Large Region Of Separated Flow Research Articles

NUMERICAL ANALYSIS AND DIFFUSER VANE SHAPE OPTIMIZATION OF A RADIAL COMPRESSOR WITH THE OPEN-SOURCE SOFTWARE SU2

In recent years, the usage of open-source computational fluid dynamics tools is on a rise both in industry and academia. SU2 is one of these open-source tools. Unlike other open-source alternatives, SU2 is equipped with boundary condition types, solvers and methods that are especially developed for the analysis and design of turbomachinery. The aim of this work is to explore and investigate the capabilities of SU2 in the prediction of performance parameters of radial compressors. Two different single stage shrouded compressor geometries, one with a vaneless diffuser and the other with a vaned diffuser have been investigated with steady state CFD. The compressors were designed by MAN Energy Solutions Schweiz AG. Computational results with SU2 showed a satisfactory agreement with both the experimental data and reference CFD solutions obtained with Fidelity Flow, which is formerly known as Numeca Fine TURBO. Only at the relatively higher mass flow rates the difference between references and SU2 were higher compared to other operating points. After performance parameters were successfully calculated with SU2, the optimization tools that come with SU2 were also used. A 2D adjoint optimization study on the vane of the vaned diffuser was carried out. The study was carried out at a single operating point that is close to choke conditions. The loss generated by the large separated flow region at the suction side of the diffuser vane was reduced by 0.55 % in the optimized geometry using minimal modifications on the existing vane geometry to keep the performance of the compressor intact at other operating points. However, the resulting modification increased the total pressure loss by 0.86 % at one of the design operating points. This performance penalty could be due to the discontinuity in the vane geometry generated by the optimizer. Overall, the study shows that SU2 has the basic numerical schemes and models that are required for the analysis of radial turbomachinery flows and geometry optimization.

Enhancement of Control Jet Performance in Supersonic Crossflows Through Pilot Injection

Injection of a control jet into a supersonic crossflow creates a large region of separated flow in the vicinity of injection, altering the pressure distribution on the parent body, and consequently the effectiveness of the injected jet in producing the desired control forces and moments. In the case of an axisymmetric parent body, this disruption typically results in a loss of effectiveness due to the spillover of the shock structures around the parent body. In this study, a novel technique for enhancing this jet effectiveness on bodies of revolution by means of employing a pilot injection upstream of the main control jet is proposed and evaluated at different external flow conditions. We show that diverting a small fraction () of the primary control jet mass flow rate through an orifice situated five to 10 jet diameters upstream causes an earlier disruption of the external flow, leading to a weaker shock system around the primary injector. This reduces the extent of adverse interaction region and results in a lower spillover of the shock and vortical structures to the bottom side, leading to a significant improvement of the control effectiveness.

Contrarotating Open Rotor Operation for Improved Aerodynamics and Noise at Takeoff

The contrarotating open rotor is, once again, being considered as an alternative to the advanced turbofan to address the growing pressure to cut aviation fuel consumption and carbon dioxide emissions. One of the key challenges is meeting community noise targets at takeoff. Previous open rotor designs are subject to poor efficiency at takeoff due to the presence of large regions of separated flow on the blades as a result of the high incidence needed to achieve the required thrust. This is a consequence of the fixed rotor rotational speed constraint typical of variable pitch propellers. Within the study described in this paper, an improved operation is proposed to improve performance and reduce rotor-rotor interaction noise at takeoff. Three-dimensional computational fluid dynamics (CFD) calculations have been performed on an open rotor rig at a range of takeoff operating conditions. These have been complemented by analytical tone noise predictions to quantify the noise benefits of the approach. The results presented show that for a given thrust, a combination of reduced rotor pitch and increased rotor rotational speed can be used to reduce the incidence onto the front rotor blades. This is shown to eliminate regions of flow separation, reduce the front rotor tip loss and reduce the downstream stream tube contraction. The wakes from the front rotor are also made wider with lower velocity defect, which is found to lead to reduced interaction tone noise. Unfortunately, the necessary increase in blade speed leads to higher relative Mach numbers, which can increase rotor alone noise. In summary, the combined CFD and aeroacoustic analysis in this paper shows how careful operation of an open rotor at takeoff, with moderate levels of repitch and speed increase, can lead to improved front rotor efficiency as well as appreciably lower overall noise across all directivities.

