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Numerical calculations were carried out in order to investigate the delay of transition to turbulence on a wing section by means of local dynamic surface deformation. Physically, the deformation ma...

Direct numerical simulations were carried out in order to reproduce the generation and control of transition on a flat plate by means of local dynamic surface modification. The configuration and fl...

Direct numerical simulations (DNSs) were carried out in order reproduce the generation and control of transition on a flat plate by means of local dynamic surface modification. The configurations and flow conditions duplicate those of previous numerical investigations, and are similar to an experimental arrangement, which employed piezoelectrically driven actuators to impart small amplitude local deformation of the plate surface. In those studies, one actuator was located in the upstream plate region, and oscillated at the most unstable frequency of 250 Hz in order to generate small disturbances, which amplified Tollmien–Schlichting instabilities. A second actuator placed downstream, was then oscillated at the same frequency, but with appropriate amplitudes in order to mitigate disturbance growth and delay the evolution of transition. Prior simulations employed an empirical process to determine optimal values of the control parameters. In the current effort, this process is replaced with a closed-loop control law. Numerical solutions are obtained to the two-dimensional and three-dimensional compressible Navier–Stokes equations, utilizing a high-fidelity numerical scheme and an implicit time-marching approach. Local surface modification of the plate is enforced via grid deformation. Results of the simulations are presented, and features of the flowfields are described. Comparisons are made between results obtained with the two control methods, and effectiveness of the closed-loop approach is evaluated.

A numerical investigation was carried out to assess the effectiveness of plasma-based flow control for a swept-wing configuration. A single dielectric barrier discharge plasma actuator was used to delay transition generated by excrescence near the leading edge. Large-eddy simulations were performed for the configuration, which had a wing airfoil section that is representative of modern reconnaissance air vehicles and maintained an appreciable region of laminar flow at design conditions. High-fidelity numerical solutions to the Navier–Stokes equations were generated for wing sweep angles of 30 and 45 deg and compared with the previously obtained unswept case. Control was generally more difficult to attain for nonzero sweep angles and transition could not be delayed for the 45 deg case unless a novel segmented swept actuator configuration was employed. Features of the computed flowfields are elucidated and effects of plasma actuation are quantified. It was found, for a sweep angle of 30 deg, that the wing l...

Compressibility effects were numerically investigated for use of plasma-based flow control, which was applied to delay transition generated by excrescence on the leading edge of a wing. The wing airfoil section incorporates 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 can be caused by the accumulation of debris, insect impacts, microscopic ice crystal formation, damage, or structural fatigue, resulting in premature transition and an increase in drag. A dielectric barrier discharge DBD plasma actuator, located downstream of the excrescence, was employed to delay transition, mitigate the effects of turbulence, decrease drag, and increase energy efficiency. Solutions were obtained for several Mach numbers, up to the transonic range. The effect of compressibility on transitional behaviour was explored, and the effectiveness of plasma-based control to delay transition with increasing Mach number was determined.

Numerical calculations were carried out to investigate control of transition on a flat plate by means of local dynamic surface deformation. The configuration and flow conditions are similar to a previous computation which simulated transition mitigation. Physically, the surface modification may be produced by piezoelectrically driven actuators located below a compliant aerodynamic surface, which have been employed experimentally. One actuator is located in the upstream plate region and oscillated at the most unstable frequency of 250 Hz to develop disturbances representing Tollmien–Schlichting instabilities. A controlling actuator is placed downstream and oscillated at the same frequency, but with an appropriate phase shift and modified amplitude to decrease disturbance growth and delay transition. While the downstream controlling actuator is two-dimensional (spanwise invariant), several forms of upstream disturbances were considered. These included disturbances which were strictly two-dimensional, those which were modulated in amplitude and those which had a spanwise variation of the temporal phase shift. Direct numerical simulations were obtained by solution of the three-dimensional compressible Navier–Stokes equations, utilizing a high-fidelity computational scheme and an implicit time-marching approach. A previously devised empirical process was applied for determining the optimal parameters of the controlling actuator. Results of the simulations are described, features of the flowfields elucidated, and comparisons made between solutions of the uncontrolled and controlled cases for the respective incoming disturbances. It is found that the disturbance growth is mitigated and the transition is delayed for all forms of the upstream perturbations, substantially reducing the skin friction.

