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DNS for flow separation control around an airfoil by pulsed jets
Abstract
Direct numerical simulation (DNS) for flow separation and transition around a NACA-0012 airfoil with an attack angle of 4° and Reynolds number of 100,000 has been reported in our previous paper. The details of flow separation, formation of the detached shear layer, Kelvin–Helmholtz instability (inviscid shear layer instability) and vortex shedding, interaction of nonlinear waves, breakdown, and re-attachment are obtained and analyzed. The power spectral density of pressure shows the low frequency of vortex shedding caused by the Kelvin–Helmholtz instability still dominates from the leading edge to trailing edge. Based on our understanding on the flow separation mechanism, we try to reveal the mechanism of the flow separation control using blowing jets and then optimize the jets. DNS simulations for flow separation control by blowing jets (pulsed and pitched and skewed jets) are reported and analyzed. The effects of different unsteady blowing jets on the surface at the location just before the separation points are studied. The length of separation bubble is significantly reduced (almost removed) after unsteady blowing technology is applied. The mechanism of early transition caused by the blowing jets was found. A blowing jet with K–H frequency, sharp shape function (very small mass blowing), pitching and skewing obtained the best efficiency based on the increase of the ratio of lift over drag and decrease of blowing mass flow. In this work, a DNS code with high-order accuracy and high-resolution developed by the computational fluid dynamics group at University of Texas at Arlington is applied.
... Separation is alleviated through the mechanism of energizing the flow in the boundary layer, either by directly injecting high-momentum fluid through blowing or by introducing high-momentum fluid from the exterior cross-flow through suction. The pulsed jet [10][11][12][13] is similar to the steady blowing but differs in that the jet is injected into the boundary layer periodically. The boundary layer flow is sufficiently excited that perturbations amplify and transition to turbulence is initiated. ...
... The smoothly curved airfoil leeward surface is modified by the exit of the actuator, denoted by the straight dashed line in Figure 2. As a result, the surface curvature is discontinuous at the junctions between the actuator exit and the airfoil surface. In the present study, the SJ is simply modeled by the time-periodic velocity boundary conditions at the actuator exit, [11][12][13]18,24 instead of a cavity-type actuator where the jet is realized by the piston stroke in the cavity. 17,20,33,38 Although phase lag exists between the piston velocity and the jet velocity, 23 and the jet will undergo viscous dissipation, we mainly focus on the effects of jet on flow separation control and neglect the implementation details of a synthetic jet. ...
... In most DNS and LES studies at relatively higher Reynolds number, the spanwise domain size is normally chosen as L z = 0.1-0.2C. 11,12,38,40 In the DNS study by Hoarau et al. 41 for incompressible flow past a NACA-0012 airfoil at Re C = 800 and 20 • AoA, the dominant flow structure in the spanwise direction is about 0.64C. In our DNS, we use a spanwise domain size of L z = C, and its adequacy is confirmed by a posteriori two-point correlation of the spanwise turbulent fluctuating velocity w ′ , defined as 42 ...
We present results of direct numerical simulations of a synthetic jet (SJ) based separation control of flow past a NACA-0018 (National Advisory Committee for Aeronautics) airfoil, at 10° angle of attack and Reynolds number 104 based on the airfoil chord length C and uniform inflow velocity U
0. The actuator of the SJ is modeled as a spanwise slot on the airfoil leeward surface and is placed just upstream of the leading edge separation position of the uncontrolled flow. The momentum coefficient of the SJ is chosen at a small value 2.13 × 10−4 normalized by that of the inflow. Three forcing frequencies are chosen for the present investigation: the low frequency (LF) F
+ = feC/U
0 = 0.5, the medium frequency (MF) F
+ = 1.0, and the high frequency (HF) F
+ = 4.0. We quantify the effects of forcing frequency for each case on the separation control and related vortex dynamics patterns. The simulations are performed using an energy conservative fourth-order parallel code. Numerical results reveal that the geometric variation introduced by the actuator has negligible effects on the mean flow field and the leading edge separation pattern; thus, the separation control effects are attributed to the SJ. The aerodynamic performances of the airfoil, characterized by lift and lift-to-drag ratio, are improved for all controlled cases, with the F
+ = 1.0 case being the optimal one. The flow in the shear layer close to the actuator is locked to the jet, while in the wake this lock-in is maintained for the MF case but suppressed by the increasing turbulent fluctuations in the LF and HF cases. The vortex evolution downstream of the actuator presents two modes depending on the frequency: the
vortex fragmentation and merging mode in the LF case where the vortex formed due to the SJ breaks up into several vortices and the latter merge as convecting downstream; the discrete vortices mode in the HF case where discrete vortices form and convect downstream without any fragmentation and merging. In the MF case, the vortex dynamics is at a transition state between the two modes. The low frequency actuation has the highest momentum rate during the blowing phase and substantially affects the flow upstream of the actuator and triggers early transition to turbulence. In the LF case, the transverse velocity has a 1%U
0 pulsation at the position 18%C upstream of the actuator.
... The airfoil chord is c = 1. Parameters R, L and d z are selected combining the results of some fast preliminary independence tests, which results are reported in Section 3.1, with available literature concerning high fidelity numerical simulations of 2D airfoils, e.g., [20][21][22][23]. The final choice is reported in Table 1. ...
... Note that in these tests L = 2R is supposed. The comparison with literature data [20][21][22][23] suggests a a shortening, then L = 17 is selected. ...
The present paper proposes an accurate characterization of a NACA 23012 airfoil at near stall conditions at a Reynolds number $Re=3\cdot 10^5$. In light of the unavoidable limits of which experiments suffer near the stall regime both in terms of effective two--dimensionality and data portability across different research groups, the present characterization is performed through high fidelity numerical simulations. Taking advantage of the local discontinuous Galerkin (LDG) LES solver, implemented in the open source library \textit{FEMilaro}, the airfoil behavior is investigated ranging from $\alpha=5^{\circ}$ to $17^{\circ}$, with a peculiar focus for $\alpha \ge 10^{\circ}$. Accuracy of the found results is ensured both from the realistic and accurate reconstructed flow physics and by means of a proper comparison with existing experimental data. Then, a reliable numerical baseline is obtained for any future investigation involving a NACA 23012 airfoil at $Re=3 \cdot 10^5$. The future study of the actual research group will be a parallel BVI analysis, but results are completely extendable as they are to any other research.
