Freestream Direction Research Articles

Effect of the yaw angle on the riblet performance

Direct numerical simulations are performed to investigate the effect of the yaw angle (α) on the drag-reduction performance by riblets in a turbulent channel flow. The yaw angles of α=0° and 90° correspond to the riblets parallel and perpendicular to the free-stream direction, respectively. It is well known that maximum drag reduction occurs at α=0°, and the drag-reduction performance by riblets is degraded with increasing α. We obtain the critical yaw angle below which drag reduction can be achieved and investigate the corresponding mechanism. From a parametric study, drag reductions are achieved for α<20°. As α increases, the form drag increases more rapidly than the skin-friction drag decreases, resulting in the increase in the total drag. At drag reducing and increasing yaw angles, the mean velocity profiles in wall units shift upward and downward, respectively, and the latter profile suggests that the drag-increasing riblet is similar to a rough surface. With increasing α, the wall-normal and spanwise velocity fluctuations and Reynolds shear stress increase, and stronger near-wall vortices are generated. By applying a scaling analysis to the momentum equation in the riblet-perpendicular direction and the FIK (Fukagata–Iwamoto–Kasagi) identity [Fukagata et al., “Contribution of Reynolds stress distribution to the skin friction in wall-bounded flows,” Phys. Fluids 14, L73–L76 (2002); Peet and Sagaut, “Theoretical prediction of turbulent skin friction on geometrically complex surfaces,” Phys. Fluids 21, 105105 (2009); Bannier et al., “Riblet flow model based on an extended FIK identity,” Flow Turbul. Combust. 95, 351–376 (2015)] to the riblet-parallel direction, the total drag is expressed as a linear function of sin2α.

An experimental investigation of the flow control of a circular cylinder in near wake with parallel plates at Re = 7500

The vortex street suppression in the near wake region of the cylinder with parallel plates was investigated using particle image velocimetry (PIV) at Reynolds number of Re = 7500. The rigid plates are designed as flat, parallel with the free-stream direction, and located on both sides of the cylinder. Two different plate lengths (0.5D and 1D based on the cylinder diameter, D) were examined at gap ratios of g/D = 0.1, 0.3, 0.5, 0.7, and 0.9, respectively. Proper Orthogonal Decomposition (POD) analyses were also employed to reveal vortex mechanism in the near wake region. The gap ratio significantly influences the interaction of the vortices. At low gap ratios, the plates directly affect the development of shear layers on both side of the cylinder. The shear layers elongate downstream with a reduction of the vortex shedding frequency. Moreover, Reynolds shear stress, vortex shedding instabilities and vortex street are suppressed in the near wake region with the help of using parallel plates. On the other hand, vortex shedding reinforces instabilities at higher gap ratios. Besides, the most effective gap ratio is different for each the plate lengths to suppress instabilities.

Effect of Roll Angle Configurations of a Reverse Delta Type Add-on Device on Wing Tip Vortex Alleviation

An add-on device in the shape of a reverse delta has shown the ability to alleviate wake vortices. The present work studies the interaction of the wing tip vortex and the reverse delta type add-on device vortices. This paper looks at the vortex interactions generated downstream of the wing tip in planes perpendicular to the free stream direction and their dependence on roll angles φ at a mean chord-based Reynolds number of Rec=2.75×105. The study reveals that the add-on device causes a reduction in the tangential velocity V and vorticity of the resultant vortex by up to 44.1% and 59.4%, respectively. Also, it is found that the resultant vortex core radius increased by 305%. The results indicate that the reverse delta type add-on device implants countersign vorticity into the wing tip vortex and modifies its roll-up process.

Flow control over a circular cylinder using a slot and axially arranged holes

Circular cylinder flow is controlled using a slot and axially arranged holes (AAH) at a Reynolds number of 32000. When the slot and AAH are placed at small incidence angles from the free-stream direction, the passive jet generated on the leeward cylinder side pushes the wake vortex downstream, reducing the drag compared to the baseline cylinder. If the incidence angle is larger than the critical angle, the alternate blowing and suction induced by the slot and AAH promote a boundary-layer transition on one side and a shear-layer transition on the other. The earlier transition to turbulent flow substantially shortens the separated shear layers on both sides, shifting the wake formation region toward the base and leading to an increase in drag. The AAH has a higher critical incidence angle than the same-size slot, as the passive jet developed from the AAH undergoes a transition to alternate blowing and suction at a higher incidence angle than the slot. Just before the incidence angle of the AAH reaches the critical angle, the weak jet flow induced on the windward cylinder side delays the boundary-layer separation through early separation and laminar reattachment without shortening the separated shear layer, thereby allowing additional drag reduction.

