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This paper reports on the SIMPLE-like solution algorithm that significantly improves velocity-pressure coupling, leading to the accelerated convergence. The algorithm is applicable to both incompressible and compressible flows. It is implemented within the framework of a cell centered finite volume method using colocated storage of flow variables on unstructured grids. The convergence acceleration is an outcome of the novel treatment of the pressure correction equation that accounts for two commonly neglected terms appearing in a general pressure correction equation. These terms, namely, the neighbor and nonorthogonal corrections, represent velocity corrections from neighboring cells and the cell-face pressure correction gradient associated with the grid nonorthogonality, respectively. Reminiscent of the PISO method, both neighbor and nonorthogonal corrections are taken into account by performing two or more correction steps. However, the full inclusion of neighbor velocity corrections can prevent the solution convergence. Following an analogy between time marching and a steady-state iterative approach, this problem has been resolved by introducing a pseudounsteady term into discretized velocity correction equations. Using this term, the contribution of neighbor corrections is underrelaxed, enabling a satisfactory convergence rate. The algorithm is applied to several benchmark test cases, covering laminar and turbulent, as well as incompressible and compressible, flows. For all test cases, the convergence rate can be significantly improved by performing two or more pressure correction steps. An optimal number of pressure corrections exists for which a meaningful reduction of computing time is possible. [ABSTRACT FROM AUTHOR]

The modified NACA 643-618 airfoil of the Aeromot 200S Super Ximango motorglider was chosen for wing section simulations at Re=300,000 and α=15 deg. A coarse-grid direct numerical simulation was carried out to obtain detailed reference data. A laminar separation bubble close to the leading edge causes transition to turbulence of the suction-side boundary layer. The turbulent boundary layer separates near midchord. The considerable computational expense of this simulation provided the motivation for hybrid simulations based on the k-ω model and a one-equation renormalization group model. An extension to the turbulence models that "seeds" synthetic turbulence velocity fluctuations in the parts of the flowfield where the grid resolution changes is proposed but was found to have little effect on the mean flow. When the model contribution was too low, the turbulent boundary layer on the airfoil suction surface separated too early. By increasing the model contribution, a reasonable agreement with the reference data was obtained. [ABSTRACT FROM AUTHOR]

This paper extends to transonic mixing layers an analysis of Lighthill ("Reflection at a Laminar Boundary Layer of a Weak Steady Disturbance to a Supersonic Stream, Neglecting Viscosity and Heat Conduction," Quarterly Journal of Mechanics and Applied Mathematics, Vol. 54, No. 3, 1950, pp. 303-325.) on the interaction between weak shocks and laminar boundary layers. As in that work, the analysis is carried out under linear-inviscid assumptions for the perturbation field, with streamwise changes of the base flow neglected, as is appropriate given the slenderness of the mixing-layer flow. The steady-disturbance profile is determined by taking a Fourier transform along the longitudinal coordinate. Closed-form analytical functions for the pressure field are derived in the small- and large-wave-number limits, and vorticity disturbances are obtained as functions of the pressure perturbations. The analysis is particularized to ethylene-air and hydrogen-air mixing layers, for which the dynamics are of current interest for hypersonic propulsion. The results provide, in particular, the effective distance of upstream influence of the pressure perturbation in the subsonic stream. The resulting value, which scales with the thickness of the subsonic layer, is much smaller than the upstream influence distances encountered in boundary layers. This study may serve as a basis to understand shock-induced autoignition and flameholding phenomena in simplified versions of non-premixed supersonic-combustion problems. [ABSTRACT FROM AUTHOR]

Boundary-layer receptivity to acoustic disturbances plays a key role in transition from laminar to turbulent flow. Acoustic disturbances interact with strong streamwise gradients at the leading edge to create Tollmien-Schlichting waves in the boundary layer. Measurement of Tollmien-Schlichting receptivity to downstream-traveling sound is complicated by the presence of Stokes waves in the boundary layer and upstream-traveling acoustic reflections that also generate downstream-traveling Tollmien-Schlichting waves. Active noise control is used to cancel reflections and enables the measurement of boundary-layer receptivity to downstream-traveling sound. Tollmien-Schlichting wave amplitudes are extracted from hot-wire data using biorthogonal decomposition of modified Orr-Sommerfeld equations that include acoustic disturbance solutions. The new method is implemented to measure the acoustic receptivity of a 20:1 modified superellipse leading edge on a flat plate. The data yield acoustic receptivity coefficients consistent with the results of previous computational studies. For comparison, pulsed-sound measurements are also obtained, which are comparable to both biorthogonal decomposition and direct numerical simulation results. The new measurement technique provides for efficient decomposition of measurements and sets the stage for renewed experimental effort in this area. [ABSTRACT FROM AUTHOR]