... The values chosen for this study are of O(0.1% − 1%), which is of similar magnitude used by previous studies for control over symmetric airfoils. [35][36][37][38] We also seek a coefficient to quantify the rotation input to the flow. Based on the vortical (circulation) input from the actuator, we can quantify the lateral momentum flux as ρr 0 u θ Γ. 27 For the velocity profiles specified, the wall-normal circulation (strength of wall-normal swirl) input is ...
... The results agree with previous works with respect to the momentum coefficient necessary to reattach the flow over a canonical airfoil. [35][36][37][38] Decreasing the coefficient of momentum to C µ → 0 (6G and 6H) we find that we can affect the flow without any momentum injection if we have C swirl > 0. The pure rotational cases (6G and 6H) on the far left of Fig. 7, show the effect on lift and drag. Figure 6 shows these two different cases (6G and 6H) with no blowing, and one with four times the swirl coefficient of the other (C swirl = 2.1% and 8.4%). ...
We numerically investigate the inuence of momentum and wall-normal vorticity injection in separation control of flow over a spanwise periodic NACA 0012 airfoil at a Reynolds number of 23,000. Large-eddy simulations are performed at angles of attack of α = 6°and 9° with flow control input prescribed through the wall boundary conditions near the natural separation point. It is shown that at α = 6°, a relatively small amount of momentum injection is needed to eliminate the separated region. At the higher angle of attack of α = 9°, it is found that stall can be suppressed with a combination of momentum and vorticity injection. The addition of vorticity enhances mixing in the spanwise direction and redirects the high momentum uid closer to the suction surface of the airfoil. For particular combination of momentum and vorticity injections we demonstrate that drag reduction and lift enhancement can be obtained.
... The sizes of the computational domain for the transitional flow are Fig. 1 Schematic of the physical model and two-dimensional grid, with the green lines denoting the surfaces of the hydrofoil. The grid is plotted in every 16 gridlines in each direction for clarity determined by experience and based on similar cases of flow around an airfoil or hydrofoil Deng et al. 2007;You and Moin 2008;Gross and Fasel 2010;Zaki et al. 2010;Catalano and Tognaccini 2011). ...
The laminar flow on a curved surface transits to turbulent induced by streamline curvature which generates pressure gradient field and separated shear layer flow. We performed a direct numerical simulation investigation on transitional flow through a linear cascade consist of S-shaped S3525 hydrofoil which has different curvature variations on the two surfaces, i.e., concave-to-convex and convex- to-concave in the streamwise direction. The objectives are to quantitatively assess the effects of streamline curvature of the hydrofoil surface on the three-dimensionality of the separated and transitional flow, including the patterns of separation and reattachment, formation and development of three-dimensional boundary layer flow, and statistics on non-homogeneous turbulent near-wall flow. Comparisons between the near-wall flows of the two surfaces demonstrate the effect of streamline curvature and its associated influential mechanisms such as pressure gradient field. Numerical data reveal that transition and occurrence of three-dimensional flow are observed earlier for the concave-to-convex surface; intermittent flow is generated in the concave section near the leading edge and convex section near the trailing edge where three-dimensionality of flow and turbulent fluctuations are the most pronounced. However, the boundary layer and near-wall flow for the convex-to-concave surface is quite stable until the concave section, thus three-dimensionality of separation and reattachment, boundary layer flow and turbulent behaviors are only notable near the trailing edge.
... Perturbations due to upstream propagating acoustic waves have been found in a laminar boundary layer before separation [18]. For a flat plate in a low free stream turbulence flow, Schubauer & Skramstad [19] found velocity fluctuations in the laminar boundary layer. ...
Airfoil self-noise at high Reynolds numbers is caused by turbulent boundary layers. In modelling such a case with Large Eddy Simulation (LES) or Direct Numerical Simulation (DNS), the acoustic source and thus surface pressure fluctuations are incoherent along the modelled spanwise length, assuming the simulated span is wide enough to capture turbulent flow structures. In literature, the spanwise coherence of surface pressure is normally inspected at 0.95% of the chord. A sound pressure level (SPL) correction for an increase in spanwise length can be applied, for example, for comparison of the modelled SPL spectra with experimental measurements. For a larger spanwise surface pressure coherence length, the SPL correction increases. In contrast, airfoils at low Reynolds numbers emit self-noise with a tonal signature. An acoustic feedback mechanism propagates instabilities upstream and is responsible for the tonal amplification. Consequently, the source of noise is not highly localized and may not be fully incoherent at all frequencies. Therefore, the chordwise position at which the surface pressure coherence length is to be calculated for a SPL correction for spanwise length, is not straightforward. In this paper, the sources of tonal acoustic noise and the effect of the position of spanwise coherence length calculation for SPL correction is investigated. A NACA 0012 airfoil is modelled using LES, at a chord-based Reynolds number of Rec = 1.1×10⁵ and an angle of attack of α = 3°. The far field acoustic pressure is calculated with the Ffowcs-Williams Hawkings acoustic analogy Formulation 1C. The airfoil source regions are separated by their boundary layer flow characteristics and their relative contributions to the far field SPL are analysed to find a chordwise point on the airfoil surface which has a spanwise coherence length that is representative of the nature of the far field acoustic radiation. The spanwise coherence length at this position can be used to calculate the SPL correction for an increase in spanwise length. Two chordwise points are identified for which the resulting corrected SPL spectra agrees well overall with experimental data.
... Table 1 summarizes domain sizes used in selected airfoil simulations in literature. The domain radius of these airfoil DNS ranges from 4c [53] to 100c [24,39]. In the present study, two different C-type meshes are applied, where each grid has a sharp trailing edge and a domain radius, R, and wake length, W , of 30c. ...
A comprehensive and detailed overview of the flow topology over a cambered NACA 65(1)-412 airfoil at Re = 20,000 is presented for angles of attack ranging from 0{\deg} to 10{\deg} using high-order direct numerical simulations. It is shown that instabilities bifurcate the flow and cause it to change at a critical angle of attack from laminar separation without reattachment over a laminar separation bubble at the trailing edge to a bubble at the leading edge. The transition of the flow regimes is governed by the Karman vortex shedding of the pressure side boundary layer at the trailing edge, Kelvin-Helmholtz instabilities within the separated shear layer on the suction side, as well as three-dimensional instabilities of elliptic flow within the vortex cores and hyperbolic flow in the shear layer between subsequent Karman vortices. As the suction side shear layer transitions and reattaches, the interaction of the two and three-dimensional instabilities results in three-dimensional tubular structures and large-scale turbulent puffs. The formation and shifting of the laminar separation bubble defines the far-wake topology several chord-lengths behind the airfoil and is accompanied by a sudden increase of the lift force and decrease in the drag that underscores the sensitive nature of low-Reynolds number airfoil aerodynamics. Lift and drag polars are presented for direct numerical simulations, wind tunnel experiments, and simplified numerical procedures where incorrect prediction of the force coefficients is caused by the failure to correctly model the low-pressure region at the trailing edge that is caused by the time-dependent generation of the Karman vortices.