Large-eddy simulations were carried out to describe the flow past a high-lift wing section. The configuration consists of a baseline laminar flow geometry, with smoothly deflected leading- and trailing-edge flaps, corresponding to a wind-tunnel model. Both flaps are deployed at a 45 deg angle with respect to the undeflected state, and blowing from internal plenums is employed to mitigate transition, increase attached flow, and enhance lift. Solutions were obtained to the Navier–Stokes equations, at the experimental chord-based Reynolds number of 1×106 and Mach number of 0.14. The numerical method is based upon a high-fidelity scheme and an implicit time-marching approach. Results were generated for two different angles of attack, in freestream conditions and within the confines of wind-tunnel walls. Comparisons are made with available experimental data in terms of surface pressure distributions, and the effect of blowing is quantified by comparison to baseline cases without control. Details of the computa...

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 numer...

Summary In this work, a study of the mechanism by which free-stream acoustic and vorticity disturbances interact with a boundary layer flow developing over a flat plate featuring a step excrescence located at a certain distance from a blunt leading edge is included. The numerical tool is a high-fidelity implicit numerical algorithm solving for the unsteady, compressible form of the Navier–Stokes equations in a body-fitted curvilinear coordinates and employing high-accurate compact differencing schemes with Pade-type filters. Acoustic and vorticity waves are generated using a source term in the momentum and energy equations, as opposed to using inflow boundary conditions, to avoid spurious waves that may propagate from boundaries. The results show that the receptivity to surface step excrescences is largely the result of an overall adverse pressure gradient posed by the step, and that the free-stream disturbances accelerate the generation of instabilities in the downstream. As expected, it is found that the acoustic disturbance interacting with the surface imperfection is more efficient in exciting the Tollmien–Schlichting waves than the vorticity disturbance. The latter generates Tollmien–Schlichting waves that are grouped in wave packets consistent with the wavelength of the free-stream disturbance. Copyright © 2015 John Wiley & Sons, Ltd.

The Navier–Stokes equations were solved using a high-fidelity time-implicit numerical scheme and an implicit large-eddy simulation approach to investigate plasma-based flow control for supersonic flow over a compression ramp. The configuration included a flat-plate region to develop an equilibrium turbulent boundary layer at Mach 2.25, which was validated against a set of experimental measurements. The fully turbulent boundary-layer flow traveled over a 24 deg ramp and produced an unsteady shock-induced separation. A control strategy to suppress the separation through a magnetically-driven surface-discharge actuator was explored. The size, strength, and placement of the model actuator were based on recent experiments at the Princeton University Applied Physics Group. Three control scenarios were examined: steady control, pulsing with a 50% duty cycle, and a case with significant Joule heating. The control mechanism was very effective at reducing the time-mean separation length for all three cases. The ste...

Large-eddy simulations are carried out to investigate the use of plasma-based actuation for the control of flows over a finite span wing at low Reynolds numbers. The wing section corresponds to the SD7003 airfoil, which is representative of those employed for micro air vehicle applications. Dielectric-barrier-discharge plasma actuators are used to modify the transitional flow and improve aerodynamic performance. Solutions are obtained to the Navier-Stokes equations, which were augmented by source terms used to represent plasma-induced body forces imparted by the actuators on the fluid. Simple phenomenological models provided the body forces generated by the electric field of the plasma surrounding the actuators. The numerical method is based upon a high-fidelity time-implicit scheme, an implicit large-eddy-simulation approach, and domain decomposition in order to perform calculations on a parallel computing platform. Flow at a chord-based Reynolds number of 40,000 is considered in the investigation, which is characterized by laminar separation on the suction surface of the wing at low angles of attack. This separation then promotes transition to a more complex state, which can be modified by the use of plasma actuation. Several aspects of control are examined, including different actuator configurations, alternative plasma-force models, both continuous and pulsed modes of operation, and the magnitude of plasma force required for control.

A large-eddy simulation (LES) was carried out in order to numerically describe the use of plasma-based actuation for the control of turbulent boundary-layer separation. The configuration consisted of a flat plate section over which the boundary layer developed, followed by a curved convex rearward-facing ramp, corresponding to an experimental arrangement. A single dielectric-barrier-discharge (DBD) plasma actuator was then employed to reduce the extent of the separated flow region. Solutions were obtained to the Navier-Stokes equations, that were augmented by source terms used to represent plasma-induced body forces imparted by the actuator on the fluid. A simple phenomenological model provided the electric field generated by the plasma, resulting in the body forces. The numerical method utilized a high-fidelity time-implicit scheme, employing domain decomposition in order to perform calculations on a parallel computing platform. Simulations were first conducted for an isolated actuator in the absence of external flow, which were used to optimize the choice of parameters inherent in the plasma model. Subsequently, actuation was applied to the plate development section without the downstream ramp section. Following these computations, the complete plate/ramp configuration was simulated. In all cases, both continuous and pulsed operation of the actuator was considered. Comparisons are made with available experimental data, and with baseline flows where no control was enforced. The large separated flow region characterizing the baseline flow for the plate/ramp configuration was well captured by the simulation. When continuous control was applied, separation was almost entirely eliminated. Pulsing operation of the actuator using a 40% duty cycle was not as eective, but did result in a considerable reduction of the recirculating zone. The control cases compared reasonably well with experimental measurements. Although deficiencies in the plasma model were apparent in the near-wall region, it appeared to be adequate for use with LES in the exploration of plasma-based control for turbulent wall-bounded flows.

