Thin-layer Navier-Stokes Code Research Articles

COMPUTATIONAL SIMULATION OF DYNAMIC STALL ON THE NLR 7301 AIRFOIL

The dynamic stall behavior of the supercritical NLR 7301 airfoil is analyzed with a 2-D thin-layer Navier–Stokes code. The code solves the compressible Reynolds-averaged Navier-Stokes equations with an upwind biased numerical scheme in combination with the Baldwin–Lomax or the Baldwin–Barth turbulence models. The effect of boundary layer transition is incorporated using the transition length model of Gostelow et al. The transition onset location is determined with Michel's formula or it can be specified as an input parameter. The two turbulence models yield significantly different steady-state lift coefficients at incidences greater than 8°. The use of the one-equation Baldwin–Barth model together with the Gostelow transition model is found to give substantially better agreement with the experimental data of McCroskeyet al. than the Baldwin–Lomax model. Also, the unsteady computations are strongly affected by the choice of the turbulence model. The Baldwin–Barth model predicts the lift hysteresis loops consistently better than the algebraic turbulence model. However, the one-equation model improves the prediction of the moment hysteresis loops only for one test case.

Shock oscillations on biconvex aerofoils

The mechanism of the origin of shock oscillations on an 18% thick biconvex aerofoil is investigated using a thin-layer Navier-Stokes code. The approach used in this study is to initiate shock oscillations on an aerofoil in transonic flow by moving the aerofoil from a regime of non-periodic flow to a regime where periodic flow does occur by withdrawal of a trailing-edge splitter plate. The results indicate that the shock induced separation plays a leading role in the origin of shock oscillations.

Transonic shock oscillations on NACA0012 aerofoil

The mechanism of the origin of shock oscillations on NACA0012 aerofoils is investigated using a moving grid thin layer Navier Stokes code. The method used to understand the mechanism is to initiate the shock oscillations on an aerofoil by moving the aerofoil from a regime of steady transonic flow into a regime of periodic flow by a change in airflow incidence. The results indicate that the shock induced bubble plays a leading role in the origin of shock oscillations and the trailing edge has an affect on its amplitude.

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.

Head-on parallel blade-vortex interaction

An experimental and computational study was carried out to investigate the parallel head-on blade-vortex interaction (BVI) and its noise generation mechanism. A shock tube, with an enlarged test section, was used to generate a compressible starting vortex which interacted with a target airfoil. The dual-pulsed holographic interferometry (DPHI) technique and airfoil surface pressure measurements were employed to obtain quantitative flow data during the BVI. A thin-layer Navier-Stokes code (BV12D), with a high-order upwind-biased scheme and a multizonal grid, was also used to simulate numerically the phenomena occurring in the head-on BVI. The detailed structure of a convecting vortex was studied through independent measurements of density and pressure distributions across the vortex center. Results indicate that, in a strong head-on BVI, the opposite pressure peaks are generated on both sides of the leading edge as the vortex approaches. Then, as soon as the vortex passes by the leading edge, the high-pressure peak suddenly moves toward the low-peak-reducing in magnitude as it moves--simultaneously giving rise to the initial sound wave. In both experiment and computation, it is shown that the viscous effect plays a significant role in head-on BVIs.

Heatshield erosion in a dusty Martian atmosphere

The study calculates the effects of dust particle impacts on the erosion of the forebody heat shield for a 26-m diameter aerobraking vehicle entering the Martian atmosphere at 8.6 km/s. An explicit thin-layer Navier-Stokes code is used to compute the dustless flowfield about the vehicle for the actual Martian atmospheric composition. The deceleration and melting of the dust particles within the forebody shock layer was computed for dust spherules having initial diameters from 1 to 19 microns. All the particles began vaporizing shortly after entering the shock layer, but most survived to hit the heat shield surface. The two different heat shield materials considered were Shuttle ceramic tiles with glassy surfaces and a moderate-density ablator. For a vehicle with a ballistic coefficient of 200 kg/sq m, a glassy heat shield surface experienced an average of 7 mm of surface erosion, or 20 times the thickness of the glassy layer on ceramic tiles.

Numerical dissipation in F3D thin-layer Navier-Stokes solution for flows with wall transpiration

Numerical boundary conditions for mass injection/suction at the wall are incorporated in the thin-layer Navier-Stokes code, F3D. The accuracy of the boundary conditions and the code is assessed by a detailed comparison of the predictions of velocity distributions and skin-friction coefficients with exact similarity solutions for laminar flow over a flat plate with variable blowing/suction, and measurements for turbulent flow past a flat plate with uniform blowing. In laminar flow, F3D predictions for friction coefficient compare well with exact similarity solution with and without suction, but the code produces large errors at moderate-to-large values of blowing. These errors are attributed to the numerical dissipation due to Steger-Warmin g flux-vector splitting in the upwinding (streamwise) direction. This numerical dissipation is shown to yield a slight Mach number dependence of skin-friction coefficient due to blowing in turbulent flow. Predicted surface pressures for turbulent flow past an airfoil with mass injection are in qualitative agreement with measurements for a flat plate.