... 81,82 Furthermore, the global instability involving the feedback loop may lead to a transition to turbulence without increasing the disturbance explicitly. For example, Deng et al. 83 analyzed the pressure fluctuations at different locations on the suction surface of the NACA-0012 airfoil through a direct numerical simulation (DNS), as shown in Fig. 10. Although no external disturbances were introduced, the initial perturbations upstream may have originated from upward traveling acoustic waves that were generated in the wake. ...
This Review summarizes the progress in research on the flow structure and aerodynamic characteristics of an airfoil at a low Reynolds number encountered by near-space low-speed aircrafts and micro-air vehicles. The structures of several kinds of laminar separation bubbles and their effect are discussed by drawing on experimental and numerical results reported in the past few decades. The transition process in the separation bubble is detailed from various perspectives, including the receptive, primary instability, secondary instability, and break-down stage. The process of evolution of a coherent vortex structure that may affect the transition is discussed by analyzing the vortex dynamics in the separation bubble. Combined with the flow characteristics at a low Reynolds number and data on the airfoil, aerodynamic characteristics of the airfoil, such as the nonlinear effect and static hysteresis effect at a low angle of attack, are discussed.
... From previous result shows that with increasing angle of attack (AOA) lift force is going to increase and dreg force is going to decrease at certain AOA, after that lift force drastically going to decrease and dreg force is going to increase, so it's call stall condition . [11][12][13][14][15][16][17][18][19][20][21][22][23]. Out of which by creating slot at leading edge to energized upper surface of airfoil by using high pressurized air from lower surface is effective. ...
In this paper mainly focused effect of fixed slot on performance of NACA0012 airfoil. Using slot lower surface high pressurized air passes through slot to energized upper surface. Analysis has been done on NACA0012 of 1m cord length at 25C of air with 5m/s air velocity. First plain airfoil at different angle of attack has been analyzed to find out stall condition and C l , C d. After that same parameter and physics, only change in geometry of airfoil with 15%C leading edge slot. Flow separation is adverse effect for performance of airfoil. Due to flow separation adverse pressure gradient effect reverse flow is there, we cause reduction of lift coefficient and increment of drag coefficient so flow separation necessary to reduce. There are different techniques to reduce flow separation but effect of slot on airfoil has been studied.
... Figure 4.20 below. This is in agreement to (Deng 2007), who investigated flow separation control around an airfoil using pulsed jets. The flow separated at high angles of attack. ...
This experiment was conducted to determine the effects of angle of attack on lift, drag, pitching moment and pressure distribution during the flow of NACA 4415 wing model of chord length of 200mm, leading edge radius of 4.96mm and span of length 450 mm using a low speed open return wind tunnel with test section of 0.47m width, 0.47m height and 1.27m length. The wing model was subjected to several tests at two different speeds of 15m/s and 25m/s which corresponds to two different Reynolds numbers of 1.8×10 5 and 3.1×10 5 respectively. The result showed that for both speeds, the coefficient of lift increased as angle of attack increased and the maximum lift coefficient was 1.45 at 15m/s while at 25m/s, the coefficient of lift is 1.40, which is slightly less than that at 15m/s. The coefficient of drag also increased as the angle of attack increased from 0 0 to 18 0 and the maximum drag was 0.15 at 15m/s, while at 25m/s, the coefficient of drag is approximately 0.2 which showed that the wing has a very low drag coefficient. For pressure distribution, it was observed that there was no flow separation at low angle of attacks from-6˚to6˚to 0˚, but the flow began to separate at moderate angle of attack of 3 0 to 9 0 and fully separated at high angle of attacks of 12 0 to 18 0 with vortices for both speeds.
... Many active device types were then developed for which, depending on the control source used, the control is of different nature (plasma actuators, fluidic actuator. etc.) [9][10][11][12]. However, it is clear that many studies for active control require complex solutions to be implemented, in terms of control, motorization or regarding the kinematics of the devices. ...
Vortex generators are used extensively as a passive flow control devices to delay or remove the boundary layer separation, which affects the hydrodynamic performance of the hydrofoil. In this paper, a new approach is introduced to overcome the boundary layer separation on the NACA S1210 hydrofoil. The outcome of tube slots combination in the S1210 hydrofoil on the boundary layer separation are numerically investigated. The performance is compared with respect to the force coefficients and glide ratio. The effects of tube slot inlet positions with different diameters on S1210 hydrofoil are presented here. The results show that the smaller diameter tube slots starting near the leading edge improves the hydrodynamics performance of the hydrofoil.
... The numerical methods were also proven to be capable of capturing weak instabilities that might not occur using simulation techniques with artificial viscosity or damping. Deng et al. (2007) performed further DNS using the same comparably coarse grid as Shan et al. (2005) in order to analyse active flow control using pulsed jets. Shan et al. (2008) performed DNS of the NACA 0012 geometry at 6 • incidence. ...
The performance of turbomachinery components and the safe flight envelope of next-generation aircraft is often limited by complex transonic flow phenomena. Since the first flights close to sonic speeds, experiments have been carried out to explore the origin of flutter phenomena, supplemented with simulations of the Reynolds-averaged Navier-Stokes equations that are dependent on turbulence models. To date, direct numerical studies of low-frequency phenomena have been limited to low Reynolds numbers. The present work explores the transonic flow regime around Dassault Aviation’s V2C laminar-flow profile at moderate Reynolds numbers, and also analyses boundary-layer instabilities on a high-pressure turbine vane. Direct numerical simulations of an un-swept wing section are carried out at Mach 0.7 and an angle of attack of 4◦ using the in-house code SBLI, which is a well-validated high-order finite-difference flow solver. While the flow at Reynolds numbers of Re = 200,000 is purely subsonic, a significant supersonic region is observed for Re ≥ 500,000. Whereas experimental investigations of the same airfoil at higher Reynolds numbers showed single shock waves, the present reference case at Re = 500,000 exhibits continuously upstream-propagating shock waves. Besides laminar/turbulent boundary-layer transition and acoustic phenomena, a low-frequency phenomenon, known as transonic buffet, is studied. In addition to spectral analyses of the flow, linear stability analysis and a dynamic mode decomposition method are used to study flow phenomena apparent at different frequency ranges between Strouhal numbers of St ≈ 20 (Kelvin-Helmholtz instabilities) and St = 0.12 (transonic buffet). Resolution of small-scale structures is established by a grid convergence study and also employing spectral error indicators to assess the grid quality. The flow characteristics are also confirmed by a simulation with a five times wider spanwise domain size (25% of the chord length) comprising more than five billion grid points. A key observation of this work is the clear distinction between an acoustic mechanism associated with the shock-wave motion (St ≈ 0.5) and a quasi-periodic mode at significantly lower frequencies causing strong fluctuations in the aerodynamic lift (St ≈ 0.12). The Strouhal number of the low-frequency phenomenon agrees well with the buffet frequency in experiments at higher Reynolds numbers.
