Skin Friction Coefficient Distributions Research Articles

Inflatable aerodynamic decelerators for CubeSat reentry and recovery: Surface properties

In this investigation, Direct Simulation Monte Carlo Method is used to compute the surface properties of inflatable aerodynamic decelerators (IAD) applied for the reentry, recovery, and reuse of cubeSats. Surface properties calculation is of fundamental importance for flexible thermal protection system materials selection that withstand the harsh reentry environment. Furthermore, simulations are extremely important to detect possible hot spots at the IAD before design and flight. In this work, three IAD configurations are exposed to a rarefied hypersonic flow with freestream conditions that correspond to those found at reentry conditions at 105 km altitude. All geometries are assumed to be axisymmetric and have a maximum radius of 0.3 m from the center line to the shoulder edge. The main differences between the geometries are the angle between the IAD and cubeSat body and the IAD curvature for geometry 3. According to the results, all IAD geometries are effective in protecting the cubeSat during the reentry phase. It is shown that the IAD geometry has a significant influence on the heat transfer, pressure, and skin friction coefficient distribution over the flexible heat shield. The maximum heat transfer coefficient computed for geometry 1, 2, and 3 is 0.90, 0.97, and 0.84, respectively.

On the Fidelity of RANS-Based Turbulence Models in Modeling the Laminar Separation Bubble and Ice-Induced Separation Bubble at Low Reynolds Numbers on Unmanned Aerial Vehicle Airfoil

The operational regime of Unmanned Aerial Vehicles (UAVs) is distinguished by the dominance of laminar flow and the flow field is characterized by the appearance of Laminar Separation Bubbles (LSBs). Ice accretion on the leading side of the airfoil leads to the formation of an Ice-induced Separation Bubble (ISB). These separation bubbles have a considerable influence on the pressure, heat flux, and shear stress distribution on the surface of airfoils and can affect the prediction of aerodynamic coefficients. Therefore, it is necessary to capture these separation bubbles in the numerical simulations. Previous studies have shown that these bubbles can be modeled successfully using the Large Eddy Simulation (LES) and Direct Numerical Simulation (DNS) but are computationally costly. Also, for numerical modeling of ice accretion, the flow field needs to be recomputed at specific intervals, thus making LES and DNS unsuitable for ice accretion simulations. Thus, it is necessary to come up with a Reynolds-Averaged Navier–Stokes (RANS) equation-based model that can predict the LSBs and ISBs as accurately as possible. Numerical studies were performed to assess the fidelity of various RANS turbulence models in predicting LSBs and ISBs. The findings are compared with the experimental and LES data available in the literature. The structure of these bubbles is only studied from a pressure coefficient perspective, so an attempt is made in these studies to explain it using the skin friction coefficient distribution. The results indicate the importance of the use of transition-based models when dealing with low-Reynolds-number applications that involve LSB. ISB can be predicted by conventional RANS models but are subjected to high levels of uncertainty. Possible recommendations were made with respect to turbulence models when dealing with flows involving LSBs and ISBs, especially for ice accretion simulations.

Free convection in a differentially heated square cavity filled with low-Prandtl-number materials: Numerical studies using transition shear stress transport model

Computing the flow field under free convection in a cavity becomes particularly challenging for low-Prandtl-number (Pr) fluids typically encountered for liquid metals. The objective of the present study is to investigate the natural convection process in a differentially heated square cavity employing the transition shear stress transport (SST) model for the Prandtl number Pr∈[0.001,0.1] and the Rayleigh number Ra∈[104,1010]. The mean flow field is visualized through streamlines, isotherms, non-dimensional velocity, and temperature profiles, turbulence intensity, contours of intermittency (γ) times, the production of turbulent kinetic energy (Pκ), and distribution of skin friction coefficient and the Nusselt number (Nu). The transition SST model can capture the mean flow field and thermal transport over the entire parametric regime successfully. An average Nusselt number (Nu¯) on the hot wall is found to scale with a certain power (n) of the Boussinesq number (Bo), the product of Ra and Pr. The value of n is 0.18 for Ra up to 106 and 0.25 for higher Ra.

