, the low Reynolds number laminar flow around a circular cylinder with periodic shedding is solved. The effect of grid density, grid extent, and time step on the Strouhal number is shown. Results from both sets of equations agree within the experimental data band. Unsteady, laminar flow computations around inclined plates and separated airfoils are also described. Strouhal numbers agree to within 5 percent of experiments for inclined plates. Differences between the complete equations and the thin-layer approximation for separated periodic flows are discussed. Computations of an impulsively started circular cylinder and airfoil yield time-accurate flowfield shapes in good agreement with experimental flow visualizations. The turbulent computation of an airfoil at a high angle-of-attack is massively separated, but shows no evidence of periodicity.

Rumsey, C. L.↗
NASA NTRS1985

Drop-turbulence interactions in a diffusion flame

A monodisperse stream of methanol drops injected along the axis of a turbulent, methane-fueled diffusion flame burning in still air is studied experimentally and theoretically, in order to determine mean and fluctuating phase velocities, mean drop number flux, drop size distributions, and mean gas phase temperature. Measured values were compared with the predictions of two separated flow analyses, namely deterministic separated flow and stochastic separated flow. The stochastic analysis yielded the best agreement with measurements, due to its providing for the turbulent dispersion of drops.

Shuen, J.-S.↗
NASA NTRS1975

Internal cowl-separation at high incidence angles

Internal flow separation at large incidence angles is studied on the basis of wind-tunnel pressure data for six axisymmetric inlet geometries. The inlet geometric variables investigated are the angle of incidence, the throat Mach number, the internal lip contraction ratio, 'sharpness' of the external shape, and freestream velocity. It is shown that an increase in lip contraction ratio delays internal flow separation, while an increase in the sharpness of the external profile tends to reduce the angle of incidence at which complete flow separation occurs. It is also shown that inlet flow separation is subject to scale effects associated with the boundary layer on the cowl surface. The scale effects are particularly pronounced at very high throat Mach numbers.

Jakubowski, A. K.↗
NASA NTRS1994

Towards Understanding the Fluid Dynamic Phenomenon of Interest to Rocket Base Heating: A Review

The significance of the base heating problem for rockets during ascent is due to the complex interaction between the rocket nozzle plumes and the external-flow which can change the flow characteristics in the base region dramatically. At lower altitudes the external-flow merges with the plume-flow, without the formation of a large separated flow region, and the cooler external-flow promotes convective cooling of the base wall. Under these conditions the majority of the base heating is due to radiative heating from the shock heated plume gases. At higher altitudes, however, the process of base heating is not so straightforward. The plume and the base flow expands dramatically and separated flow regions occur in the base area. Hot exhaust gases from the rocket nozzle will be entrained into the separated flow regions and produce a convective component to the base wall heating. Further, if the rocket exhaust-gas contains soot, the soot can increase the emission from the gas and dramatically increase the wall absorption coefficient for radiative heating if it is deposited on the walls . In addition, if the rocket exhaust gas is fuel rich, the fuel can bum in the separated flow regions and further increase the base heating. The base burning phenomenon, and the increased base heating caused by it at higher altitudes, have been observed for the Space Shuttle and Saturn Rocket. Under these conditions, the total heating is significantly higher than the heating without separated flow in the base region, and the increase in heating is directly attributable to the fluid dynamic complexity of the base region. Realistic simulation of the base heating requires that the calculated flow environment reproduce the fluid dynamic flow features accurately. Thus, it will be necessary to introduce into the CFD codes the capability for the flow to respond to the complex vehicle geometry, the effect of turbulence, the ability to accurately reproduce the plume shock/shear layer structures and capture such phenomenon as shock induced base separation and base-burning phenomenon. A survey of experimental, theoretical and computational work that details the fluid dynamics of the base flow environment will be presented in the proposed paper. CFD simulations of rocket base flows using standard CFD codes such as OVERFLOW or GASP will be explored to capture these phenomenon accurately. Merits and limitations of these codes for base flow environment predictions will be explored.

Venkatapathy, E.↗