Direct numerical simulation is employed in order to describe the subsonic flow past an array of micron-sized discrete roughness elements, which were mounted near the leading edge of a 30-degree swept wing at a chord Reynolds number of 7.4 x 10 6 . The flow conditions correspond to flight receptivity experiments that were conducted to investigate the effects of roughness on crossflow instabilities. To make the computations tractable, the geometry is scaled by the radius of the wing leading edge, which magnifies the region of interest and enhances resolution. The leading-edge region is then approximated by the flow past an infinite parabolic cylinder. The numerical method is based upon a sixth-order-accurate time-implicit scheme to attain high fidelity and was used in conjunction with an eighth-order low-pass Pade-type nondispersive filter operator to maintain stability. A high-order overset-grid approach preserved spatial accuracy on a local mesh system representing the roughness elements, using domain decomposition to perform calculations on a parallel computing platform. The direct simulation for the flow about the roughness elements was used to capture crossflow vortices and served as input to the nonlinear parabolized stability equations, which were then solved in order to determine receptivity of the flow to the geometric perturbations. Three different geometric roughness elemental shapes were investigated in the study. For one shape, the effect of element height was examined. Features of the roughness-element flowfields are elucidated, and findings of the stability calculations are compared. Results are presented for receptivity of the crossflow instability to the size and shape of elements, as obtained by the direct numerical simulation and by two different stability approaches.

Usage of large-eddy simulation (LES) methods for calculation of turbulent flows has increased substantially in recent years. This paper attempts to 1) provide an assessment of the current capabilities of LES, 2) outline some recommended practices for using LES, and 3) identify future research needs. The assessment considers flow problems for which LES can be successfully applied today and flow problems for which LES still has limitations. The availability of LES and hybrid Reynolds-averaged Navier-Stokes (RANS)/LES in general-purpose codes is discussed. Several important issues for which the LES community has not yet reached a consensus are discussed. These include grid sensitivity studies, application of unstructured grid methods, upwind-biased solvers, and turbulence (subgrid) modeling including continuous hybrid RANS/LES approaches. A section on recommended practices and key considerations tries to provide guidance on some of the important items that need to be addressed in using LES. The paper concludes with a discussion of future research directions, with a focus on work needed to advance the capabilities and reliability of LES for analysis of turbulent flows.

The purpose of this article is to summarize a computational approach, which developed and matured over an extended period of time, and has been shown to be useful for performing large-eddy simulation (LES) of flows with active control. Because of the nature of active flow control, simulation of this class of problems typically cannot be carried out accurately by methods less sophisticated than LES. Active control flowfields are highly unsteady, and can be characterized by small-scale fluid structures which are produced by the control process, but may also be inherent in the original uncontrolled situation. The numerical scheme is predicated upon an implicit time-marching algorithm, and utilizes a high-order compact finite-difference approximation to represent spatial derivatives. Robustness of the scheme is maintained by employing a low-pass Pade-type nondispersive spatial filter, which also accounts for the fine-scale turbulent dissipation that otherwise is traditionally provided by an explicitly added subgrid-scale (SGS) stress model. Geometrically complex applications are accommodated by an overset grid technique, where spatial accuracy is preserved through use of high-order interpolation. Utility of the method is illustrated by specific computational examples, including suppression of acoustic resonance in supersonic cavity flow, leading-edge vortex control of a delta wing, efficiency enhancement of a transitional highly loaded low-pressure turbine blade, and separation control of a wall-mounted hump model. Control techniques represented in these examples are comprised of both steady and pulsed mass injection or removal, as well as plasma-based actuation. For each case, features of the flowfield are elucidated and the solutions are compared to the baseline situation where no control was enforced. Where available, comparisons are also made with experimental data.