Numerical Study of Aerodynamic Characteristics of Airflow Around NACA 0012 and NACA 4412 Airfoils at Re = 170000

In this paper, a numerical study of the aerodynamic characteristics using the CFD method of the two-dimensional flow around NACA 0012 and NACA 4412 airfoils at -10 degrees to 25 degrees of angle of attack operating at Re 170000 is presented. The formation of flow patterns, pressure distribution, pressure coefficient (Cp), lift coefficient (Cl), drag coefficient (Cd), moment coefficient (Cm), and skin friction coefficient distribution along the chord The Y+ wall value curve is a validation that the mesh formation is in accordance with the requirements, i.e., a Y+ wall value below 10 for viscous dominant flows and for Spalart-Allmaras turbulent model viscous. The geometry of NACA 0012 and NACA 4412 airfoils is created using SolidWork, and CFD analysis is carried out using Ansys Workbench 2018. The results show that the stall condition at NACA 0012 is at 15 degrees of angle of attack, while at NACA 4412 it is at 16 degrees. Under stall conditions, Cl of NACA4412 is approximately 13 percent higher than NACA 0012, while the value of Cd shows almost the same value for both NACA 0012 and NACA 4412.

A high-order fluid–structure interaction framework with application to shock-wave/turbulent boundary-layer interaction over an elastic panel

Within this work, a loosely-coupled high-order fluid–structure interaction (FSI) framework is developed in order to investigate the influence of an elastic panel response on shock-wave/turbulent boundary-layer interaction (SWTBLI). Since high-order methods are expected to determine the future of high-fidelity numerical simulations, they are employed in the construction of both fluid and structure solvers. Specifically, a split-form arbitrary Lagrangian-Eulerian discontinuous Galerkin spectral element method is employed in the fluid solver and a Legendre spectral finite element method in the structure solver. Shock capturing by an improved adaptive filter method, which confines the filtering effect to the vicinity of shocks, is found to be well-behaved in accuracy, efficiency and flexibility.After being validated by two benchmark FSI problems, the developed FSI framework is applied to simulate SWTBLI over an elastic panel. A comparison with a previous simulation of SWTBLI over a rigid panel reveals that: (1) A larger amplitude of the pressure variation, observed on the elastic panel surface, implies a larger threat to the structural integrity; (2) The shock-induced separation flow over the elastic panel changes both in size and shape, leading to a different skin-friction coefficient distribution; (3) A new low-frequency flow unsteadiness of the same magnitude as the elastic panel vibration is detected, which may affect the flow dynamics inside the turbulent boundary layer.

Magnetized supercritical third-grade nanofluid flow from a vertical cylinder using a Crank–Nicolson implicit scheme

The current study investigates the supercritical convective radiative buoyancy-driven flow and heat transfer in external coating flow of an electrically conducting viscoelastic third-grade nanofluid from hotmoving/stationary cylinder under constant radial magnetic field. A new computational thermodynamic model has been proposed to study the supercritical behavior of third-grade aqueous nanofluid in terms of supercritical water. The Redlich–Kwong equation of state (RK-EOS) has been deployed to calculate thermal expansion coefficient for free convective flow of supercritical nanofluid in terms of temperature, compressibility factor and pressure. A validation test has been conducted for RK-EOS with available experimental results. A well-tested unconditionally stable Crank–Nicolson scheme has been implemented to obtain numerical solutions. Graphical results for flow variables, heat transfer and skin-friction coefficient distributions are presented for variation of nanoscale, rheological, magnetic, radiative and thermodynamic parameters. Transient velocity is reduced whereas temperature is elevated with amplified values of third-grade fluid parameter, reduced pressure, reduced temperature, and decreased values of volume fraction of nanofluid for a stationary cylinder under supercritical conditions. Special cases of the current model validated against previously published results. Important applications of the current study include nuclear reactor vessels, deposition of smart (functional magnetic) nano-coatings and solar collector energy systems.