... Pulsed blowing is also a way of active boundary layer control, the effect of which on control performance was studied by Hecklau [13] and Deng [14]. In most cases excitation is incorporated at the leading edge to affect the boundary layer upstream of the point of separation, with suction and blowing (steady or periodic) [15]. ...
... The modifications of lift and drag forces with flow control for different levels of wall-normal momentum input are summarized in Fig. 7 over the range of 0% ≤ C µ ≤ 2.1%. This range is selected to follow the previous studies achieving effective flow control over symmetric airfoils [27,38,58,59]. Shown by the color of the symbols in this figure is the magnitude of angular velocity input u θ,max /U ∞ . ...
The objective of this computational study is to quantify the influence of wall-normal and angular momentum injections in suppressing laminar flow separation over a canonical airfoil. Open-loop control of fully separated, incompressible flow over a NACA 0012 airfoil at $\alpha = 9^\circ$ and $Re = 23,000$ is examined with large-eddy simulations. This study independently introduces wall-normal momentum and angular momentum into the separated flow using swirling jets through model boundary conditions. The response of the flow field and the surface vorticity fluxes to various combinations of actuation inputs are examined in detail. It is observed that the addition of angular momentum input to wall-normal momentum injection enhances the suppression of flow separation. Lift enhancement and suppression of separation with the wall-normal and angular momentum inputs are characterized by modifying the standard definition of the coefficient of momentum. The effect of angular momentum is incorporated into the modified coefficient of momentum by introducing a characteristic swirling jet velocity based on the non-dimensional swirl number. With this single modified coefficient of momentum, we are able to categorize each controlled flow into separated, transitional, and attached flows.
... They found that turbulence intensity can alter the aerodynamic performance and cause more diffusion in the wingtip vortex core. The flow field structure was investigated by Deng et al. [11] via DNS simulation to evaluate the effects of pulse jet. Applying a pulsative jet led to a reduction in separation zone as well as the drag coefficient. ...
Evolution of wing tip vortex has been widely studied by many researchers. Winglet, jet, and suction devices have been implemented close to the wing tip to passively or actively mitigate the tip vortex effects. This study aims to investigate the effects of employing a synthetic jet at the tip of a half wing model. The flow was assumed to be incompressible, low speed and the Reynolds number based on chord length with aspect ratio equal to two for half wing was 1.8 × 10 5. Different reduced frequencies and momentum coefficients were applied. A Detached Eddy Simulation by considering Spalart-Allmaras as the turbulence model for subgrid scale zone and near the walls was employed to simulate the flow field study. Results showed large diffusivity in vortex core. Also, a reduction in longitudinal and total velocity magnitude has observed at vortex core region in the near wake.
... Pulsed blowing was introduced to overcome this. Studies conducted by Hecklau et al. [11] and Deng et al. [12] shows that pulsed blowing will not reduce the control performance. Huang et al. [13] conducted elaborate studies to analyse and determine the physical mechanism that govern the suction and blowing flow control. ...
Boundary layer separation over an airfoil causes large energy losses and strong adverse pressure gradients. This in turn leads to a reduction in the lift force and an increase in the drag force. Therefore delaying or if possible, eliminating the flow separation is mandatory. The elimination of flow separation would permit higher angles of attack for many practical applications. Steady blowing on the suction side of the airfoil is found to be effective in controlling the boundary layer separation. Flow around NACA0012 and LA203A airfoils are analysed in the present study, with the position of the secondary blowing jet at 60 percent of the chord length and angles of attack ranging from 2 to 18 degree for NACA0012 and 2 to 20 degree for LA203A. The secondary blowing velocity is varied from 0 percent to 40 percent of the free stream velocity. The lift curves of all the cases studied are plotted. The results show that the secondary blowing helps to control flow separation and cause an increase in the lift and delay the stalling of airfoils in both cases.
... Passive devices were replaced by active ones, which can be turned off when not necessary in order to avoid additional drag (i.e. during cruise flight of an aircraft for instance). Many active device types were then developed for which, depending on the control source used, the control is of different nature (plasma actuators, fluidic actuators, etc.) [9][10][11][12]. However, it is clear that many studies for active control require complex solutions to be implemented, often starting from existing technological bricks, in terms of control, motorization or regarding the kinematics of the devices. ...
The present paper provides an experimental optimization of a NACA 4415 airfoil equipped with vortex generators (VGs) to control its flow separation. To build this optimal configuration an experimental parametric study was conducted on five geometrical parameters: thickness and height of vortex generators, position, orientation angle with respect to the mean flow direction, spacing in the spanwise direction. Moreover, a new configuration that includes micro generators behind the conventional ones was also investigated as a potentially interesting solution. For all these cases wind tunnel tests were performed and compared for different angles of attack and various Reynolds numbers up to 2 105. These experiments enabled us to highlight the main trends to get an optimal design, for which quantitative improvement can be achieved by passive means in terms of aerodynamic performances on NACA4415 airfoil. The results reveal that triangular shape vortex generators are best suited to control boundary layer separation. An optimum angle of VGs is obtained for 12°with a 3 mm distance between vortex generators located at 50% of the chord. It was found that micro vortex generators are very effective in controlling the flow with less parasite drag. The maximum lift coefficient for an airfoil with coupled vortex generators increases by 21% and a flow separation is delayed by 17°. However, this very good performance is counterbalanced by the appearance of parasitic drag. Indeed, it creates a counter-rotating array of vortices with the second raw of micro-vortex generators that reinforce the vortexes strength without any increase in device height.