Graph machine learning for predicting wake interaction losses based on SCADA data

Six wind turbine generators (WTGs) within a larger wind farm were chosen to model wake interaction losses by means of graph machine learning. 10-minute averaged measurement data values at each turbine over a period of two years were available. Based on the measured power output and ”free-stream” (undisturbed) wind conditions, determined at an upstream WTG unaffected by wakes, a data-driven free-stream power curve model was built. The cluster of WTGs was then represented by a graph data structure in the form of an adjacency matrix, where the turbines are nodes, which in turn are connected by edges. The edges are represented by weightings, indicating the strength of wake interactions between connected turbines. To learn the weightings, the predicted power outputs of the free-stream power curve model were compared to the measured power outputs and then optimised to account for these differences, which are attributed to wake interaction losses. After that a new data-driven model was trained, with free-stream wind speed, direction and yaw misalignment as features and the optimised weightings as targets. It was shown that the model was able to predict power curves that show clear patterns due to wake effects and reduced the root mean squared error, based on the measured and predicted values, by 16.29% compared to the free-stream power curve model.

Analysis of stagger effects on Busemann supersonic biplane airfoil in shock tube tests by point diffraction interferometer method

The Busemann biplane concept is proposed for low-boom low-drag supersonic aircraft. Studies on the supersonic biplane have been focused on supersonic flow, whereas the understanding of transonic aerodynamic characteristics is essential for aircraft design. Significantly, the complicated flow field with a shock and boundary layer interaction between Busemann biplane elements in transonic flow remains unclear. In the present study, the transonic aerodynamic characteristics of the two-dimensional Busemann biplane are investigated by shock tube tests and numerical simulations, focusing on the flow field between the biplane elements. In addition, the stagger effects, which change the relative position of the lower and upper elements in the freestream direction to avoid the choked flow between the biplane elements, are also analyzed. A shock tube system creates a test flow at Mach numbers of 0.6 to 0.8, and the Reynolds number is set at 2.85 × 105. The Point Diffraction Interferometer (PDI) method is applied to determine the density distribution around the model and the pressure coefficient on the surfaces of the biplane elements. Symmetrical density distribution is confirmed in the Busemann biplane (the baseline model, or the biplane with no stagger). A normal shock wave is generated between the biplane elements. When the Mach number increases, the normal shock wave moves downstream. The normal shock position is quantitatively stable with small fluctuation amplitudes. The fringes are strongly curved near the surfaces of the biplane elements due to the flow separation near the airfoil surfaces. When the Mach number increases, the separation area increases. When the Mach number increases, the peak value of the pressure coefficient on the surfaces of the biplane elements increases. On the other hand, the normal shock wave between the biplane elements is eliminated in the staggered models. The changes in density distribution between the biplane elements are minor compared to the case of the baseline model, which indicates that the stagger reduces the choked flow between the biplane elements. The peak value of the pressure coefficient on the wing surface increases when the stagger value increases. The stagger effects on reducing the total drag coefficient of the biplane models are confirmed.

Carrier-Domain Method for high-resolution computation of time-periodic long-wake flows