Turbulent flow simulations of the common research model on Cartesian grids using recursive fitting approach

In this study, we propose a new method that uses the recursive fitting method and wall function for the three-dimensional turbulent flow simulation. The recursive fitting method based on the Cartesian grid method is employed for the automatic generation around two- and three-dimensional geometries. The grid approximately represents the concave and convex features, i.e., a wing–fuselage juncture and wing trailing edges. The wall boundary conditions can be imposed on the wall faces; thus, the conservation laws are strictly satisfied. Meanwhile, using the wall function allows turbulent boundary layers to be accurately reproduced. The turbulent flow simulation using the wall function is straightforward because the flow variables at the first cell centroid are employed for friction velocity estimation. However, the distance between the body surface and the first cell centroid varies since cells with different shapes are generated. Therefore, the simulation using the proposed method is validated through turbulent flow simulations in two dimensions. The results show the accurate predictions of the turbulent boundary layers. To investigate the capacity of the proposed method, three-dimensional simulation around aircraft at a transonic cruise condition is conducted. The automatically generated grids by the proposed method approximately represent the geometric features. The results of flow simulation show that the surface pressure and skin friction coefficient distributions agree with those of the reference simulation on a conventional body-fitted grid. Further, predicted pressure and viscous drag coefficients are reasonable compared with those of the reference simulation.

Importance of induced magnetic field and exponential heat source on convective flow of Casson fluid in a micro-channel via AGM

The study of the natural convective flow of a fluid in the presence of an induced magnetic field has always been of considerable importance due to its many applications in various areas of science, technology, and industry, such as the operation of magnetohydrodynamic generators. This study addresses an analysis of exponential heat source and induced magnetic field on the second-class convection of Casson fluid in a microchannel. The flow is in a vertical microchannel organized by two vertical plates. The answer to governing equations has been grabbed for temperature field, induced magnetic field, and velocity via Akbari-Ganji's method (AGM). Nusselt number, skin friction coefficient, and current density are approximated. Graphs that describe the conclusion of influential physical variables on velocity, temperature, current density, induced magnetic field, and skin friction coefficient distributions are shown. Comparison of results with numerical method (Runge-Kutta-Fehlberg, RKF-45), homotopy perturbation method (HPM), and AGM confirms the accuracy of answers obtained with AGM.

Numerical analysis of the effect of train length on train aerodynamic performance

The improved delayed detached eddy simulation is adopted in the present study to investigate the influence of the train length on its aerodynamic performance. The low y+ wall treatment and the cubic constitutive relation are adopted to resolve the viscous flows and model the anisotropic turbulence within the boundary layer. The analysis implied that the distribution region and intensity of velocity fluctuation are strengthened, resulting in a larger turbulence kinetic energy distribution and a higher boundary layer thickness as the train length increases. A reduction in the streamwise velocity and the negative pressure with the increasing train length on the tail train is observed, resulting in lower drag and lift coefficients. As the length of the train increases, both the mean and instantaneous slipstream velocities are increased. The boundary layer thickness and the skin friction coefficient are compared with flat plate theory, reduced-scale, and full-scale experiments, proving the ability of numerical simulation to model the boundary layer velocity profile and skin friction coefficient distribution correctly. The wake structures are identified by the Spectral Proper Orthogonal Decomposition method, the dominant mode frequency decreases, and the wavelength becomes larger as the length of the train becomes longer due to the thickening boundary layer.

Propeller Position Effects over the Pressure and Friction Coefficients over the Wing of an UAV with Distributed Electric Propulsion: A Proper Orthogonal Decomposition Analysis

New propulsive architectures, with high interactions with the aerodynamic performance of the platform, are an attractive option for reducing the power consumption, increasing the resilience, reducing the noise and improving the handling of fixed-wing unmanned air vehicles. Distributed electric propulsion with boundary layer ingestion over the wing introduces extra complexity to the design of these systems, and extensive simulation and experimental campaigns are needed to fully understand the flow behaviour around the aircraft. This work studies the effect of different combinations of propeller positions and angles of attack over the pressure coefficient and skin friction coefficient distributions over the wing of a 25 kg fixed-wing remotely piloted aircraft. To get more information about the main trends, a proper orthogonal decomposition of the coefficient distributions is performed, which may be even used to interpolate the results to non-simulated combinations, giving more information than an interpolation of the main aerodynamic coefficients such as the lift, drag or pitching moment coefficients.