Large-eddy simulations were carried out in order to explore flow control strategies for a high-lift wing section. The configuration consisted of a baseline laminar flow geometry, with smoothly deflected leading and trailing-edge flaps, that were are deployed at a 45 deg angle with respect to the undeflected state. In previous investigations of the NASA based design, continuous blowing from internal plenums was utilized to mitigate transition, increase attached flow, and enhance lift. While that approach was successful, it required a substantial amount of mass flow to improve performance. The current effort examines alternative control systems that may be more energy efficient. Solutions were obtained to the Navier-Stokes equations, at the experimental chord-based Reynolds number of 1.0×10 and Mach number of 0.14. The numerical method is based upon a high-fidelity scheme and an implicit time-marching approach. An overset mesh system was employed to represent blowing nozzles interior to the wing surface. Simulations considered two different angles of attack, and control consisting of pulsed blowing and the segmented arrangement of nozzles. Details of the computations are described, effectiveness of the control is quantified, and physical features of the computed flowfields are characterized. Comparisons are provided with baseline cases without control, and with continuous blowing situations, where experimental data is also available. It was found that pulsed blowing can reduce mass flow requirements by more then 30% with minimal loss of lift. For segmented blowing, mass flow was reduced by over 50%, however control was found to be largely ineffective.

This work describes an implicit large-eddy simulation (LES) for active control of flow over a wall-mounted hump. Both steady suction and oscillatory blowing and suction are compared to baseline conditions with no flow control. This simulation models an experiment conducted by NASA which was one of the test cases in their 2004 CFD Validation on Synthetic Jets and Turbulent Separation Control Workshop. A previous LES of the baseline case with the current scheme demonstrated significantly better agreement with the experimental flow physics than RANS in the separated region downstream of the hump. The current work concludes that this also holds for the cases using active control. The LES is accomplished using an implicit parallel flow solver that is based on an approximately-factored time-integration method using fourth-order spatial compact-dierence formulations and a high-order filtering strategy. To properly resolve the flow for this LES, the Reynolds number of 9.36 ◊ 10 5 used in the experiment was reduced to 2.0 ◊ 10 5 . Eects of lowering the Reynolds number were previously investigated using RANS. Flow features of the active control cases are compared with the baseline case, each other, and experimental data. The active control LES cases matches experimental data significantly better in the recirculation region than RANS. The baseline and steady suction LES separation reattachment lengths are within 2% and 4%, respectively, of the experimental locations. These simulations achieve very good agreement with the experimental surface pressure coecient, skin friction coefficient, mean velocity profiles, Reynolds stresses, and flow reattachment location. Because of the lower Reynolds number, the oscillatory blowing and suction flow control only has about 25% of the eectiveness of that observed in the experiment. The LES separation reattachment length is 10% longer than its experimental counterpart. Other flow quantities display favorable agreement with experimental data. Results demonstrate that both steady suction and oscillatory blowing and suction can be properly simulated by LES and eectively reduce the size of the separated flow region.

Plasma-based active flow control was simulated numerically for the subsonic flow through a highly loaded low-pressure turbine. The configuration corresponded to previous experiments and computations which considered flow at a Reynolds number of 25,000 based upon axial chord and inlet conditions. In this situation, massive separation occurs on the suction surface of each blade due to uncovered turning. The present exploratory numerical study was performed to investigate the use of asymmetric dielectric-barrier-discharge actuators for mitigating separation, thereby decreasing turbine wake losses and increasing efficiency. Solutions were obtained for the Navier-Stokes equations, which were augmented by a phenomenological model that was used to represent plasma-induced body forces imparted by the actuator on the fluid. The numerical method used a high-fidelity time-implicit scheme, employing domain decomposition to carry out calculations on a parallel computing platform. A high-order overset grid approach preserved spatial accuracy in a locally refined embedded region. The magnitude of the plasma-induced body force required for control is examined, and both continuous and pulse-modulated actuations are considered. Novel use of counterflow actuation is also investigated, and the effects of pulsing frequency and duty cycle are considered. Features of the flowfields are described, and resultant solutions are compared with each other, with previous mass-injection control cases, and with the baseline situation where no control was enforced.