Large-eddy simulations are carried out to investigate plasma-based flow control that is used to delay transition generated by excrescence on the leading edge of a wing. The wing airfoil section has a geometry that is representative of modern reconnaissance air vehicles and has an appreciable region of laminar flow at design conditions. Modification of the leading edge, which can be caused by the accumulation of debris, insect impacts, microscopic ice crystal formation, damage, or structural fatigue, may result in premature transition and an increase in drag. A dielectric barrier discharge plasma actuator, located downstream of the excrescence, is employed to mitigate transition, decrease drag, and increase energy efficiency. Numerical solutions are obtained to the Navier-Stokes equations that were augmented by source terms used to represent the body force imparted by the plasma actuator on the fluid. A simple phenomenological model provided this force resulting from the electric field generated by the plasma. The numerical method is based upon a high-fidelity numerical scheme and an implicit time-marching approach. An overset mesh system is employed to represent excrescence in the leading-edge region. Solutions are generated for both uniform and distributed excrescence geometries, as well as for the clean wing configuration without leading-edge modification. Results are obtained for two different values of the plasma field strength. Features of the computational flowfields are elucidated, and the effectiveness of control is quantified by comparison with baseline results without plasma actuation. It is found that plasma control can reestablish the laminar flow region lost to excrescence-generated transition and increase the lift-to-drag ratio by up to 8.7%. [ABSTRACT FROM AUTHOR]

Two-dimensional direct numerical simulations are used to study the impact of deep gaps on laminar-turbulent transition in compressible boundary-layer flow. For these, the gap depth-to-width ratio is always larger than five. They are located on a flat plate without pressure gradient. A steady base flow is used with a Mach number of 0.6, free-stream temperature of 288 K, and free-stream pressure of 1 bar. Subsequently, Tollmien-Schlichting waves are introduced by suction and blowing at the wall, and their growth over the gap is evaluated by N factors. The influence of the gap on laminar-turbulent transition is quantified by the difference ΔN compared with the N factor obtained for a flat plate without gap. A periodic influence of the gap depth on ΔN is observed. In the direct numerical simulations, acoustic waves enter the gap and form a standing wave due to reflections, similar as occurring in organ pipes. The feedback of the standing wave on the boundary-layer flow above is essential for the observed ΔN variations. In a second case, the influence of a specific gap placed in front of a forward-facing step is studied as well. Here, a reduction of the N factor and hence a delay of transition, relative to the flow with step alone, are reached due to the presence of the gap. [ABSTRACT FROM AUTHOR]

The friction component is responsible for more than 40 % of typical civil aircraft drag. As a consequence, the issue of laminar flow has been of prime importance in aeronautics for many years now. This article is focused on Tollmien-Schlichting-induced transition and drag predictions of two-dimensional laminar airfoils obtained with experimental and numerical methods. In 2012, a test campaign in the ONERA-S2MA wind tunnel, including infrared acquisitions, pressure sensors, and wake analyses, allowed substantial data to be obtained on such airfoils in transonic conditions. To complete this study, two-dimensional fluid dynamics computations have been performed, either with a Reynolds-averaged Navier-Stokes solver using transition criteria or with a boundary-layer code combined with direct stability analysis. Furthermore, experimental (wake survey) and numerical (far-field theory) techniques allowing airfoil drag breakdown have been employed. Wind-tunnel and computational fluid dynamics transition predictions have been compared. Good agreement has been observed but the transition criteria may show some limitations in particular situations, such as long separation bubble development. The gains in lift and drag due to laminar flow have been quantified (natural vs triggered transition). Concerning drag reduction, the importance of the viscous pressure component has been highlighted. Finally, the effects of parameters such as angle of attack, Mach number, and Reynolds number on transition location and drag have been investigated. [ABSTRACT FROM AUTHOR]