... An examination of the literature, especially the DNS and LES studies, reveals that the aspect ratio of the airfoil has profound effect on both the time-averaged and time-dependent behavior of the global quantities and turbulent statistics. Although it is shown in Table 2 that a smaller aspect ratio around 0.2 is often considered acceptable for the low AoA airfoil [24,28,[42][43][44][47][48][49] with separatedreattached flow, a systematic investigation of the aspect ratio effect is still lacking. Low AoA transition and reattachment patterns of the turbulent flow are complex and indeed quite different from those of the massively separated flow at high AoA. ...
... Dans cette configuration, la couche limite transitionne à cause de la présence d'une bulle de recirculation sur l'extrados du profil. Liu[118] et Deng et al.[43] ont aussi étudié la même configuration avec le même code que Shan et al.[154]. Cependant, ils se sont focalisés sur l'étude de la transition de la couche limite sur le profil (K-type ou N-type[96]). ...
The study of the mechanisms of noise generation by the flow over an airfoil is essential to reduce the airframe noise. A direct aeroacoustic solver has been developed to shed some light on these mechanisms. Low dissipation and low dispersion numerical schemes are designed to preserve the weak acoustic waves. A multi-size mesh multi-time step algorithm has been developed to realize local grid refinements on a structured mesh and to reduce the calculation cost of the direct noise computation. The present solver is used to compute the noise generated by a 2-D NACA0012 airfoil at a low Reynolds number and to study the effect of the angle of attack on both flow and acoustic fields. Furthermore, the presence of the tonal noise from a NACA0012 airfoil at a moderate Reynolds number of 200 000 is investigated. A numerical insight into the effect of the experimental confinement due to the wind tunnel walls is then given for a 3-D NACA0018 airfoil at Reynolds 160 000. Finally, a direct computation of the flow over a 3-D truncated NACA0012 airfoil at a high Reynolds number (2.32 millions) is performed by large eddy simulation. The solution is compared with an experimental database named EXAVAC. The main noise generation echanisms are well reproduced with the multi-domain approach.
... They found that turbulence intensity can alter the aerodynamic performance and cause more diffusion in the wingtip vortex core. The flow field structure was investigated by Deng et al. [11] via DNS simulation to evaluate the effects of pulse jet. Applying a pulsative jet led to a reduction in separation zone as well as the drag coefficient. ...
An important phenomenon in three-dimensional flow over a wing is the existence of wingtip vortex. It has significant effects on the aerodynamics of flying vehicles. In this computational study, we investigate the effects of geometry of the wingtip on the structure of the wing-tip vortices. Here, we consider a rectangular half-wing with NACA0012 airfoil as cross section. The aerodynamic coefficients and the flow-field variables are computed at low Reynolds numbers below 50,000. As the edge-shape parameter is increased the wing tip vortex is weakened. This influence is higher at higher values of Reynolds number. But, the increase of angle of attack does not change the shape or rate of this increase.
... Unfortunately, three-dimensional (3D) flow control surveys are severely limited. Deng et al. [20] examined blowing flow control via the DNS method to optimize the blowing jets. They studied the effects of different unsteady blowing jets on the surface at locations just before the separation points, and the separation bubble length was significantly reduced after unsteady blowing was applied. ...
A three-dimensional study of suction flow control has been performed to investigate the aerodynamic characteristics of a rectangular wing with NACA 0012 airfoil section, and also the optimum length of suction jet has been determined. In this study, the RANS equations were employed in conjunction with a k-ω SST turbulent model. Perpendicular suction at leading edge was applied on the upper surface of the wing, with two different types of slot distribution, center suction and tip suction. The suction jet lengths were varied from 0.25 to 2 of chord length and the jet velocity was also selected 0.5 of freestream velocity. Most importantly, the results indicated that the lift to drag ratio increase as the suction jet length rises in both cases. However, the improvement of aerodynamic characteristics is more significant for center suction, and it is extremely close to the entire wing suction situation, which the jet length is equal to the wingspan. Moreover, in the center suction case the vortexes are frequently abated or moved downstream and interestingly more vortexes are removed in comparison to tip suction, in similar conditions. Ultimately, when the jet length is less than of a half the wingspan, tip suction, and while the jet length is greater than of a half the wingspan, center suction is more suitable.
... The authors presented three-dimensional undulated large-scale vortices row with a regular spanwise wavelength which is very similar to that of bluff body wakes. Deng [4] conducted direct numerical simulations for flow separation and transition around a NACA0012 airfoil with an attack angle of 4 • and a Reynolds number of 100,000 and the details of flow separation, formation of detached shear layer, Kelvin-Helmholtz instability, vortex shedding, interaction of non-linear waves, breakdown and reattachment investigated. ...
Low Reynolds number aerodynamics is important for various applications including micro-aerial vehicles, sailplanes, leading edge control devices, high-altitude unmanned vehicles, wind turbines and propellers. These flows are generally characterized by the presence of laminar separation bubbles. These bubbles are generally unsteady and have a significant effect on the overall resulting aerodynamic forces. In this study, the time-dependent unsteady calculations of low Reynolds number flows are carried out over an Eppler 387 airfoil in both two-and three-dimensions. Various instantaneous and time-averaged aerodynamic parameters including pressure, lift and drag coefficients are calculated in each case and compared with the available experimental data. An observed anomaly in the pressure coefficient around the location of the separation bubble in two-dimensional simulations is attributed to the lack of spanwise flow due to three-dimensional instabilities.
... For a reduced frequency of 4.29 a gain in lift over drag of 325% relative to the uncontrolled flow was demonstrated for 8.3deg angle of attack. Deng et al. [15] investigated AFC using pulsed jets for a NACA-0012 airfoil. They observed that for large amplitudes AFC lead to bypass transition while for small amplitudes unstable modes of the flow were excited and amplified by the flow leading to transition. ...