Direct numerical simulation was used to describe the subsonic flow past an array of distributed cylindrical roughness elements mounted on a flat plate. Solutions were obtained for element heights corresponding to a roughness-based Reynolds number (Re k ) of both 202 and 334. The numerical method used a sixth-order-accurate centered compact finite difference scheme to represent spatial derivatives, which was used in conjunction with a tenth-order low-pass Pade-type nondispersive filter operator to maintain stability. An implicit approximately factored time-marching algorithm was employed, and Newton-like subiterations were applied to achieve second-order temporal accuracy. Calculations were carried out on a massively parallel computing platform, using domain decomposition to distribute subzones on individual processors. A high-order overset grid approach preserved spatial accuracy on the mesh system used to represent the roughness elements. Features of the flowfields are described, and results of the computations are compared with experimentally measured velocity components of the time-mean flowfield, which are available only for Re k = 202. Flow about the elements is characterized by a system of two weak corotating horseshoe vortices. For Re k = 334, an unstable shear layer emanating from the top of the cylindrical element generated nonlinear unsteady disturbances of sufficient amplitude to produce explosive bypass transition downstream of the array. The Re k = 202 case displayed exponential growth of turbulence energy in the streamwise direction, which may eventually result in transition.

This work describes the development and validation of a parallel high-order compact finite difference Navier---Stokes solver for application to large-eddy simulation (LES) and direct numerical simulation. The implicit solver can employ up to sixth-order spatial formulations and tenth-order filtering. The parallelization of the solver is founded on the overset grid technique. LES were then performed for turbulent channel flow with Reynolds numbers ranging from Re ?=180 to 590, and flow past a circular cylinder with a transitional wake at Re D =3900. The channel flow solutions were obtained using both an implicit LES (ILES) approach and a dynamic sub-grid scale model. The ILES method obtained virtually identical solutions at half the computational cost. The original vector and new parallel solvers produce indistinguishable mean flow solutions for the circular cylinder. Repeating the cylinder simulation on a much finer mesh resulted in significantly better agreement with experimental data in the near wake than the coarse grid solution and other previous numerical studies.

Active control was simulated numerically for the subsonic flow through a highly loaded low-pressure turbine. The configuration approximated cascade experiments that were conducted to investigate a reduction in turbine stage blade count, which can decrease both weight and mechanical complexity. At a nominal Reynolds number of 25,000 based upon axial chord and inlet conditions, massive separation occurred on the suction surface of each blade due to uncovered turning. Vortex generating jets were then used to help mitigate separation, thereby reducing wake losses. Computations were performed using both steady blowing and pulsed mass injection to study the effects of active flow control on the transitional flow occurring in the aft-blade and near-wake regions. The numerical method utilized a centered compact finite-difference scheme to represent spatial derivatives, that was used in conjunction with a low-pass Pade-type nondispersive filter operator to maintain stability. An implicit approximately factored time-marching algorithm was employed, and Newton-like subiterations were applied to achieve second-order temporal accuracy. Calculations were carried out on a massively parallel computing platform, using domain decomposition to distribute subzones on individual processors. A high-order overset grid approach preserved spatial accuracy in locally refined embedded regions. Features of the flowfields are described, and simulations are compared with each other, with available experimental data, and with a previously obtained baseline case for the noncontrolled flow. It was found that active flow control was able to maintain attached flow over an additional distance of 19–21% of the blade chord, relative to the baseline case, which resulted in a reduction of the wake total pressure loss coefficient of 53–56%.

The development of a high-order spatial discretization for a k-e turbulence model and its application to flow over a wall-mounted hump is described. The high-order implementation is validated for a flat plate and subsequently applied to the more complex wall-mounted hump for conditions with and without flow control. Results for the hump flow are compared to experimental data. The turbulence model is incorporated in an implicit parallel flow solver that is based on an approximately factored time-integration method coupled with spatially fourth- and sixth-order compact-difference formulations and a high-order filtering strategy. Both second-order and high-order discretizations of the k-e turbulence equations were included in the compact solver. Validation using flow over a flat plate demonstrated that use of a second-order scheme for the k-∈ turbulence equations dominates the solution even when high-order compact differencing is used for the flow equations. This validation also demonstrated that significant computational savings are possible because less mesh resolution is required when using a high-order discretization of the k-e turbulence equations. Comparison of the high-order and second-order solutions was also performed for the wall-mounted hump. Qualitative agreement was achieved with experimental data for both high-and low-order schemes. High-order solutions on a coarse grid agreed very well with second-order solutions on a considerably finer grid.