Turbulence model effects on separated flow about a prolate spheroid

The three-dimensional separated flow about a prolate spheroid at high incidence is numerically investigated using the F3D thin-layer Navier-Stokes code. The effect of different turbulence models on the flowfield solution and the characteristics of the predicted flow are analyzed. The models used in this study are the Baldvvin-Lomax algebraic model, the Baldvvin-Lomax model as modified for crossflow separation by Degani and Schiff, and a modified version of the Johnson-King model with and without the Degani-Schiff crossflow modifications applied. The Johnson-King model is applied to assess the importance of modeling nonequilibrium effects in predicting flow about a slender body at high incidence. The computations are made for steady-state, fully turbulent flow. The results are compared with experimental pressure data and with computational results obtained by Panaras and Steger, using the identical code and grid, but with their own modifications to the Baldwin-Lomax model. In addition, the computed solutions are analyzed using surface flow patterns, helicity density contours, and turbulent eddy-viscosity profiles. The results of this analysis provide insight into the effects of the turbulence models on flow characteristics and demonstrate the effect of the models on the accurate prediction of highly separated and vortical flows about a slender body.

Computation of three-dimensional turbulent vortical flows on bodies at high incidence

As Navier-Stokes computations are carried out for high-angle-of-attack flows, the various numerical and physical factors affecting the accuracy of the flow predictions must be determined. In an effort to further this understanding, numerical predictions of the complex three-dimensional separated flow about bodies of revolution in supersonic and subsonic flows have been performed. Predictions have been made for the supersonic flow over a secant ogive cylinder at relatively low angles of attack using a thin-layer Navier-Stokes code, showing good agreement with experimental pressure measurements as well as boundary-layer velocity profiles. The predictions are also compared with parabolized Navier-Stokes computations from a previous work. Subsonic predictions at high angles of attack were made for a tangent ogive cylinder and were found to be in good agreement with available experimental data. The accuracy of the predictions are shown to depend on using grids that are sufficiently fine to resolve the details of both the viscous boundary layer and the off-surface separated flow structures and, for turbulent flow cases, the use of a turbulence model that 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.

Modeling of subsonic flow through a compact offset inlet diffuser

This paper presents a comparison of computational and experimental results that were obtained during an investigation of compressible flow through a compact, highly offset diffuser. Entrance values for the Mach number and the Reynolds number were 0.341 and 5.75(10)5, respectively. Strong curvature of the diffuser centerline produced extensive separation along one wall, which resulted in an exhaust flow with highly complex crossflow patterns and spatial nonuniformities. Our objective was to evaluate the use of an existing thin-layer Navier-Stokes code (ARC3D) to predict effects of diffuser shape and inlet flow properties on pressure recovery and exit flow quality. Comparisons are shown between computed and measured flow velocity components and wall pressure distributions. Agreement was satisfactory except in regions of separated flow. The Baldwin-Lomax algebraic turbulence model used in the code does not appear to represent adequately this complex separated region. We present preliminary results of computations made with a one-half-equation turbulence model, which accounts for some of the history effects in computing the turbulence length scale. The computed results show that the details of the axial velocity distribution in the separated flow region were quite different for the two turbulence models, but other features of the flow, such as the static pressure distributions, were only slightly affected.

Unsteady separation over maneuvering bodies

Abstract : Development of a new boundary layer code has reached the status to permit meaningful applications. Computations have been carried out for the initiation of separation in the symmetry plane of a prolate spheroid of slenderness ratio 1/4, impulsively started into forward motion at zero incidence and also at 50 degrees angle. This case serves as validation by comparing with previously published results of Xu and Wang (ref.1), and also demonstrates that our method gives information of flow near the rear stagnation point not available in the literature. More studies have been performed on the optimization of surface suction to delay or prevent the unsteady separation for an impulsively started circular cylinder. Here the methodology should be of interest. Work on an unsteady three-dimensional thin-layer Navier-Stokes code, however, is progressing slowly. Keywords: Unsteady separation, three dimensional moving body, Separation control.

Computations of supersonic flow over a missile afterbody containing an exhaust jet

Abstract : A thin-layer Navier-Stokes code, developed for projectile aerodynamics, has been used to compute the supersonic flow over a missile afterbody containing a centered exhaust jet. The thin-layer, compressible, Navier-Stokes equations are solved using a time dependent, implicit numerical algorithm. A unique flow field segmentation procedure is used which preserves the sharp base corner and facilitates the adaption of the grid to the free shear layer in the base region. Solutions have been obtained for an axisymmetric, boattailed afterbody where the free stream Mach number is 2.0 and the jet exit Mach number is 2.5. Computations were made at various jet static pressure to free stream static pressure ratios from 1 through 9. Qualitative features of the base region flow field seen experimentally are very well observed in the computed results. Quantitative comparisons of base pressure with experiment indicate good agreement at high pressure ratios and some disagreement at low pressure ratios.

Computational investigation of a pneumatic forebody flow control concept

The effectiveness of a tangential slot blowing concept for generating lateral control forces on an aircraft forebody is analyzed using computational fluid dynamics. The flow about a fighter forebody is computed using a multiple-zone, thin-layer Navier-Stokes code. Tangential slot blowing is modeled by the use of an actuator plane. The effects of slot location and slot length on the efficiency of the system are analyzed. Results of the study indicate that placement of the slot near the nose of the aircraft greatly enhances the efficiency of the system, while the length and circumferential location of the slot are of secondary importance. Efficiency is defined by the amount of side force or yawing moment obtained per unit blowing coefficient. The effect of sideslip on the system is also analyzed. The system is able to generate incremental changes in forces and moments in flows with sideslip angles up to 10 deg comparable to those obtained at zero sideslip. These results are used to determine a baseline configuration for an experimental study of the tangential slot blowing concept.