Large-Eddy Simulation of Plasma-Based Turbulent Boundary-Layer Separation 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.

Large-eddy simulation of the flow around a freight wagon subjected to a crosswind

The purpose of the research reported in this paper is to investigate the aerodynamic response of a freight train subjected to a crosswind at 90° yaw angle using large-eddy simulation (LES). The freight train under investigation consists of many identical wagons. The wagons of the train are of the single-stacked container type that is having a box-like shape. In order to reduce the computational cost, attention is confined to the aerodynamic loads and flow structures around a selected single wagon from the middle of the train. The influences of the neighbouring wagons are simulated by imposing spanwise periodicity. The Reynolds number of the flow, based on the time-averaged speed of the crosswind and the height of the wagon from the ground, is 300,000. The standard Smagorinsky model with model constant of 0.1 and the Van Driest damping function, implemented in the commercial CFD package CFX, are used. The assessment of LES accuracy was carried out by performing three different computations using three different meshes; coarse, medium and fine. The fine mesh simulation gives results similar to that of the medium mesh simulation. The LES results showed that the flow moves in the gaps between the containers and under the wagon with high speed resulting in a much more complex flow topology in the wake of the wagon compared to that previously published around passenger trains. The flow separates at the sharp windward edge of the container to form a large separated flow region on the roof of the container. Aerodynamic forces on the container and freight wagon were computed and their time history was used to reveal the characteristic frequencies of the flow motion around the body. Further, the effect of the moving ground on aerodynamic coefficients is investigated by performing simulation on a moving ground. As a result of the floor motion, the side force coefficient was reduced by about 2.5% while the lift force was increased by about 11%.

Synthetic jet을 이용한 스마트 무인기(SUAV) 유동제어 Part 2 : 천이 비행 모드에서 synthetic jet을 이용한 유동제어

스마트 무인기 익형 주위의 유동 구조를 파악하고 이를 바탕으로 synthetic jet을 이용하여 천이 비행 모드에서의 수익하중 감소 여부를 파악하였다. 스마트 무인기의 실제 비행 모드에서의 유동 구조를 분석하여 앞전 및 뒷전에서 발생하는 와류에 의해서 수익하중이 크게 증가함을 밝혔다. 이에 앞전과 뒷전에서 발생하는 유동의 박리를 효과적으로 제어하기 위하여 Part 1의 결과를 바탕으로 0.01c, <TEX>$0.95c_{flap}$</TEX> 지점에 jet을 위치시키고 각각 무차원 주파수(F+)를 0.5, 5로 작동시켜 그에 따른 유동구조 변화와 항력 감소율을 알아보았다. 그 결과 천이 비행 모드에서의 유동 제어를 위해서는 앞전에 위치한 jet만을 작동시킬 경우 가장 효과적으로 수익하중을 감소시킬 수 있음을 밝혔다. 이에 정지 비행 모드에서 뿐만 아니라 천이 비행 모드에서 synthetic jet을 이용하여 유동을 제어한다면 스마트전 비행 모드에서의 비행성능과 안정성을 동시에 향상시킬 수 있을 것이다. In order to reduce the download around the Smart UAV(SUAV) at Transition mode, flow control using synthetic jet has been performed. Many of the complex tilt rotor flow features are captured including the leading and trailing edge separation, and the large region of separated flow beneath the wing. Based on the results of part 1 of the present work, synthetic jet is located at 0.01c, <TEX>$0.95c_{flap}$</TEX> and it is operated with the non-dimensional frequency of 0.5, 5 to control the leading edge and trailing edge separation. Consequently, download is substantially reduced compared to with no control case at transition mode using leading edge jet only. The present results show that the overall flight performance and stability of the SUAV can be remarkably improved by applying the active flow control strategy based on synthetic jet.