A substantial number of large-eddy simulations are conducted on separated flow controlled by a dielectric barrier discharge plasma actuator at a Reynolds number of 63,000. In the present paper, the separated flow over a NACA 0015 airfoil at an angle of attack of 12 deg, which is just poststall, is used as the base flow for separation control. The effects of the location and operating conditions of the plasma actuator on the separation control are investigated by a parametric study. The control effect is evaluated based on the improvement of not only the lift coefficient but also the drag coefficient over an airfoil. The most effective location of the plasma actuator for both lift and drag improvement is precisely confirmed to be upstream of the natural separation point. Even a low burst ratio is found to be sufficient to obtain the same improvements as the cases with a high burst ratio. The effective nondimensional burst frequency F+ is observed at 4 ≥ F+ ≥ 6 for the improvement in the lift coefficient and at 6 ≥ F+ ≥ 20 for that in the drag coefficient. The lift/drag ratio shows a clear peak at 6 ≥ F+ ≥ 10. To clarify the mechanism of the laminar-separation control, the effect of a turbulent transition is investigated. There is a clear relationship between the separation control effect and the turbulent transition at the shear layer. An earlier and smoother transition case shows greater improvements in the lift and drag coefficients. Flow analyses show that the cases with early and smooth turbulent transition can attach the separated flow further upstream, resulting in a higher suction peak of the pressure coefficient. In addition, another mechanism of the separation control is observed in which the lift coefficient is improved, not by the reattachment through the turbulent transition but by the large-scale vortex shedding induced by the actuation. It is possible to separate these two dominant mechanisms based on the effect of the turbulent transition on the separation control. [ABSTRACT FROM AUTHOR]

The influence of forward-facing steps on the amplification of disturbances in transonic flows is studied numerically. The main goal is to find a function for an additional amplification factor ΔN that can be incorporated into the eN method. In the present study, both two-dimensional direct numerical simulation of disturbances as well as linear stability theory are used to gain detailed information on the flow. Good agreement between A-factor results from direct numerical simulation and linear stability theory is obtained. Subsonic and slightly supersonic results show differences in the flow topology and in the shape of the N-factor curve, but effective ΔN factors are in the same order of magnitude. [ABSTRACT FROM AUTHOR]

Separation control by two-dimensional periodic excitation is studied theoretically for nonsymmetrical flows over a parabola using a mathematical model of incompressible laminar flow with periodic perturbations. This model is based on the representation of the flow as the sum of the mean and the periodic components. The governing equations for both components are derived from the Navier-Stokes equations using time averaging. The momentum equations for the mean component comprise additional source terms that result from averaging the quadratic periodic terms. These source terms allow the control of separation. A numerical procedure has been developed to solve these equations along with equations of periodic flow, and to study the influence of inflow conditions, namely, frequency, amplitude, and the station where the excitation is inserted on the mean flow. Using this procedure, one can check if the given inflow conditions suppress the leading-edge separation of the flow over a parabola. [ABSTRACT FROM AUTHOR]

Nonlinear parabolized stability equations and secondary-instability analyses are used to provide a computational assessment of the potential use of the discrete-roughness-element technology for extending swept-wing natural laminar flow at chord Reynolds numbers relevant to transport aircraft. Computations performed for the boundary layer on a natural-laminar-flow airfoil with a leading-edge sweep angle of 34.6 deg, freestream Mach number of 0.75, and chord Reynolds numbers of 17 x 106, 24 x 106, and 30 x 106 suggest that discrete roughness elements could delay laminar-turbulent transition by about 20% when transition is caused by stationary crossflow disturbances. Computations show that the introduction of small-wavelength stationary crossflow disturbances (i.e., discrete roughness element) also suppresses the growth of most amplified traveling crossflow disturbances. [ABSTRACT FROM AUTHOR]