The flow over a NREL S822 wind turbine airfoil was simu-lated for a chord Reynolds number of 100,000 and an angle of attack of 5deg. These conditions approximately match the blade element conditions at 80% radius of a 2m diameter turbine op-erating at 300rpm. A simulation of the uncontrolled flow with steady approach flow conditions shows boundary layer separa-tion on the suction side which is consistent with University of Illinois at Urbana-Champaign experimental data. Active flow control has the potential to locally (and on demand) reduce the unsteady loads on individual turbine blades during non-nominal operation, thereby increasing turbine life. In addition, flow con-trol may help lower the cut-in wind speed. Unsteady flow control for reducing the suction side separation using pulsed vortex gen-erator jets, flip-flop jets, and plasma actuators were evaluated. It was found that very low actuation amplitudes were already suffi-cient for eliminating the suction side separation. The high effec-tiveness and efficiency is traced back to hydrodynamic instabili-ties that lead to a downstream growth of the forced disturbances. Too high actuator amplitudes resulted in early disturbance satu-ration which made the control inefficient. NOMENCLATURE B jet blowing ratio c chord length, wave speed c d drag coefficient c l lift coefficient c f skin friction coefficient c µ momentum coefficient f frequency k spanwise mode number Q vortex identification criterion s arc-length measured from leading edge S i j strain rate tensor t time T period u, v velocity components v ∞ approach flow velocity W i j shear stress tensor α angle of attack, pitch angle, streamwise wave number β skew angle, spanwise wave number τ duty cycle ρ density ω z spanwise vorticity INTRODUCTION Modern large horizontal axis wind turbine (HAWT) rotors are subjected to significant unsteady loads resulting from wind shear, gusts, large scale atmospheric turbulence, yaw misalign-ment, upwind wakes, and aeroelastic vibration of the blades among others. Wind gusts (sudden change in wind speed and/or direction) can quickly drive the turbine operating state out of the nominal region and into rotational augmentation or dynamic stall which are characterized by large unsteady aerodynamic loads and high acoustic noise. Dynamic stall describes the complex series of events that results in the dynamic delay of stall on air-foils and wings experiencing unsteady motion with angles of at-tack significantly beyond those of static stall. In the presence of turbulence, the aerodynamic loads associated with dynamic stall generate the highest fatigue cycles for both, blade and rotor shaft bending [1]. Other phenomena such as the tower blade passing, wind shear, shaft tilt, wakes from upwind turbines, etc. contribute to the unsteady loads. Larger turbines with higher aspect ratio blades have a better power conversion efficiency but are also ex-posed to larger structural and fatigue loads. Implementing new and innovative flow control techniques could decrease excessive loads, which affect the rotor and accompanying components. A reduction or control of the unsteady aerodynamic loads would allow for a lighter design and lengthen the turbine life and thus make wind energy more competitive.
... The large-eddy simulation for this flow demonstrated that synthetic-jet actuation could effectively delay the onset of flow separation and cause a significant increase in the lift coefficient. Shan et al. [14,15] and Deng et al. [16] used the Direct Numerical Simulation (DNS) method to study the flow separation and control of the separation over a NACA 0012 aerofoil using vortex generators and pulsed and blowing jets. Segawa et al. [17] studied the performance of an alternating suction/blowing array of jet actuators for a range of Reynolds numbers from 3.8 × 10 5 to 5.7 × 10 5 , and they showed that clockwise and counterclockwise longitudinal vortices appeared to be formed side-by-side in the near wall region. ...
Transition-sensitive, single point eddy viscosity models are fairly new and performance assessment of these models is required. In this study, the performance of transition and turbulence models is first evaluated for predicting low Re number flows having laminar separation bubbles that are traditionally quite difficult to predict for RANS-based CFD methods. Second, suppression of the laminar separation bubbles using blowing and/or suction is investigated for a single aerofoil. A prior experimental study using hot-wire anemometry for a clean (no jet) NACA 2415 aerofoil at α=8° at a Reynolds number of 2×105 shows the presence of a laminar separation bubble. For this flow, the recently developed k–kL–ω transition model is first shown to accurately predict the location and extent of the experimentally measured separation bubble. Following this, the same transition model was used to predict the flow over the NACA 2415 aerofoil using single or simultaneous blowing or suction. In the single blowing or suction cases, the separation bubble is not completely eliminated, but either abated or moved downstream. Smaller blowing velocity ratios cause more effective suppression of the separation bubble than larger blowing ratios, independent of the blowing locations. In contrast, larger suction velocity ratios are better than smaller suction ratios for the suppression. Moreover, the lift/drag ratios increase as the jet location moves from the leading edge to a downstream direction in both cases. In the simultaneous blowing and suction cases, the k–kL–ω transition model is shown to suppress the separation bubble by using a mix of jet parameters which result in increased lift/drag ratios.
The separated flow at low Reynolds number around the NACA0012 airfoil is numerically studied by large-eddy simulation. Strategies of boundary layer suction to control flow separation are investigated. A method of using two-zone suctions, near the leading edge and near the trailing edge, are calculated. Based on verification with direct numerical simulation (DNS) and experimental data, the results of the lift and the drag, the vortices, and the strength of near-field pressure fluctuations, are checked. The results show that the two-zone suctions can supress flow transition and separation, thereby increase the lift and reduce the drag. The shedding of vortices is weakened, and the near-field pressure fluctuations are attenuated. For comparison with the two-zone suctions, the strategies of suction near the leading edge only and suction near the trailing edge only are also studied. It is found that suction near the leading edge only may suppress transition and delay separation when the suction zone is large enough, but the flow property deteriorates due to shedding vortices in the wake. The suction near the trailing edge only may improve the flow performance by reducing the size of the vortices in the rear section of the airfoil and in the wake region, but it has little effect on the separation bubble and transition.
To take of aircraft within short distance and to increase efficiency of turbo-machinery lift force required to increase and drag required to reduce. These mainly depend on the lift coefficient and drag coefficient. When flow pass over airfoil at some angle due to the adverse pressure effect flow is going to separate from surface so it reduce the lift and increase drag. So in this paper mainly focused on the different technique to reduce flow separation and some general idea about different modal of CFD.
In this chapter we present considerations on the impact of the turbulence characteristics of most flows of industrial interest on the flow control interpretations or strategies. In the second part we will give an overview of the CFD approaches, including actuator integration methods. The actuators characteristics will then be presented with the associated detection methods. Considerations on the spatial and time characteristics of flow control will be provided.
The modified NACA 643-618 airfoil of the Aeromot 200S Super Ximango motorglider was chosen for wing section simulations at Re = 300;000 and a = 15 deg. A coarse-grid direct numerical simulation was carried out to obtain detailed reference data. A laminar separation bubble close to the leading edge causes transition to turbulence of the suction-side boundary layer. The turbulent boundary layer separates near midchord. The considerable computational expense of this simulation provided the motivation for hybrid simulations based on the k-? model and a one-equation renormalization group model. An extension to the turbulence models that "seeds" synthetic turbulence velocity fluctuations in the parts of the flowfield where the grid resolution changes is proposed but was found to have little effect on the mean flow. When the model contribution was too low, the turbulent boundary layer on the airfoil suction surface separated too early. By increasing the model contribution, a reasonable agreement with the reference data was obtained. © 2015 by the authors. Published by the American Institute of Aeronautics and Astronautics, Inc..