CFL3D/OVERFLOW Results for DLR-F6 Wing/Body and Drag Prediction Workshop Wing

A series of overset grids was generated in response to the Third AIAA CFD Drag Prediction Workshop (DPW-III) which preceded the 25th Applied Aerodynamics Conference in June 2006. DPW-III focused on accurate drag prediction for wing/body and wing-alone configurations. The grid series built for each configuration consists of a coarse, medium, fine, and extra-fine mesh. The medium mesh is first constructed using the current state of best practices for overset grid generation. The medium mesh is then coarsened and enhanced by applying a factor of 1.5 to each (I, J, K) dimension. The resulting set of parametrically equivalent grids increase in size by a factor of roughly 3.5 from one level to the next denser level. Computational fluid dynamics simulations were performed on the overset grids using two different Reynolds-averaged Navier-Stokes flow solvers: CFL3D and OVERFLOW. The results were postprocessed using Richardson extrapolation to approximate grid-converged values of lift, drag, pitching moment, and angle of attack at the design condition. This technique appears to work well if the solution does not contain large regions of separated flow (similar to that seen in the DLR-F6 results) and appropriate grid densities are selected. The extra-fine grid data helped to establish asymptotic grid convergence for both the OVERFLOW FX2B wing/body results and the OVERFLOW DPW-W1/W2 wing-alone results. More CFL3D data are needed to establish grid convergence trends. The medium grid was used beyond the grid convergence study by running each configuration at several angles of attack so drag polars and lift/pitching moment curves could be evaluated. The alpha sweep results are used to compare data across configurations as well as across flow solvers. With the exception of the wing/body drag polar, the two codes compare well qualitatively showing consistent incremental trends and similar wing pressure comparisons.

The numerical computation of near-wall turbulent flow over a steep hill

The present work performs a detailed comparison between numerical computations for the flow over a two-dimensional steep hill and some newly obtained laboratory data. Six turbulence models were tested: four eddy-viscosity models ( κ – ε , RNG- ε , κ – ω , SST) and two second-moment models ( SSG– RSM- ε , BSL– RSM- ω ). The experiments were conducted in a water channel and were specially planned such that the large separated flow region that is formed on the lee side of the hill could be well scrutinized. The experimental results include complete profiles of the mean velocity components and of the two-dimensional Reynolds stress tensor and were obtained through the laser Doppler anemometry. A particular concern of this work has been to achieve a detailed experimental and numerical characterization of the near-wall flow region. As such, for most of the measuring stations, at least eight points were located in the viscous sublayer. The work also shows the distribution of wall-shear stress in detail. The ω -equation-based models were observed to perform much better than the ε -equation-based models. The length of separated flow region, mean velocity profiles and wall-shear stress were all reasonably well predicted. The flow properties on the hill top were particularly difficult to describe. The turbulence properties in the reversed flow region were best simulated by the BSL– RSM model.

A comparison of two meshing schemes for CFD analysis of the impulse turbine for wave energy applications

The results of two three-dimensional CFD models of the 0.6 m diameter, 0.6 hub-to-tip ratio Impulse Turbine are compared qualitatively and quantitatively with experimental data for a geometrically similar turbine under similar conditions. One model uses hexahedral cells exclusively to discretise the computational domain, the other uses a hybrid meshing scheme, utilising hexahedral cells at the rotor blade and the guide vanes, tetrahedral cells in the bulk of the remaining computational domain and pyramidal cells to mediate the interface between these two types of cell. Quantitatively, there is a visible difference in the predicted non-dimensional pressure drop across the turbine. The main qualitative difference between the results of the two models is a large region of separated flow extending from the tip of the rotor blade, on the pressure side, to roughly the 50% radial station. This is predicted using the hybrid scheme and that is essentially absent in the hexahedral model. Simple design calculations are used to illuminate the flow regime presented to the rotor by the guide vanes and conclusions are drawn as to the performance of each grid in predicting performance and providing insight into the fluid flow. Both Models produced results in reasonable agreement with experiment. The unstructured, hybrid grid predicted a pressure drop that was greater than the structured hexahedral grid and both grids underpredicted pressure drop to some extent. The hybrid, unstructured grid quality was higher than that of the structured all hexahedral grid. The hybrid, unstructured grid was somewhat easier to create than the structured, all hexahedral grid.