Numerical analysis of heat transfer between multiple jets and flat moving surface

The current work represents the numerical analysis of heat transfer between single and three inline circular jets impinging on an orthogonally flat moving surface. The analysis is carried out at different surface-to-jet velocity ratios (r) = 0.25,0.50,0.75, 1.0 and 2.0 and three jet-to-jet spacing (s) = 2D,4D and 6D for multiple jets. All cases have been performed at the jet-to-surface distance (H) = 6D and Reynolds number = 26000 where D is a jet diameter. The SST-k-ε turbulence model is used for solving turbulent flow problem. The influence of surface velocity on heat transfer as well as flow field over the surface has been analyzed. The results show that localized Nusselt number varies with velocity ratio and stagnation values of Nusselt number lower with velocity ratios. For single jet, spanwise heat transfer of surface dominated by the surface motion for r > 0.75, and the associated flow field also modify with surface motion. For multiple jets, jet-to-jet spacing has an influence in secondary peak in Nusselt number between neighboring jets at low values of velocity ratio. For a single jet, the average Nusselt number remains almost constant for 0 ≤ r ≤ 0.75 and then increases drastically after r > 0.75 but for multiple jets, it increases smoothly with r. Skin friction coefficient distribution at the symmetry line of the target surface is uniform on both sides of the stagnation point and it becomes more uneven with increasing velocity ratio (r).

Aerodynamic characteristics of a high-speed train exposed to heavy rain environment based on non-spherical raindrop

The aerodynamic characteristics of high-speed trains operating in severe environment have gained an increasing interest in the past few years. As rain is a common weather phenomenon, understanding its effects on the aerodynamic characteristics of the high-speed train provides valuable information for the design of a train. In the present paper, the aerodynamic model of a high-speed train operating in a dry condition is first set up, and the simulation results are compared with the existing wind tunnel test data to validate accuracy of the mesh resolution and CFD technique. For the simulation of raindrops, the applicability of spherical raindrop assumption in numerical simulations is discussed in this paper. The results demonstrate that, when the equivalent spherical diameters are greater than 2 ​mm, the simulation errors of terminal velocity of raindrops exceed 3%. The non-spherical characteristics of raindrops need to be considered for large particle sizes, and the particle sphericity for various equivalent spherical diameters is investigated. After that, a two-way coupled Euler-Lagrange approach is developed to evaluate the aerodynamic characteristics of a high-speed train in heavy rain environment based on non-spherical raindrop assumption. The flow around the train, pressure coefficient distribution, skin friction coefficient distribution, and aerodynamic force coefficients of the high-speed train exposed to heavy rain environment are then investigated. It is found that the aerodynamic coefficients of the high-speed train increase approximately linearly with the rainfall intensity. The explicit equations for predicting the aerodynamic force coefficient which is correlated with train speed and rainfall intensity are proposed in this study.

Feasibility of skin-friction field measurements in a transonic wind tunnel using a global luminescent oil film