The subsonic flow through highly loaded low-pressure turbines is simulated numerically using a high-order method. The configuration approximates cascade experiments that were conducted to investigate a reduction in turbine stage blade count, which can decrease both weight and mechanical complexity. At a nominal Reynolds number of 25 × 10 3 based upon axial chord and inlet conditions, massive separation occurs on the suction surface of each blade as a result of uncovered turning. Pulsed injection vortex generator jets were then used to help mitigate separation, thereby reducing wake losses. Computations were performed for both uncontrolled and controlled cases and reproduced the transitional flow occurring in the aft-blade and near-wake regions. The numerical method utilizes a centered compact finite difference scheme to represent spatial derivatives, which is used in conjunction with a low-pass Pade-type nondispersive filter operator to maintain stability. An implicit approximately factored time-marching algorithm is employed, and Newton-like subiterations are applied to achieve second-order temporal accuracy. Calculations were carried out on a massively parallel computing platform, using domain decomposition to distribute subzones on individual processors. A high-order overset grid approach preserved spatial accuracy in locally refined embedded regions. Features of the flowfields are elucidated, and simulations are compared with each other and with available experimental data. Relative to the uncontrolled case, it was found that pulsed injection maintained attached flow over an additional 15% of the blade chord, resulting in a 22% decrease of the wake total pressure loss coefficient.

A high-order numerical method is described for performing large-eddy simulations (LESs) of supersonic flowfields. Spatial derivatives are represented by a compact stencil that is used in conjunction with a tenth-order non-dispersive filter. The scheme employs a time-implicit approximately-factored finite-difference algorithm, and applies Newton-like subiterations to achieve second-order temporal and sixth-order spatial accuracy. Details of the method are summarized and LESs are carried out for a spatially evolving supersonic turbulent boundary layer. Two different subgrid-scale models are incorporated in the simulations to account for the spatially under-resolved stresses and heat flux. Comparisons are made between the respective computations, as well as with available experimental data and with previous numerical results.

Large-eddy simulations of supersonic cavity flowfields are performed using a high-order numerical method. Spatial derivatives are represented by a fourth-order compact approximation that is used in conjunction with a sixth-order nondispersive filter. The scheme employs a time-implicit approximately factored finite difference algorithm, and applies Newton-like subiterations to achieve second-order temporal and fourth-order spatial accuracy. The Smagorinsky dynamic subgrid-scale model is incorporated in the simulations to account for the spatially underresolved stresses. Computations at a freestream Mach number of 1.19 are carried out for a rectangular cavity having a length-to-depth ratio of 5:1. The computational domain is described by 2.06×10 7 grid points and has been partitioned into 254 zones, which were distributed on individual processors of a massively parallel computing platform. Active flow control is applied through pulsed mass injection at a very high frequency, thereby suppressing resonant acoustic oscillatory modes

This work investigates a high-order numerical method which is suitable for performing large-eddy simulations, particularly those containing wall-bounded regions which are considered on stretched curvilinear meshes. Spatial derivatives are represented by a sixth-order compact approximation that is used in conjunction with a tenth-order non-dispersive filter. The scheme employs a time-implicit approximately factored finite-difference algorithm, and applies Newton-like subiterations to achieve second-order temporal and sixth-order spatial accuracy. Both the Smagorinsky and dynamic subgrid-scale stress models are incorporated in the computations, and are used for comparison along with simulations where no model is employed. Details of the method are summarized, and a series of classic validating computations are performed. These include the decay of compressible isotropic turbulence, turbulent channel flow, and the subsonic flow past a circular cylinder. For each of these cases, it was found that the method was robust and provided an accurate means of describing the flowfield, based upon comparisons with previous existing numerical results and experimental data. Published in 2003 by John Wiley & Sons, Ltd.

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 pla...

This work investigates the application of a high-order finite difference method for compressible large-eddy simulations on stretched, curvilinear and dynamic meshes. The solver utilizes 4th and 6th-order compact-differencing schemes for the spatial discretization, coupled with both explicit and implicit time-marching methods. Up to 10th order, Pade-type low-pass spatial filter operators are also incorporated to eliminate the spurious high-frequency modes which inevitably arise due to the lack of inherent dissipation in the spatial scheme. The solution procedure is evaluated for the case of decaying compressible isotropic turbulence and turbulent channel flow. The compact/filtering approach is found to be superior to standard second and fourth-order centered, as well as third-order upwind-biased approximations. For the case of isotropic turbulence, better results are obtained with the compact/filtering method (without an added subgrid-scale model) than with the constant-coefficient and dynamic Smagorinsky models. This is attributed to the fact that the SGS models, unlike the optimized low-pass filter, exert dissipation over a wide range of wave numbers including on some of the resolved scales. For channel flow simulations on coarse meshes, the compact/filtering and dynamic models provide similar results, with no clear advantage achieved by incorporating the SGS model. However, additional computations at higher Reynolds numbers must be considered in order to further evaluate this issue. The accuracy and efficiency of the implicit time-marching method relative to the explicit approach are also evaluated. It is shown that a second-order iterative implicit scheme represents an effective choice for large-eddy simulation of compressible wall-bounded flows.

