Navier-Stokes Code Research Articles

Installed F/A-18A inlet flow calculations - A grid study

Full Navier-Stokes calculations on the F/A-18A High Alpha Research Vehicle inlet at 30 deg angle of attack, 0 deg yaw, and a freestream Mach number of 0.2, have been obtained in this study using an algebraic turbulence model with two grids (original and revised). In addition, a limited set of results were obtained using a thinlayer Navier-Stokes (TLNS) code with the revised grid. The limitedness of the results was due to the inadequacy of the grid to meet the requirements of the TLNS code. Results obtained with the original grid were used to determine where further grid refinements and additional geometry were needed. The calculations are compared to a limited amount of available experimental data. The predicted forebody/fuselage surface static pressures compared well with data for all solutions. The predicted trajectory of the vortex generated under the leadingedge extension was different for each solution. These discrepancies are attributed to differences in the grid resolution, turbulence modeling, artificial dissipation, and differencing schemes. All solutions predict that this vortex is ingested by the inlet. The predicted 40-probe rake inlet total pressure recoveries are lower than data and the distortions are higher than data. The results obtained with the revised grid were significantly improved from the original grid results.

Navier-Stokes computations and experimental comparisons for multielement airfoil configurations

Navier-Stokes computations and experimental comparisons for multielement airfoil configurations

Prediction of nonequilibrium turbulent flows with explicit algebraic stress models

An explicit algebraic stress equation, developed by Gatski and Speziale, is used in the framework of the K-UJ and K-e formulations to predict two-dimensional separated turbulent flows. The nonequilibrium effects are modeled through coefficients that depend nonlinearly on both rotational and irrotational strains. The proposed models were implemented in the Courant-Friedrichs-Lewy three-dimensional Navier-Stokes code. Comparisons with the experimental data are presented, which clearly demonstrate that explicit algebraic stress models improve the response of standard two-equation models to nonequilibrium effects. I. Introduction

Endwall and Unsteady Flow Phenomena in an Axial Turbine Stage

Detailed experimental and numerical studies have been performed in a subsonic, axial-flow turbine stage to investigate the secondary flow field, the aerodynamic loss generation, and the spanwise mixing under a stage environment. The experimental study includes measurements of the static pressure distribution on the rotor blade surface and the rotor exit flow field using three-dimensional hot-wire and pneumatic probes. The rotor exit flow field was measured with an unsteady hot-wire probe, which has high temporal and spatial resolution. Both steady and unsteady numerical analyses were performed with a three-dimensional Navier–Stokes code for the multiple blade rows. Special attention was focused on how well the steady multiple-blade-row calculation predicts the rotor exit flow field and how much the blade interaction affects the radial distribution of flow properties at the stage exit. Detailed comparisons between the measurement and the steady calculation indicate that the steady multiple-blade-row calculation predicts the overall time-averaged flow field very well. However, the steady calculation does not predict the secondary flow at the stage exit accurately. The current study indicates that the passage vortex near the hub of the rotor is transported toward the midspan due to the blade interaction effects. Also, the structure of the secondary flow field at the exit of the rotor is significantly modified by the unsteady effects. The time-averaged secondary flow field and the radial distribution of the flow properties, which are used for the design of the following stage, can be predicted more accurately with the unsteady flow calculation.

Hypersonic shockwave/turbulent boundary-layer interactions on a porous surface

The passive control of hypersonic shock/boundary-layer interactions is numerically studied. A previously developed Navier-Stokes code is applied, after modifications to the code to account for passive control have been implemented.

A Lagrangian-Eulerian scheme for flow around an airfoil in rain

In the interest of assessing airfoil performance in rain, a Lagrangian particle tracking algorithm for a general body-fitted co-ordinate system has been developed and linked with a thin layer incompressible Navier-Stokes code. Non-deforming spherical particles are tracked through the two-dimensional, incompressible air flow field surrounding a NACA 64-210 airfoil section. Each tracked particle represents a distribution of raindrops. Impacts on the airfoil surface and the resulting splashback are modeled, and the steady state fluid field and droplet distribution are determined utilizing an iterative, two-way momentum coupled approach. Details of the splashback effect on the boundary layer are examined. A 1–2° rain-induced decrease in stall angle of attack is predicted.