An introduction to CFD fundamentals and using commercial CFD software to solve engineering problems, designed for the wide variety of engineering students new to CFD, and for practicing engineers learning CFD for the first time. Combining an appropriate level of mathematical background, worked examples, computer screen shots, and step by step processes, this book walks the reader through modeling and computing, as well as interpreting CFD results. The first book in the field aimed at CFD users rather than developers. New to this edition: A more comprehensive coverage of CFD techniques including discretisation via finite element and spectral element as well as finite difference and finite volume methods and multigrid method. Coverage of different approaches to CFD grid generation in order to closely match how CFD meshing is being used in industry. Additional coverage of high-pressure fluid dynamics and meshless approach to provide a broader overview of the application areas where CFD can be used. 20% new content.
Control of flow over a NACA2415 aerofoil which experiences a laminar separation bubble for a transitional Reynolds number of 2×10 5 is numerically simulated under the effects of blowing and suction. An earlier experimental study using hot-wire anemometry for a clean (no jet) NACA 2415 aerofoil at α = 8° shows a laminar separation bubble over the one-third of the airfoil upper surface. In the no jet case, the recently developed k-k L-ω and k-ω SST transition models accurately predict the location and extent of the separation bubble. Later, single or multiple jets with a width of 2.5% the chord length are placed on the aerofoil upper surface for simulating the isolated or simultaneous blowing and suction jets. For single jets with blowing or suction, whereas the k-ω SST model suppress the separation bubble in both cases, the k-k L-ω transition model doesn't completely eliminate the separation bubble but moves it downstream in the suction case. Overall, the blowing/suction control mechanism appears to be suppression of the separation bubble and reduction of the upper surface pressure to increase the lift and decrease the drag. For simultaneous blowing and suction, firstly, a blowing jet at 10%c and a suction jet at 36%c are placed at the beginning and end of the separation bubble based on the best results of the single jets. Then, the locations of the blowing and suction jets are reversed. In both cases, both transition models eliminate the separation bubble resulting in increase in the lift and decrease in the drag. However, the best results in terms of the L/D ratios are still obtained with the single suction jets. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.
The flow over a NREL S822 wind turbine airfoil was simulated for a chord Reynolds number of 100,000 and an angle of attack of 5deg. These conditions approximately match the blade element conditions at 80% radius of a 2m diameter turbine operating at 300rpm. A simulation of the uncontrolled flow with steady approach flow conditions shows boundary layer separation on the suction side which is consistent with University of Illinois at Urbana-Champaign experimental data. Active flow control has the potential to locally (and on demand) reduce the unsteady loads on individual turbine blades during non-nominal operation, thereby increasing turbine life. In addition, flow control may help lower the cut-in wind speed. Unsteady flow control for reducing the suction side separation using pulsed vortex generator jets, flip-flop jets, and plasma actuators were evaluated. It was found that very low actuation amplitudes were already sufficient for eliminating the suction side separation. The high effectiveness and efficiency is traced back to hydrodynamic instabilities that lead to a downstream growth of the forced disturbances. Too high actuator amplitudes resulted in early disturbance saturation which made the control inefficient.
The flow over a National Renewable Energy Laboratory S822 wind turbine airfoil was simulated for a chord Reynolds number 01 100,000 and an angle of attack of 5 deg. These conditions approximately match the blade element conditions at 80% radius of a 2-m-diameter turbine operating at 300 rpm. Simulations of the uncontrolled flow show boundary-layer separation on the suction side, which is consistent with University of Illinois at Urbana Champaign experimental data. Active flow control has the potential to locally (and on demand) reduce the unsteady loads on individual turbine blades during nonnominal operation, thereby increasing turbine life. In addition, flow control may help lower the cut-in wind speed. Unsteady flow control for reducing the suction side separation using pulsed vortex generator jets, flip-flop jets, and plasma actuators were evaluated. Actuation frequencies of 2.5 and 5 (normalized with freestream velocity and axial chord length) and blowing velocities of 0.01, 0.1, and 1 (normalized with the freestream velocity) were investigated. It was found that blowing ratios as low as 0.1 were already sufficient for eliminating the suction side separation, resulting in a more than four times increase of the lift-to-drag ratio. The high effectiveness and efficiency is traced back to hydrodynamic instabilities that lead to a downstream growth of the forced disturbances. Too high actuator amplitudes resulted in early disturbance saturation, which made the control inefficient.
Flow induced aerodynamic̈ instability in transonic flow field is initiated by the shock-boundary layer interaction which is frequent in turbornachinery. In recent years, the effect of bump wall on the flow field around an airfoil has been investigated experimentally and as a result, it was observed that the bump wall is effective for the control of shock wave on the airfoil. In the transonic or supersonic flow field, a rapid expansion of moist air or steam gives rise to non-equilibrium condensation. In the present study, the effect of non-equilibrium condensation of moist air on the self-excited shock wave oscillation around a circular arc blade with or without a bump on the blade was investigated numerically. The results showed that the nonequilibrium condensation and bump on the blade significantly reduced the flow field unsteadiness compared to the case without the non-equilibrium condensation and without bump.
Numerical simulations of the low-Reynolds number (Re=104~105) flows over a SD7003 airfoil at moderate incidences (<10°) are performed in the current paper. A lowfrequency convective instability is observed to dominate the spectrum near the leading edge and be responsible for the growth of the disturbance in the attached boundary layer. The characteristic frequency, the growth rate and the wave shape are investigated based on the numerical results. The growth of the low-frequency instability is not in agreement with parallel flow stability theory, nor with leading edge receptivity theory. And it has a higher growth rate than the Tollmien-Schlichting (T-S) wave. The effects of the angle-of-attack (AoA), the Reynolds number and the airfoil geometry on the low-frequency instability are investigated and discussed. © 2011 by Ying Zhou and Z. J. Wang. Published by the American Institute of Aeronautics and Astronautics, Inc.