Large-eddy simulations of supersonic compression-ramp flowfields were performed by a high-order numerical method, utilizing the Smagorinsky dynamic subgrid-scale model to account for spatially underresolved stresses. Computations were carried out at a freestream Mach number of 3.0 for ramp angles of 8, 16, 20, and 24 deg. Extensive comparisons are made between the respective solutions and available experimental data that were collected at higher Reynolds numbers. These include surface pressure, skin friction, and both mean and fluctuating velocity profiles. For the 24-deg case, a number of experimentally measured statistical quantities are compared to the simulation

Direct numerical simulations are carried out to explore the use of flow control that delays transition generated by excrescence on a plate-like geometry in subsonic flow. Both forward-facing and rearward-facing steps of small roughness heights are considered in the investigation. These are representative of joints and other surface imperfections on wing sections that disrupt laminar flow, thereby increasing skin friction and configuration drag. Unlike previous studies, the steps have a finite lateral extent, such that sharp edges occur in both the spanwise and streamwise directions, and provide a more realistic characterisation of misaligned panels in aerodynamic configurations. The effect of spanwise corners upon transition is examined, and dielectric barrier discharge plasma-based flow control is applied to delay transition and increase the extent of the laminar flow region. Solutions are obtained to the Navier– Stokes equations that were augmented by source terms used to represent body forces imparted by plasma actuators 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-fidelity scheme and an implicit time-marching approach, on an overset mesh system that is used to represent the finite-span steps. Very small-amplitude numerical forcing is employed 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 considered, and the effectiveness of the control is quantified. Transition with the forward-facing step is considerably exacerbated by the presence of a spanwise edge. Plasma control is minimally effective, even with the use of multiple actuators and increased applied force. For the rearward-facing step, transition is substantially delayed by plasma control with small force application.

Numerous studies conducted over the years have shown that the transition onset in boundary layer flows is strongly dependent on the receptivity to various environmental disturbances. The objective of this paper is to study the mechanism by which free-stream acoustic and vorticity disturbances interact with a boundary layer flow developing over a flat-plate featuring a small excrescence located at a certain distance from a blunt leading edge. The numerical tool is a high-fidelity implicit numerical algorithm solving for the unsteady, compressible form of the Navier-Stokes equations in a body-fitted curvilinear coordinates, and employing high accurate compact differencing schemes with Pade type filters. Acoustic and vorticity waves are generated using a source term in the momentum and energy equations, as opposed to using inflow boundary conditions, to avoid spurious waves that may propagate from boundaries. The results show that the receptivity to surface excrescences is largely the result of an overall adverse pressure gradient posed by the step, and that the free-stream disturbances accelerate the generation of instabilities in the downstream. As expected, it is found that the acoustic disturbance interacting with the surface imperfection is more efficient in exciting the Tollmien-Schlichting waves than the vorticity disturbance. The latter generates TS waves that are grouped in wave packets with the length consistent with the wavelength of the free-stream disturbance.

A high-order method is used to perform large-eddy simulations of a supersonic compression-ramp flowfield. The procedure employs an implicit approximately factored finite difference algorithm, which is used in conjunction with a 10th-order nondispersive filter. Spatial derivatives are approximated by a sixth-order compact scheme, and Newton-like subiterations are applied to achieve second-order temporal accuracy. In the region of strong shock waves, the compact differencing of convective fluxes is replaced locally by an upwind-biased scheme. Both the Smagorinsky and dynamic subgrid-scale stress models are incorporated in the simulations. Details of the method are summarized, and a number of computations are carried out. Comparisons are made between the respective solutions as well as with available experimental data and with previous numerical results

The flowflelds surrounding a synthetic-jet actuating device are investigated numerically by direct simulation. Solutions are obtained to the unsteady compressible Navier-Stokes equations for both the interior of the actuator cavity and for the external jet flowfield. The interior results are generated on an overset deforming zonal mesh system, whereas the jet flowfield is obtained by a high-order compact-difference scheme. Newton-like subiterations are employed to achieve second-order temporal accuracy. Details of the computations are summarized, and the quality of the results is assessed via grid resolution and time-step size studies. Several aspects of the actuator configuration are investigated, including cavity geometry and Reynolds number. Differences between two-dimensional and three-dimensional external unsteady flowfields are elucidated, and comparison is made with experimental data in terms of the mean and fluctuating components of the jet velocity