The feasibility of skin-friction field measurements using the global luminescent oil-film skin-friction field estimation method was evaluated for a challenging case of a supercritical airfoil model under transonic wind-tunnel conditions (freestream Mach number of 0.72) at a high Reynolds number (10 million, based on the model chord length). The oil-film thickness and skin-friction coefficient distributions were estimated over the airfoil model upper surface for a range of angles of attack (from -0.4^{circ } to 2.0^{circ }), thus enabling the study of different boundary-layer stability situations with laminar–turbulent transition, including cases with shock-wave/boundary-layer interaction. Conventional pressure measurements on the surface and in the wake of the model as well as Schlieren flow visualizations were conducted to support the oil-film based investigations. In the laminar-flow regions, the oil-film thickness could be generally kept below the critical limit of roughness that would induce premature boundary-layer transition. The skin friction in this region could be estimated with a moderate confidence level, as confirmed for portions of the chord by the reasonable agreement with numerical data obtained via laminar boundary-layer computations. Moreover, the location of transition onset was evaluated from the skin-friction estimations with relatively low uncertainty, thus enabling the examination of the transition location evolution with varying angle of attack. The estimated locations of transition onset were shown to be in general agreement with reference transition locations measured via temperature-sensitive paint. On the other hand, the oil-film thickness in the turbulent-flow regions was larger than the height of the viscous sublayer, which led to an hydraulically rough surface with increased skin friction, as compared to the clean configuration. For this reason, quantitative skin-friction estimations were not feasible in the turbulent-flow regions. The global effects of the oil-film setup on the flow around the airfoil were evaluated from the estimations of the aerodynamic coefficients. In particular, it was shown that the presence of the specific base coat used for the application of the oil film already induced a significant increase in airfoil drag, as compared to the clean configuration, whereas a thin oil film led to negligible or small additional increases in drag. Based on the present observations, considerations for the further improvement of the global luminescent oil-film skin-friction field estimation method in transonic flow experiments at high Reynolds numbers are elucidated.Graphic abstract

Evolution Characteristics of Separated Vortices and Near-Wall Flow in a Centrifugal Impeller in an Off-Designed Condition

Flow separation is undesirable and lowers the efficiency of centrifugal impellers. In this study, the evolution characteristics of separated vortices in a centrifugal impeller are studied under the off-designed flow rate condition. Unsteady Reynolds-Averaged Navier–Stokes (URANS) with standard k-ε turbulent model is applied to simulate the alternating stall in the six-blade centrifugal impeller. We present and analyze the distributions of pressure gradient (either adverse or favorable) and skin friction coefficients on both sides of the blade for the stalled and unstalled passages to study the relationship between pressure gradient and separation of boundary layer flow. The evolution of skin friction coefficient is also presented at various axial cross sections. Numerical results reveal that, for the stalled passage, the increase in adverse pressure gradient on the pressure surface near the middle of the blade (S/S0 = 0.4) is much larger than that of the suction surface during a vortex formation cycle. The skin friction coefficient on the pressure surface also increases in magnitude sharply and the variation shows a peak-valley trend, while the coefficient on the suction surface increases slowly. Comparing the distribution of skin friction coefficient on the pressure surface of the same blade at different axial cross sections, it is found that the skin friction coefficient notably increases at S/S0 = 0.6 on the middle axial cross section (Z/b2 = 0.5). For the unstalled passage, both the pressure and suction surfaces produce favorable pressure gradients. The skin friction coefficient on the pressure surface shows an increasing trend around S/S0 = 0.5, and a large vortex can be seen at the exit of the impeller. The variation of skin friction coefficient on the suction surface is relatively mild; thus, the flow is relatively stable. It is clarified that the effect of adverse pressure gradient and wall shear stress jointly cause separation of the boundary layer; thus, the separated vortices are generated in the rotating impeller and deteriorate the performance of the impeller.