We are introducing the Carrier-Domain Method (CDM) for high-resolution computation of time-periodic long-wake flows, with cost-effectives that makes the computations practical. The CDM is closely related to the Multidomain Method, which was introduced 24 years ago, originally intended also for cost-effective computation of long-wake flows and later extended in scope to cover additional classes of flow problems. In the CDM, the computational domain moves in the free-stream direction, with a velocity that preserves the outflow nature of the downstream computational boundary. As the computational domain is moving, the velocity at the inflow plane is extracted from the velocity computed earlier when the plane’s current position was covered by the moving domain. The inflow data needed at an instant is extracted from one or more instants going back in time as many periods. Computing the long-wake flow with a high-resolution moving mesh that has a reasonable length would certainly be far more cost-effective than computing it with a fixed mesh that covers the entire length of the wake. We are also introducing a CDM version where the computational domain moves in a discrete fashion rather than a continuous fashion. To demonstrate how the CDM works, we compute, with the version where the computational domain moves in a continuous fashion, the 2D flow past a circular cylinder at Reynolds number 100. At this Reynolds number, the flow has an easily discernible vortex shedding frequency and widely published lift and drag coefficients and Strouhal number. The wake flow is computed up to 350 diameters downstream of the cylinder, far enough to see the secondary vortex street. The computations are performed with the Space–Time Variational Multiscale method and isogeometric discretization; the basis functions are quadratic NURBS in space and linear in time. The results show the power of the CDM in high-resolution computation of time-periodic long-wake flows.

Effects of counterflow jet on the performance of a generic rocket at Mach 4 to 9

Effects of a jet emanating from the nose of a generic rocket in the opposite direction of the incoming freestream in reducing the drag and the heat loading of the rocket for flight Mach numbers 4 to 9 are reported. Computations to determine the counterflow jet total pressures that would result in most drag reduction are made for 40,000 ft (12,192 m) atmospheric conditions using an ideally expanded Mach 1.5 jet flow as the counterflow jet. In comparison to the baseline case, counterflow jet results in ∼20% reduction in drag at every Mach number studied. The area-weighted average surface temperature is also reduced by 30–70% with increasing flight Mach number.Additional computations were made using the jet total pressure that resulted in most drag reduction for the Mach 6 case to determine the effects of the jet total temperature on the area-weighted average surface temperature. For this purpose, the jet total temperature was varied between 300 and 1000 K. Results indicate that the area-weighted average surface temperature is reduced by 500 K even for jet total temperature of 1000 K.Computations were also made using the jet total pressure that resulted in most drag reduction for the Mach 6 case but for flight Mach numbers ranging from 4 to 9, while jet total temperature was kept as a constant at 500 K. Results indicate that use of a constant jet total pressure value results in 14–22% drag reduction with the peak reduction observed at Mach 6. Results also indicate that area-weighted average surface temperature is decreased by 33–50% with peak reduction observed at Mach 6.

Numerical investigation on streamwise vortex generation by plasma actuator

To clarify vortex formation and development, numerical simulations of a streamwise vortex generation by a dielectric-barrier-discharge plasma actuator were performed. The simulations are set based on the experiment by Jukes and Choi [“Dielectric-barrier-discharge vortex generators: Characterisation and optimisation for flow separation control,” Exp. Fluids 52, 329 (2012)], which is a laminar boundary-layer flow with a plasma actuator aligned to the freestream direction (x). The input momentum by the plasma actuator varies in the low (CaseL), medium (CaseM), and high (CaseH) cases. The streamwise vortex distributions for CaseL and CaseM are similar, and the vortex developments qualitatively agree with the experimental result. In these cases, the single streamwise vortex with the negative x-vorticity is mainly generated by the plasma actuator. For CaseH, however, the generated vortex is made up of both the main streamwise vortex with negative x-vorticity and the helical vortices with positive x-vorticity surrounding the main vortex. These encircling vortices twist the main vortex. Scaling the vortex characteristics, as proposed by Jukes and Choi [“On the formation of streamwise vortices by plasma vortex generators,” J. Fluid Mech. 733, 370 (2013)], demonstrates that the vortex characteristics can be scaled for simulation results. The streamline visualization clarifies the streamwise vortex formation by the induced flow from the plasma actuator and entrainment of the freestream flow to the streamwise vortex. The transient behavior of the streamwise vortex formation is examined for CaseM and CaseH. CaseM and CaseH have distinct transient behavior in the formation of the main streamwise vortex and the surrounding vortices.