Numerical Analysis of Nozzle and Afterbody Flow of Hypersonic Transport Systems

Using an implicit Finite-Volume Navier–Stokes code, the flow field in a Single Expansion Ramp Nozzle (SERN) for a hypersonic aircraft is studied. Comparisons between experimental data and CFD calculations for certain components of the integrated exhaust system (cold two-dimensional nozzle flow, high temperature reacting three-dimensional combustion chamber flow, and two-dimensional nozzle flow with external flow) are presented. To show the sensitivity of the considered components to off-design operating conditions, comprehensive numerical studies have been carried out. For the determination of nozzle performance a detailed two-dimensional analysis from transonic to hypersonic flight Mach numbers has been performed. A direct optimization method has been used to investigate the influence of the lower nozzle flap shape on the thrust vector.

Modeling of Nitrogen and Oxygen Recombination on Partial Catalytic Surfaces

In connection with recombination coefficients derived from experimental data described in the literature, a reaction scheme including detailed rate expressions for O and N atom recombination on the surface of re-entry vehicles is established, consisting of elementary reaction steps. To validate the reaction mechanism derived, surface chemistry and fluid mechanical processes are coupled assuming a one-dimensional stagnation flow field. A quantitative agreement is achieved between recombination coefficients resulting from the numerical computations and those calculated from experiments. The temperature dependence of the recombination coefficient is explained by elementary reaction steps. Furthermore, the reaction scheme established is implemented in a two-dimensional Navier–Stokes code computing the re-entry flow around a simple geometry to show the importance of a detailed modeling of surface reactions.

3-D Navier-Stokes analysis of crossing glancing shocks/turbulent boundary layer interactions

Three-dimensional viscous flow analysis is performed for a configuration where two crossing and glancing shocks interact with a turbulent boundary layer. A time marching 3-D full Navier—Stokes code, called PARC3D, is used to compute the flow field and the solution is compared to the experimental data obtained at the NASA Lewis Research Center 1 ft × 1 ft supersonic wind tunnel facility. The study is carried out as part of the continuing code assessment program in support of the Generic Hypersonic Research at NASA Lewis. Detailed comparison of static pressure fields and oil flow patterns is made with the corresponding solution on the wall containing the shock/boundary layer interaction in an effort to validate the code for hypersonic inlet applications.

Current distribution on isothermal rod electrodes in combustion MHD generators

'Drummond, J. P., and Hussaini, M. Y., A Detailed Numerical Model of a Supersonic Reacting Mixing Layer, AIAA Paper 861427, June 1986. Riggins, D. W., and McClinton, C. R., A Computational Investigation of Flow Losses in a Supersonic Combustor, AIAA Paper 90-2093, July 1990. Eklund, D. R., Northam, G. B., andHartfield, R. J., ADetailed Investigation of Staged Normal Injection into a Mach 2 Flow, 27th J ANNAF Combustion Meeting, F. E. Warren Air Force Base, Cheyenne, WY, Nov. 1990. Eklund, D. R., Northam, G. B., and Fletcher, D. G., A Validation Study of the SPARK Navier-Stokes Code for Nonreacting Scramjet Combustor Flowfields, AIAA Paper 90-2360, July 1990. Drummond, G. B., A Two-Dimensional Numerical Simulation of a Supersonic, Chemically Reacting Mixing Layer, NASA TM 4055, Dec. 1988. Hitch, B. D., and Senser, D. W., Reduced H2-O2 Mechanisms for Use in Reacting Flow Simulation, AIAA Paper 88-0732, Jan. 1988.

Ignition and Combustion Performance of Scramjet Combustors with Fuel Injection Struts

Ignition and combustion performance of a scramjet combustor with a fuel injection strut was experimentally investigated with Mach 2.5 vitiated air. Five strut models with different leading-edge geometry were tested without fuel injection to select the less flow-disturbing configuration. The nonreacting flowfields were also investigated by computation with a two-dimensional Navier—Stokes code. Using the selected strut, combustion and ignition tests were conducted. A pitot pressure and gas composition survey was carried out to deduce mixing and combustion efficiencies. It was found that mixing and combustion with a less flow-disturbing strut was considerably worse than those with a more flow-disturbing strut. Autoignition and forced ignition with plasma torches were tested for hydrogen. Ignition characteristics of parallel and perpendicular injection were quite different. The plasma igniters could successfully ignite both parallel and perpendicular fuel jets without a noticeable time delay between both sides of the strut.