Plasma-based flow control for low-pressure turbines at low-Reynolds-number

Purpose Reynolds number in small-size low-pressure turbines (LPT) can drop below 2.5 · 104 at high altitude cruise, which in turn can lead to laminar boundary layer separation on the suction surface of the blades. The purpose of this paper is to investigate the potential of an alternate current (AC)-driven Single Dielectric Barrier Discharge Plasma Actuator (AC-SDBDPA) for boundary layer control on the suction side of a LPT blade, operating at a Reynolds number of 2 · 104.Design/methodology/approach Experimental and numerical analyses were conducted. The experimental approach comprised the actuator testing over a curved plate with a shape designed to reproduce the suction surface of a LPT blade. A closed loop wind tunnel was employed. Sinusoidal voltage excitation was tested. Planar velocity measurements were performed by laser Doppler velocimetry (LDV) and particle image velocimetry (PIV). The device electrical power dissipation was also calculated. Computational fluid dynamics (CFD) simulations using OpenFOAM© were conducted, modelling the actuator effect as a body force calculated by the dual potential algebraic model. Unsteady RANS (Reynolds Averaged Navier-Stokes equations), also known as URANS approach, with the k-ε Lam-Bremhorst Low-Reynolds turbulence model was used.Findings The AC-SDBDPA operation brought to a reduction of the separation region; in particular, the boundary layer thickness and the negative velocity values decreased substantially. Moreover, the flow angle in both the main flow and in the boundary layer was reduced by the plasma control effect. The actuation brought to a reduction of the 17 per cent in the total pressure loss coefficient. The pressure coefficient and skin friction coefficient distributions indicated that under actuation the reattacnment point was displaced upstream, meaning that the flow separation was effectively controlled by the plasma actuation. Adopting slightly higher actuation parameters could bring to a full reattachment of the flow.Practical implications The work underlines the potentialities of an AC-SDBDPA to control separation in LPTs of aeroengines.Originality/value The present work sets a methodological framework, in which the validated procedure to obtain the body force model combined with CFD simulations can be used to study a configuration with multiple actuators allocated in array without requiring further experiments.

Large-eddy simulation of shock-wave/turbulent boundary layer interaction with and without SparkJet control

The efficiency and mechanism of an active control device “SparkJet” and its application in shock-induced separation control are studied using large-eddy simulation in this paper. The base flow is the interaction of an oblique shock-wave generated by 8° wedge and a spatially-developing Ma=2.3 turbulent boundary layer. The Reynolds number based on the incoming flow property and the boundary layer displacement thickness at the impinging point without shock-wave is 20000. The detailed numerical approaches were presented. The inflow turbulence was generated using the digital filter method to avoid artificial temporal or streamwise periodicity. The numerical results including velocity profile, Reynolds stress profile, skin friction, and wall pressure were systematically validated against the available wind tunnel particle image velocimetry (PIV) measurements of the same flow condition. Further study on the control of flow separation due to the strong shock-viscous interaction using an active control actuator “SparkJet” was conducted. The single-pulsed characteristic of the device was obtained and compared with the experiment. Both instantaneous and time-averaged flow fields have shown that the jet flow issuing from the actuator cavity enhances the flow mixing inside the boundary layer, making the boundary layer more resistant to flow separation. Skin friction coefficient distribution shows that the separation bubble length is reduced by about 35% with control exerted.

LARGE-EDDY SIMULATION OF SHOCK-WAVE/TURBULENT BOUNDARY LAYER INTERACTION AND ITS CONTROL USING SPARKJET

Large-eddy simulation (LES) of an oblique shock-wave generated by an 8° sharp wedge impinging onto a spatially-developing Mach 2.3 turbulent boundary layer and their interactions has been carried out in this study. The Reynolds number based on the incoming flow property and the boundary layer displacement thickness at the impinging point without shock-wave is 20,000. The detailed numerical approaches are described and the inflow turbulence is generated using the digital filter method to avoid artificial temporal or streamwise periodicity. Numerical results are compared with the available wind tunnel PIV measurements of the same flow conditions. Further LES study on the control of flow separation due to the strong shock-viscous interaction is also conducted by using an active control actuator “SparkJet” concept. The single-pulsed characteristics of the control device are obtained and compared with the experiments. Instantaneous flowfield shows that the “SparkJet” promotes the flow mixing in the boundary layer and enhances its ability to resist the flow separation. The time and spanwise averaged skin friction coefficient distribution demonstrates that the separation bubble length is reduced by maximum 35% with the control exerted.