Effects of Coanda jet direction on the aerodynamics and flow physics of the swept circulation control wing

Over the last decade, many flight tests of the conventional/canard/flying-wing layout UAVs (unmanned aerial vehicle) have proven that the trailing-edge circulation control (CC) is an effective method to generate a significant rolling moment for the rolling control. However, the Coanda jet’s three-dimensional effects, especially the Coanda jet direction, on the design of UCAVs (unmanned combat air vehicle) have not been revealed and discussed clearly yet. In order to solve this problem, the effects of the Coanda jet direction on the swept wing’s aerodynamics and flow physics are investigated in detail in this paper. An in-house CFD(computational fluid dynamics) solver HyperFLOW with the Spalart–Allmaras for Rotation and Curvature turbulence model and the velocity inlet boundary has been validated by 2-D and 3-D CC cases and is used in this work. The research reveals that compared with the Coanda jet in the freestream direction (FJ jet), the Coanda jet perpendicular to the trailing edge (PJ jet) has a more significant impact on the aerodynamic loads and flow structures for swept wings, including increasing lift, drag, and nose-down moment, enhancing the wingtip vortex and weakening the thickness vortex. In contrast, the FJ jet performs better in improving the lift-to-drag ratio of swept wings. On a typical UCAV, the generalizability of the conclusion about the Coanda jet direction effects on the aerodynamic characteristics is verified. The current research gets an insight into the three-dimensional effects of the trailing-edge CC and provides solid support for implementing the CC technology into practical aircraft concepts.

Experimental and numerical analysis of wake vortex evolution behind transport aircraft with oscillating flaps

The aim of this study is to impose a disturbance on the velocity field already in the near field of the wake vortex system of a generic high-lift transport aircraft configuration by oscillating flaps. This should result in an excitation of long wave instabilities in the downstream development of the wake vortex system. For transient fast-response pressure probe measurements, experiments are carried out in the Göttingen-type low-speed wind tunnel C of the Chair of Aerodynamics and Fluid Mechanics of the Technical University of Munich. The vortex wake is captured at cross flow planes orientated perpendicularly to the freestream direction at nondimensional distances with respect of wing span b of x/b=0.5 and x/b=6. By means of the experimental data at x/b=0.5, a large eddy simulation (LES) is initialized, which solves the downstream development of the wake vortex system. A baseline configuration with statically deflected flaps is compared to a configuration with actuated flaps. It can be shown experimentally as well as numerically that a significant disturbance can be imposed already in the extended near field of the wake vortex system by oscillating flaps. Furthermore, there is a good agreement between the experimental and numerical data. Based on these results, the numerical model can be used for investigations of the farther downstream development of the wake vortex system up to vortex decay in future work.

Investigations of improved cooling effectiveness for ramp film cooling with compound angle film cooling jets

PurposeThis paper aims to investigate the film cooling effectiveness (FCE) and mixing flow characteristics of the flat surface ramp model integrated with a compound angled film cooling jet.Design/methodology/approachThree-dimensional numerical simulation is performed on a flat surface ramp model with Reynolds Averaged Navier-Stokes approach using a finite volume solver. The tested model has a fixed ramp angle of 24° and a ramp width of two times the diameter of the film cooling hole. The coolant air is injected at 30° along the freestream direction. Three different film hole compound angles oriented to freestream direction at 0°, 90° and 180° were investigated for their performance on-ramp film cooling. The tested blowing ratios (BRs) are in the range of 0.9–2.0.FindingsThe film hole oriented at a compound angle of 180° has improved the area-averaged FCE on the ramp test surface by 86.74% at a mid-BR of 1.4% and 318.75% at higher BRs of 2.0. The 180° film hole compound angle has also produced higher local and spanwise averaged FCE on the ramp test surface.Originality/valueAccording to the authors’ knowledge, this study is the first of its kind to investigate the ramp film cooling with a compound angle film cooling hole. The improved ramp model with a 180° film hole compound angle can be effectively applied for the end-wall surfaces of gas turbine film cooling.

Simulations of stably stratified flow past two spheres at Re = 300

Flows past two spheres immersed in a horizontally moving, linearly stratified fluid are investigated at a moderate Reynolds number of 300. Characterization of flow patterns considers representative geometrical configurations defined by varying both the distance between the spheres and their relative orientation to the free stream direction. Simulations are performed on unstructured meshes chosen to accurately resolve the dynamics of fluids in regions close to the spheres for Froude numbers Fr∈[0.25,∞]. Results illustrate the evolution of boundary layers, separation, and the wakes interaction under the influence of a gravity induced buoyancy force. Computations utilize a limited area, nonhydrostatic model employing non-oscillatory forward-in-time integration based on the multidimensional positive definite advection transport algorithm. The model solves the Navier–Stokes equations in the incompressible Boussinnesq limit, suitable for describing a range of mesoscale atmospheric flows. Results demonstrate that stratification progressively dominates the flow patterns as the Froude number decreases and that the interactions between the two spheres' wakes bear a resemblance to atmospheric flows past hills.