Numerical simulations are carried out to explore flow control that delays transition generated by excrescence on a platelike geometry in subsonic flow. Both forward-facing and rearward-facing steps of small roughness heights are considered in the simulations. These are representative of joints and other surface imperfections on wing sections that disrupt laminar flow, thereby increasing skin friction and drag. Dielectric barrier discharge plasma-based flow control is employed to delay transition and increase the extent of the laminar flow region. Solutions are obtained to the Navier-Stokes equations, which were augmented by source terms used to characterize the body force imparted by a plasma actuator on the fluid. A simple phenomenological model provided these forces resulting from the electric field generated by the plasma. The numerical method is based upon a high-order numerical scheme and an implicit time-marching approach on overset mesh systems used to describe the steps. Very small-amplitude numerical forcing is applied to generate perturbations, which are amplified by the geometric disturbances and result in transition, similar to the physical situation. Both continuous and pulsed operations of actuators are investigated. Features of the flowfields are described, and comparisons are provided between the baseline and control cases. It is found that use of plasma actuators can maintain laminar flow for the entire length of the computational domain, resulting in a reduction of the integrated drag by up to 70%. [ABSTRACT FROM AUTHOR]

Tests were conducted at the Cryogenic Ludwieg-Tube Gottingen to investigate the influence on onset of boundary-layer transition of thermal gradients on the model surface, whereby the temperature-sensitive paint technique has been used to detect transition. The laminar airfoil under investigation was tested at high Reynolds and at both low and high subsonic Mach numbers. The influence on boundary-layer transition of the temperature difference between the flow and the model surface due to the temperature drop occurring during a test run at the Cryogenic Ludwieg-Tube Gottingen was examined. Various combinations of flow conditions and angles of attack were employed to investigate different stability situations. Linear-stability computations, based on the experimental data, were also carried out. The unfavorable impact of surface temperatures larger than the adiabatic-wall temperature was identified to be dependent on the considered boundary-layer-stability situation. For certain stability situations, it was shown that a nonadiabatic model surface can strongly influence the boundary-layer transition location. [ABSTRACT FROM AUTHOR]

Effusion cooling of a laminar flat-plate boundary-layer flow at Mach 2.7 is investigated using direct numerical simulations. Air is employed as hot and cooling gas and is blown in the wall-normal direction through discrete orifices into the boundary layer to generate a cooling film. In the present investigations, the inclusion of the flow in the blowing pipes enables a comparison with modeled blowing where the coolant mass flux and temperature are prescribed in the respective areas at the wall. Attention is turned to 1) the blowing through multiple rows of holes, under the assumption of a fixed plenum pressure for all pipes; 2) the temperature modeling, which plays an important role when the cooling gas and wall temperature differ; and 3) the flow tripping to turbulence by the blowing. It is demonstrated that simple modeling fails for narrow hole spacing, which lowers (together with the pipe/main-flow interaction) the critical blowing ratio for flow tripping. [ABSTRACT FROM AUTHOR]

At transitional Reynolds numbers (104 - 105), many smooth airfoils experience laminar flow separation and possible turbulent reattachment, where the occurrence of either state is strongly influenced by small changes in the surrounding environment. The Eppler 387 airfoil is one of many airfoils that can have multiple lift and drag states at a single wing incidence angle. Prestall hysteresis and abrupt switching between stable states occur due to sudden flow reattachment and the appearance of a separation bubble close to the leading edge. Here, we demonstrate control of the flow dynamics by localized acoustic excitation through small speakers embedded beneath the suction surface. The flow can be controlled not only through variations in acoustic power and frequency, but also through spatial variations in forcing location. Implications for control and stabilization of small aircraft are considered. [ABSTRACT FROM AUTHOR]

A new laminar-turbulent transition model for low-turbulence external aerodynamic applications is presented that incorporates linear stability theory in a manner compatible with modem computational fluid dynamics solvers. The model uses a new transport equation that describes the growth of the maximum Tollmien-Schlichting instability amplitude in the presence of a boundary layer. To avoid the need for integration paths and nonlocal operations, a locally defined nondimensional pressure-gradient parameter is used that serves as an estimator of the integral boundary-layer properties. The model has been implemented into the OVERFLOW 2.2f solver. Comparisons of predictions using the new model with high-quality wind-tunnel measurements of airfoil section characteristics confirm the predictive qualities of the model, as well as its improvement over the current state of the art in computational fluid dynamics transition modeling at approximately half the computational expense. [ABSTRACT FROM AUTHOR]