Application of various RANS based models towards predicting turbulent slot jet impingement

Numerical simulation near the stagnation zone for an impinging jet is challenging because of the presence of sharp gradient due to the change of velocity from zero to the system scale velocity. A comparative study of four Reynolds averaged Navier–Stokes (RANS) based turbulence models has been done by numerically simulating a two-dimensional turbulent slot jet impinging air normally on a flat plate. Nozzle-to-plate spacings equal to 4 and 9.2 have been taken as two test cases with nozzle exit Reynolds number of 2×104. The turbulence intensity at the exit of the nozzle is kept constant at 1%. The capability to predict the fluid flow of the standard k−ε model, the low Reynolds number (LRN) k−ε models proposed by Launder and Sharma (LS) and Yang and Shih (YS), and the standard k−ω model have been evaluated by comparing the computational results with the experimental data available in the literature. The numerically obtained results have been compared with the experimental results on impingement surface pressure, free jet mean velocity profiles, spanwise velocity profiles, spanwise distribution of root mean square velocity fluctuation, skin friction coefficient distribution, non-dimensional velocity profile in wall coordinates. The standard k−ε model predicted the largest recirculation loop among the turbulence models used. The skin friction coefficient distribution is reproduced by YS and k−ω model for both the spacings whereas the other LRN model LS and the standard k−ε model are unable to predict the results. The linear law of the wall is found to be valid up to Y+≃4.0 in the recirculation region. The velocity profiles in wall coordinates in the recirculation region are behaving similar to those in the wall jet region of an impinging jet and the developed region of a wall jet.

A DPIV study on the effects of separation distance upon the vortical behaviour of jet–cylinder impingements

An experimental investigation based on laser-induced fluorescence (LIF) and digital particle image velocimetry (DPIV) was conducted to provide a better understanding into the effects of jet–cylinder separation distance on the vortical structures and dynamic behaviour resulting from jet impingements upon convex cylinders. Separation distances of 1, 2 and 4 jet diameters were investigated, while cylinder–jet diameter ratio was maintained at 2 throughout. LIF and DPIV results show that jet ring-vortex initiation, wall-separated vortex initiation, vortex dipole formation and vortex separation occur further away from the impingement point as the separation distance decreases. Varying the separation distance also influences the recirculating wake region size at the cylinder lee sides, as well as producing two distinct flow modes between adjacent vortex dipoles along the cylinder straight edges. Mean and instantaneous skin friction coefficient distributions determined from DPIV results not only showcase the different effects of separation distance on the wall shear stress, but also illustrate how the flow dynamics influence the local wall shear stress levels. On the other hand, wall-separated vortex initiation, vortex dipole formation and separation events incur little distinctive changes upon the local wall shear stress distributions that may ease their unique identification.

二次元オフセット噴流の流動特性と熱伝達

The flow field of an offset jet is complex and is frequently found in many engineering devices. In this paper, the fluid flow and heat transfer characteristics of two-dimensional offset jet are investigated experimentally. The offset jet is produced by the flow of air which issue from the end of a long parallel channel. The exit Reynolds number Re is changed in 1.5×103 to 7.0×103, and the offset ratio H/h (H : step height, h : channel height) is changed in 0.5, 1.0 and 1.5. The wall shear stress is measured using a micro flow sensor. Flow visualization is carried out using a CCD camera and laser light. The velocity profiles are measured by a PIV system. When the flow at the channel exit is laminar, fluid in the outer region has a counter-clockwise vorticity and fluid in the inner region has a clockwise vorticity, and a lifting-off of the wall jet flow into the ambient fluid is occurred. The Strouhal number of the vortex frequency is almost constant against the offset ratio. The formation of those vortices is not seen at the turbulent flow. Two peaks exist in the wall static pressure, Nusselt number and skin friction coefficient distributions of the laminar flow with H/h = 0.5 and 1.0. The 1st peak appears near the reattachment point, and the 2nd peak appears in the neighborhood of the lifting-off of the wall jet flow. The two peaks coalesce into one at H/h = 1.5. As the offset ratio increases, the reattachment point moves to downstream direction and Nu becomes large both the laminar and turbulent flows. The reattachment point of the turbulent flow moves to upstream direction compared with the laminar flow because Coanda-effect is strengthen.