Effects of counterflow jet on the performance of a generic rocket

Effects of a jet emanating from the nose of a generic rocket in the opposite direction of the incoming freestream have been studied to determine the effects of the jet in reducing the drag and the heat loads on the rocket. The effects are comparatively evaluated for different jet exit Mach numbers from ideally expanded nozzles, designed using the method of characteristics, operating at a total temperature of 500 K and at total pressures with predetermined ratios of the freestream total pressure. Axisymmetric computations were conducted with the generic rocket flying at a Mach number of 6.0 and 40,000 ft altitude with zero-degree angle of attack. Results indicate that a jet generated by the Mach 1.5 nozzle, operating at a total pressure corresponding to 1/4th of the freestream total pressure, resulted in the best combined overall drag and temperature reduction. The area-averaged surface temperature and drag were reduced by 52%, and 22%, respectively, in comparison to the body without the flow control.

Converging near-elliptic shock waves

Abstract

Bayesian inference of a multivariate model in a phase change situation around a cylinder in staggered arrangement

The complexity of problem of fluid flow and heat transfer over an array of circular cylinders is common in industrial applications of fluid dynamics. The complex nature of the problem encountered in industry gives rise to certain significant dimensions in fluid dynamics theory. Some of them are fluid flow interaction, interferences in flow and vortex dynamics which are typically found in compact heat exchangers, cooling of electronic equipment, nuclear reactor fuel rods, cooling towers, chimney stacks, offshore structures, hot-wire anemometry, and flow control. The mentioned structures are subjected to air or water flows and therefore, experience flow induced forces which can lead to their failure over a long period of time. Basically, with respect to the free stream direction, the configuration of two cylinders can be classified as tandem, side-by-side, and staggered arrangements. The Reynolds Averaged Navier-Stokes (RANS) equations have been used to compute the flow and Eulerian model has been used to understand phase change situation in a staggered arrangement of cylinders. In the present study, nucleate boiling has been the cause of heat and mass transfer between the phases for different cylinder configurations in staggered arrangement. The study was carried out by keeping cylinders stationary as well as rotating to understand the difference and impact of these situations on VOF. The profound effects of arrangement of cylinders, the location of cylinder surface, surface temperature of the cylinder were investigated. To have a deeper and meaningful insight into the phase change phenomenon of water into water vapor, Bayesian inference of a multivariate model involving certain significant factors such as Nusselt number (Nu), Prandtl number (Pr), and Stanton number (St) along with volume fraction of vapor phase of water was carried out. The posterior probabilities computed from Bayesian inference for two statistically significant and experimentally verified datasets were obtained during the study. Through Principal Component Analysis (PCA) for the multivariate datasets used in the study, it was observed that some factors are positively and negatively correlated and individually contributing towards a meaningful finding from the study.

A numerical study of the effect of a single cylinder and plate deflector toward the Savonius wind turbine performance

This study has investigated the effect of the circular cylinder placed at the side of advancing and plate deflector placed in front of returning blade. The CFD approach is done to solve the incompressible URANS. The numerical validation is done by comparing the experimental results using torque coefficient data for Cm=0.185 and TSR=1.078. The turbine using conventional Savonius 1 m of diameter and 1 m of height. A circular cylinder have diameter of 0.7D placed at x/D = 0.5D and y/D = 0.7D. The plate deflector is placed at 1D in front and center of turbine by varying the deflector plate on perpendicular free stream in placed in front of returning at distance variations (y/D) of 0, 0.1, 0.2, 0.3 and 0.4 with deflector angle of 30 degree towards the free stream direction and the Reynolds number of 4.32 x105. The study indicates that the deflector and circular cylinder can increase the turbine Performance at y/D of 0.3 in about 9.71 %.