Three-Dimensional Flows in a Low-Specific-Speed Diffuser Pump Stage. Study by 3 D Stage Flow Calaculations and Flow Visualization.

A viscous stage calculation method was used for investigating complex three-dimensional flow phenomena, particularly under off-design operating conditions, occurring in a low-specific-speed diffuser pump stage with a centrifugal impeller and a vaned bowl diffuser. The incompressible version of Dawes three-dimensional Navier-Stokes code was extended for the stage calculation by introducing an inter-row mixing process proposed by Denton. The numerical prediction succeeded in predicting major flow patterns obtained by a multicolor oil-film flow visualization technique which featured the use of tri-primary-color fluorescent powder as pigment. The meridional and blade-to-blade secondary flows, including the inpeller inlet recirculation at the shroud side, the diffuser inlet recirculation at the hub side, and the formation of the extensive vortex on the diffuser hub surface, were investigated numerically and experimentally in detail. The flow mechanism leading to the onset of impeller inlet recirculation was also discussed in detail.

On the bifurcation structure of axisymmetric vortex breakdown in a constricted pipe

The bifurcation structure is presented for an axisymmetric swirling flow in a constricted pipe, using the pipe geometry of Beran and Culick [J. Fluid Mech. 242, 491 (1992)]. The flow considered has been restricted to a two-dimensional parameter space comprising the Reynolds number Re and the relative swirl V0 of the incoming swirling flow. The bifurcation diagram is constructed by solving the time-dependent axisymmetric Navier–Stokes equations. The stability of the steady results presented by Beran and Culick, obtained from a steady axisymmetric Navier–Stokes code, has been confirmed. Further, the steady solution branch has also been extended to much larger V0 values. At larger V0, a stable unsteady solution branch has been identified. This unsteady branch coexists with the previously found stable steady solution branch and originates via a turning point bifurcation. The bifurcation diagram is of the type described by Benjamin [Proc. R. Soc. London Ser. A 359, 1 (1978)] as the canonical unfolding of a pitchfork bifurcation. This type of bifurcation structure in the two-dimensional parameter space (Re,V0), suggests the possibility of hysteresis behavior over some part of parameter space, and this is observed in the present study. The implications of this on the theoretical description of vortex breakdown and the search for a criterion for its onset are discussed.

High Reynolds number and turbulence effects on turbine heat transfer

Experimental data on pressure distribution and heat transfer on a turbine airfoil were obtained over a range of Reynolds numbers from 0.75 to 7.0 x 10(exp 6) and a range of turbulence intensities from 1.8 to about 15%. The purpose of this study was to obtain fundamental heat transfer and pressure distribution data over a wide range of high Reynolds numbers and to extend the heat transfer data base to include the range or Reynolds numbers encountered in the Space Shuttle main engine turbopump turbines. The results of this study indicated that Reynolds number and turbulence intensity have a large effect on both the transition from laminar to turbulent flow and the resulting heat transfer. For a given turbulence intensity, heat transfer for all Reynolds numbers at the leading edge can generally be correlated with the Frossling number developed for lower Reynolds numbers. For a given turbulence intensity, heat transfer for the airfoil surfaces downstream of the leading edge can be approximately correlated with a dimensionless parameter. Comparisons of the experimental results were also made with a numerical solution from a two-dimensional Navier-Stokes code.

Computation of the poststall behavior of a circulation controlled airfoil

The physics of the circulation controlled airfoil is complex and poorly understood, particularly with regards to jet stall, which is the eventual breakdown of lift augmentation by the jet at some sufficiently high blowing rate. The present paper describes the numerical simulation of stalled and unstalled flows over a two-dimensional circulation controlled airfoil using a fully implicit Navier-Stokes code, and the comparison with experimental results. Mach numbers of 0.3 and 0.5 and jet total to freestream pressure ratios of 1.4 and 1.8 are investigated. The Baldwin-Lomax and k-epsilon turbulence models are used, each modified to include the effect of strong streamline curvature. The numerical solutions of the post-stall circulation controlled airfoil show a highly regular unsteady periodic flowfield. This is the result of an alternation between adverse pressure gradient and shock induced separation of the boundary layer on the airfoil trailing edge.