Computations of Steady and Unsteady Low-Speed Turbulent Separated Flows

Numerical computations are presented for two-dimensional steady and unsteady separated flows. The first case is a low-speed, converging-diverging duct with a rapid expansion, creating a large separated flow region. The second case is the Massachusetts Institute of Technology flapping foil experiment, where a stationary hydrofoil is subject to gust loading. Numerical solutions are obtained by solving the incompressible Navier-Stokes equations. These equations are solved in a time accurate manner using the method of artificial compressibility. The Johnson and King turbulence model is employed for modeling the turbulent flow. Modifications to the model are suggested that take into account the normal stress production of energy and the strong adverse pressure gradient associated with separating flows. The performance of the Johnson and King model and its modifications are studied for both steady and unsteady flow conditions

Computation of internal turbulent flow with a large separated flow region

AbstractAn implicit two‐equation turbulence solver, KEM. in generalized co‐ordinates, is used in conjunction with the three‐dimensional incompressible Navier–Stokes solver, INS3D, to calculate the internal flow in a channel and a channel with a sudden 2:3 expansion. A new and consistent boundary procedure for a low Reynolds number form of the κ‐ε turbulence model is chosen to integrate the equations up to the wall. The high Reynolds number form of the equations is integrated using wall functions. The latter approach yields a faster convergence to the steady‐state solution than the former. For the case of channel flow, both the wall‐function and wall‐boundary‐condition approaches yield results in good agreement with the experimental data. The back‐step (sudden expansion) flow is calculated using the wall‐function approach. The predictions are in reasonable agreement with the experimental data.

Navier–Stokes Solutions of Unsteady Flow in a Compressor Rotor

In order to achieve more accurate predictions of unsteady flow in a transonic compressor rotor an existing numerical approach has been modified by incorporating a turbulence model. The computations are performed by solving the complete time-dependent compressible Navier–Stokes equations using MacCormack’s explicit finite difference algorithm. These equations are solved for the flow through two adjacent rotor blades at a streamsurface near the blade tip subjected to the wakes emitted from upstream stators. At this radial location the flow enters the blade passage at an absolute Mach number of 0.66. The high blade curvature at this radial location produces a large region of separated flow on the suction surface with laminar flow. To more accurately resolve the features of this flow separation the Baldwin–Lomax algebraic eddy-viscosity turbulence model is incorporated into the numerical procedure in regions near the blade surface. The unsteady flow features are represented at the inflow boundary through the use of characteristic variables involving the upstream and downstream running Riemann invariants and the entropy variation expressed in terms of the total pressure profile. At the outflow boundary the concept of a “second throat” or choke point is implemented in conjunction with supersonic outflow conditions. The results are compared with numerical results obtained without the use of a turbulence model (laminar) for a single blade passage. Improved agreement with limited experimental data is also noted.

Fluid Stresses in the Vicinity of Disk, Ball, and Tilting Disk Prosthetic Heart Valves From In-Vitro Measurements

Experiments were conducted in a steady flow test apparatus with an axisymmetric aortic shaped test chamber using hot-wire anemometry to obtain local momentum transfer and turbulence data in the vicinity of the model disk, Kay-Shiley disk, Starr-Edwards ball and Bjork-Shiley tilting disk prosthetic heart valve configurations. These data for Reynolds numbers of 2000, 4000, and 6000 were used to locate regions of high stress where erythrocytes may be lysed or severely strained. Nondimensionalized velocity profiles downstream of the valves indicated a similarity of behavior for all Reynolds numbers at an axial location. Velocity gradients of magnitudes potentially damaging to erythrocytes were found. These gradients were largest for disk-type occluders. Large regions of separated flow were found to occur behind the valve sewing ring, distal to the valve occluder, and along the test chamber wall distal to the valve for all valves tested. Relatively high turbulence was found to exist distal to all valves tested. Turbulent stresses of magnitudes potentially hemolytic were measured.

Nose pressure distribution and separation on an inclined axisymmetric body.

When a body of revolution is inclined to a small angle of attack, a large region of separated flow occurs on the leeward side of the body. The effect of this viscous phenomenon on the surface pressure distribution of a hemispherical nose with a cylindrical afterbody is examined.