Time-Accurate Experimental Investigation of Hypersonic Inlet Buzz at Mach 5

Hypersonic inlet buzz is investigated experimentally in a ramjet intake at Mach 5. A two-dimensional planar inlet featuring a double-compression ramp with 10 and 22° inclination to the freestream direction and a rectangular-cross-section duct is tested. A solenoid is used to raise a barrier into the flow and initiate buzz at high throttling ratios. Static pressure is measured at a data sample rate of 10 kHz, and schlieren images are collected at an image frame rate of 30 kHz. Oscillations in the flow not linked to buzz are present at both zero and full throttle, and they cause a high-amplitude pressure peak at very high frequency (higher than 3 kHz). At full throttle, buzz is spotted, in the contemporary presence of shear layers being ingested by the intake (Ferri criterion) and of separated flow lying on the second compression ramp (Dailey criterion). Both features are analyzed by means of the standard deviation of the frames captured. The kind of buzz hereby observed activates high-amplitude pressure oscillations at both low frequencies and high frequencies. The unstarted flowfield is largely subsonic inside the intake and allows for the use of the linear acoustic theory. Indeed, one of the peaks recorded matches the quarter-wave resonator frequency (1411 Hz) of the inlet, suggesting that the barrier acts as a solid wall and that the inlet model behaves as a duct with a closed end.

Adaptive mesh refinement immersed boundary method for simulations of laminar flows past a moving thin elastic structure

One of the critical issues in numerical simulation of fluid-structure interaction problems is inaccuracy of the solutions, especially for flows past a stationary thin elastic structure where large deformations occur. High resolution is required to capture the flow characteristics near the fluid-structure interface to enhance accuracy of the solutions within proximity of the thin deformable body. Hence, in this work, an algorithm is developed to simulate fluid-structure interactions of moving deformable structures with very thin thicknesses. In this algorithm, adaptive mesh refinement (AMR) is integrated with immersed boundary finite element method (IBFEM) with two-stage pressure-velocity corrections. Despite successive interpolation of the flow field by IBM, the governing equations were solved using a fixed structured mesh, which significantly reduces the computational time associated with mesh reconstruction. The cut-cell IBM is used to predict the body forces while FEM is used to predict deformation of the thin elastic structure in order to integrate the motions of the fluid and solid at the interface. AMR is used to discretize the governing equations and obtain solutions that efficiently capture the thin boundary layer at the fluid-solid interface. The AMR-IBFEM algorithm is first verified by comparing the drag coefficient, lift coefficient, and Strouhal number for a benchmark case (laminar flow past a circular cylinder at Re = 100) and the results showed good agreement with those of other researchers. The algorithm is then used to simulate 2-D laminar flows past stationary and moving thin structures positioned perpendicular to the freestream direction. The results also showed good agreement with those obtained from the arbitrary Lagrangian-Eulerian (ALE) algorithm for elastic thin boundaries. It is concluded that the AMR-IBFEM algorithm is capable of predicting the characteristics of laminar flow past an elastic structure with acceptable accuracy (error of ∼0.02%) with only ∼1% of the computational time for simulations with full mesh refinement.

Mechanisms of drag reduction due to flow control around circular disk using coaxial type dielectric barrier discharge plasma actuator at low Reynolds numbers

Mechanisms of drag reduction due to flow control around a circular disk using a coaxial type dielectric barrier discharge plasma actuator (DBD-PA) are numerically investigated at Reynolds numbers (Re) of 5000 and 10 000. The disk is placed normal to the freestream direction and the coaxial type DBD-PA is installed on the upstream side of the disk surface. Flow induced by the coaxial type DBD-PA in quiescent air and a time-averaged drag coefficient of the case without the control are in favorable agreement with experimental results. It is found that two phenomena on flow dynamics are responsible for the drag reduction due to the flow control of the coaxial type DBD-PA. First one is the generation of low-pressure region because of an interaction between the induced flow and the freestream on the upstream side. Second is the delayed roll-up motion of the separated shear layer from the edge of the disk. Most of cases demonstrate the reduction of the time-averaged drag and the reduction rate is up to 2.3% at the Re = 5000 and 3.0% at the Re = 10 000, respectively. Moreover, the control effects on surface pressure distribution at both Reynolds numbers present trade-off and the large relative diameter of the coaxial type DBD-PA is the favorable impacts in terms of the time-averaged drag coefficient and the drag fluctuation.