Computational/experimental investigation of staged injection into a Mach 2 flow

The stage normal injection of two jets of air into a Mach 2 freestream behind a rearward-facing step has been investigated using laser-induced iodine fluorescence and laser doppler anemometry techniques. A detailed data set has been compiled that includes profiles of pressure, temperature, two components of velocity, and mole fraction of the injectant at 44 locations, plus planar surveys of pressure, temperature, velocity, and mole fraction. A companion numerical study was performed using the SPARK Navier-Stokes code. Good overall agreement was observed for this complex, highly three-dimensional flowfield. Discrepancies were observed in the computed and measured total temperatures, turbulent mixing, and shock strengths. The effect of grid resolution was investigated by calculating solutions on grids of 60,000 points, 200,000 points, and 450,000 points. Differences in the solutions on the two finer grids were small. The effect of turbulence modeling was investigated by calculating solutions with three different algebraic models for the jet turbulence. Overall, the turbulence models were found to have the greatest effect on the numerical solutions, followed by the grid resolution and the injectant Mach number.

Three-dimensional simulation of a translating strut inlet

A three-dimensional Navier-Stokes code is used to numerically simulate the flow through a translating strut scramjet inlet. The inlet has variable geometry for efficient operation over a wide speed range. Overall flow-field features such as the corner flow, topwall separation, shockwave coalescence, cowl pressure increase, and flow distortion at the throat are investigated. Comparisons are made with experimental results to provide for the assessment of the present analysis. Effects of boundary-layer ingestion on the overall flow features are also investigated.

Experimental and computational studies of flow in a simplified HVAC duct

The analysis of air flow in an automotive air–conditioning (A/C) duct is complicated by large cross–sectional area variations, abrupt changes in flow direction and flow maldistribution. In this study, the complex three–dimensional turbulent flow found in a generic A/C duct was investigated experimentally and computationally. Outlet flow split, exit velocity profile and surface pressure distribution along the duct were measured for use in validating computational flow capabilities. A three–dimensional Navier–Stokes code was utilized to predict the overall flow distribution and surface pressure. This work has validated the performance of the computer code for the overall flow prediction. The present study also identified the area of further improvements especially for the flow in regions of turbulent separation.

Numerical Simulation of the Flow Field Around Supersonic Air-Intakes

The demand for efficiency in today’s and in future civil aircraft is such that experimental studies alone do not suffice to optimize aircraft aerodynamics. In this context, much effort has been spent in the past decade to develop numerical methods capable of reproducing the phenomena that occur in the engine flow field. This paper presents some studies in Computational Fluid Dynamics related to supersonic inlets. Two approaches are considered. First, there is a need for a code capable of calculating in a cost-efficient way the entire flow field around a two-dimensional or three-dimensional inlet, e.g., to perform parametric studies. To this effect, a computing method based on grid construction by mesh generator dedicated to inlet shapes and on the discretization of the unsteady Euler equations with an explicit upwind scheme was developed. The treatment of complex geometries led us to adopt a multiblock grid approach. Therefore particular attention was paid to the treatment of the boundary conditions between the different domains. Second, there is a need for a code that can capture local phenomena in order to get a better understanding of inlet behavior (shock/shock, shock/boundary layer interactions, etc.). To this effect a two-dimensional turbulent Navier-Stokes code is used. The two-equation k-ε turbulence model included in the program seems to be one of the most successful models for calculating flow realistically. Correction of the near-wall influence extends its capability to complex flow configurations, e.g., those with separated zones.

Computational investigation of circular-to-rectangular transition ducts

A study demonstrating that flows inside transition ducts with unusual geometry can be analyzed by employing proper selections of a Navier-Stokes code, grid topology, and turbulence modeling is presented. Based on comparison between existing experimental data and the computed results for the same configurations, reasonable agreements were obtained for wall static pressure in the transition duct. Static pressure comparisons in the supersonic nozzle section were excellent, as well as agreement between computed and measured mass flow and thrust performance.