High-Reynolds-number supersonic e owe elds were generated numerically to assess the performance of three explicit algebraic Reynolds-stress turbulence models. The cone gurations consist of a shock/boundary-layer interaction and a 24-deg compression ramp, both of which exhibit an appreciable region of separated e ow. Solutions were also obtained using standard zero-equation and k‐ models. Details of the computations are summarized, and the accuracy of numerical results is established via grid resolution studies. Comparisons are made with experimental data in terms of surface pressure and skin friction, as well as off-surface proe les of mean velocity and componentsoftheReynolds-stresstensor. Forthee owsconsidered here, it isfound that thealgebraic-stressmodels offer little improvement over existing closures.

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 numer...

Two-dimensional numerical simulations were carried out in order to investigate the effect of geometric disturbances on boundary-layer stability. The configurations consisted of abrupt steps with sharp edges and a smooth bump of small heights, on an otherwise smooth aerodynamic surface, which are representative of excrescence on laminar flow wings. Well resolved solutions to the Navier-Stokes equations were obtained by a high-order numerical method, for which it was found that the flows were inherently unsteady. Small-magnitude forcing was also applied to generate perturbations that were amplified by the change in surface shape. The applied forcing was representative of that commonly imposed for reproducing transition through the use of direct numerical simulation (DNS) or large-eddy simulation (LES). For each geometric disturbance, several roughness-based Reynolds numbers were considered, and the growth of perturbations was quantified. Linear stability analysis was also employed to analyze and characterize this growth. A parametric study was conducted to assess influence of the forcing frequency on evolution of fluid structures due to the disturbances. A combined DNS/LES computation was carried out for all cases in order to establish physically correct transition locations. Predictions based upon the two-dimensional results and linear stability analyses were quantified by comparison to the DNS/LES solutions using N-factor growth rates. Overall, it appeared that utilizing growth rates from forced solutions resulted in the most accurate prediction of transition locations.

Steady high-Reynolds-number flowfields about a finite-span rectangular wing mounted upright in a supersonic stream were generated numerically to investigate formation of the tip vortex. Solutions were obtained by integration of the time-dependent three-dimensional compressible Euler and thin-layer laminar and mass-averaged turbulent Navier-Stokes equations. In the turbulent case, a two-equation (k-e) closure model was employed, which included low-Reynolds-number terms and a compressibility correction. Details of the computations are summarized, and a grid resolution study is provided to assess the accuracy of the numerical results. Comparisons are made between solutions of the various governing equations, as well as with experimental data in terms of pressure and Mach-number component distributions in the vortex core.

Measurements and computations of the separation and reattachment locations are reported for the steady flow over a SD7003 airfoil at different angles of attack and chord Reynolds number in the range 2 × 104–4 × 104. The experiments are based on multi-line molecular tagging velocimetry, and the computations employ an implicit large eddy simulation approach. Comparisons of experimental results with current computations and previous experiments point to challenges involved in the experimental determination of the separation bubble characteristics. The results also underline the importance of the facility-dependent freestream turbulence level on the experimental data. The collective effect of experimental uncertainty and facility-dependent issues, examined systematically herein, appear to clarify the discrepancy among the various experimental and computational results. The findings also suggest that accurate characterization of the separation bubble over airfoils at low Reynolds number is more difficult than generally recognized and presents a challenge in comparing results between different experiments, and between experiments and computations. Moreover, this complicates the validation of computational data against experiments within this flow regime.

Numerical computations are carried out to predict the transition generated by excrescence on a platelike geometry in subsonic flow. Both forward-facing and rearward-facing steps of small roughness height are considered in the investigation. These are representative of joints and other surface imperfections on wing sections that disrupt laminar flow, thereby increasing skin friction and drag. Solutions are obtained via a high-fidelity numerical scheme and an implicit time-marching approach on an overset mesh system that is used to represent the steps. Very-small-amplitude numerical forcing is employed to generate perturbations, which are amplified by the geometric disturbances, similar to the physical situation. The flowfield just downstream of the steps is characterized by the growth and breakdown of two-dimensional fluid structures. Because all significant scales of the flow are fully resolved in this region, the solutions there correspond to direct numerical simulations. Further downstream where the flo...