We carried out a twin-rudder performance test in a circulating water channel using particle image velocimetry (PIV). We set the distance between two single rudders to 0.5C, 0.75C, and 1.0C, where C is the chord length (C=100 mm), and considered angles of attack of 10°, 20°, and 30°, respectively. A two-frame gray-level cross-correlation PIV method was used to obtain flow information. The twin-rudder model was constructed using a submerged NACA 0018 airfoil with a row arrangement. Throughout the experiments, the Reynolds number of the NACA 0018 airfoil was kept at 1.5×104. The measured results were compared with results of each case involving a single rudder. For the single rudder, flow visualization, time-mean velocity fields, and vorticity information were measured at angles of attack of 10°, 15°, 20°, and 25°. For the twin rudder, streamlines, instantaneous and time-mean velocity fields, vorticity, and turbulence intensity were measured and compared with each other to determine the critical distance between the single rudders. It was found that this critical distance should be less than 1.0C in order to decrease turbulence flow and vortices.
The ultrasonic flowmeter has been widely used in the industrial flow measurement. The flow measurement accuracy depends on the relationship of the profile-linear average velocity. But this relationship of the transition zone is not available at present.In this paper, the characteristics of transition flow with specific Re number in pipes are researched. The k-ɛ model and LES(Large Eddy Simulation) model are respectively used to calculate the flow field of the transition zone, and the experiment results show that the LES model is more effective than the k-ɛ model. The relationship of the profile-linear average velocity for the transition zone in pipes is obtained by the calculated results of the LES model and it is proved that there is a big error using the tradition relationship based on the turbulence flow to calculate the profile-linear average velocity relationship of the transition flow. The research results of this paper can improve the measurement accuracy of ultrasonic flowmeters and provide theoretical basis for the research on the whole transition flow.
Stability characteristics of aerofoil flows are investigated by linear stability analysis of time-averaged velocity profiles and by direct numerical simulations with time-dependent forcing terms. First the wake behind an aerofoil is investigated, illustrating the feasibility of detecting absolute instability using these methods. The time-averaged flow around an NACA-0012 aerofoil at incidence is then investigated in terms of its response to very low-amplitude hydrodynamic and acoustic perturbations. Flow fields obtained from both two- and three-dimensional simulations are investigated, for which the aerofoil flow exhibits a laminar separation bubble. Convective stability characteristics are documented, and the separation bubble is found to exhibit no absolute instability in the classical sense; i.e. no growing disturbances with zero group velocity are observed. The flow is however found to be globally unstable via an acoustic-feedback loop involving the aerofoil trailing edge as a source of acoustic excitation and the aerofoil leading-edge region as a site of receptivity. Evidence suggests that the feedback loop may play an important role in frequency selection of the vortex shedding that occurs in two dimensions. Further simulations are presented to investigate the receptivity process by which acoustic waves generate hydrodynamic instabilities within the aerofoil boundary layer. The dependency of the receptivity process to both frequency and source location is quantified. It is found that the amplitude of trailing-edge noise in the fully developed simulation is sufficient to promote transition via leading-edge receptivity.
On considère un écoulement compressible bidimensionnel, autour d'une cavité ouverte. Des d'instabilité, auto-entretenues par l'effet de rétroaction de l'écrasement de la couche de cisaillement sur le bord aval de la cavité, génèrent des émissions acoustiques qu'il faut réduire. Des simulations numériques directes (DNS) permettent d'obtenir, avec ou sans actionnement, un modèle précis de l'écoulement. A partir des champs issus de la simulation, des décompositions orthogonales de modes propres (POD) sont proposées pour bâtir, par projection de Galerkin sur les équations isentropiques, des modèles d'ordre réduit non linéaires en prenant en compte l'actionnement (le contrôle). Pour éviter la divergence temporelle, les coefficients du système dynamique non forcé sont calibrés par diverses approches originales dont une basée sur la sensiblité modale. A partir du système dynamique forcé par un actionnement multifréquentiel (présent aussi dans les DNS), un contrôle en boucle fermée linéaire quadratique gaussien est proposé sur un système linéarisé. La reconstruction de l'état est basée sur une estimation stochastique linéaire sur 6 points de pression. Le contrôle optimal obtenu s'avère être périodique à la fréquence du second mode de Rossiter, qui est exactement celles des instabilits auto-entretenues dans la cavité. Par introduction de ce contrôle dans les simulations numériques directes, nous avons obtenu une réduction du bruit (faible) sur la fréquence du contrôle. ABSTRACT : We consider a two dimensional compressible flow around an open cavity. The Flow around a cavity is characterised by a self-sustained mechanism in which the shear layer impinges on the downstream edge of the cavity resulting in an acoustic feedback mechanism which must be reduced. Direct Numerical Simulations (DNS) of the flow at a representative Reynolds number has been carried to obtain pressure and velocity fields, both for the case of unactuated and multi frequency actuation. These fields are then used to extract energy ranked coherent structures also called as the Proper Orthogonal Decomposition (POD) modes. A Reduced Order Model is constructed by a Galerkin projections of the isentropic compressible equations. The model is then extended to include the effect of control. To avoid the divergence of the model while integrating in time various calibration techniques has been utillized. A new method of calibration which minimizes a linear functional of error, based on modal sensitivity is proposed. The calibrated low order model is used to design a feedback control of the Linear Quadratic Gaussian (LQG) type, coupled with an observer. For the experimental implementation of the controller, a state estimate based on the observed pressure measurements at 6 different locations, is obtained through a Linear Stochastic Estimation (LSE). The optimal control obtained is periodic with a frequency corresponding to the second Rossiter mode of the cavity. Finally the control obtained is introduced into the DNS to obtain a decrease in spectra of the cavity acoustic mode.
In the present book we try to give a panoramic view of the origin of turbulence in near-wall shear layers by binding its different aspects for prototypical incompressible flows. Basically, the foregoing considerations reflect the authors’ experience in experimental studies of the transition process. We expect that the result of our efforts will be of interest to researchers, high-school teachers, and students involved in hydro-aerodynamics, stability and transition problems. To be in line with this general idea, there was a need to pick out the material for inclusion from plenty of available research data. In this way, our priority was a conceptual treatment of different - in particular modern - aspects of the problem, rather then delving too much into the details. An interested reader can find these details in recent publications on the transition problem which are referred to in the book.
The three-dimensional separated flow around a slender flat-plate delta wing with sharp leading-edge at a 12.5 deg angle of attack has been studied by solving the full compressible Navier-Stokes equations in the generalized curvilinear coordinates. The time integration is carried out by using the second-order LU-SGS implicit scheme. A fourth-order centered compact difference scheme is used for spatial derivatives. A sixth-order implicit filter is employed to reduce numerical oscillation. Non-reflecting boundary conditions are imposed at the far-field and outlet boundaries to avoid possible non-physical wave reflection. Parallel computing based on Message Passing Interface (MPI) has been utilized to improve the performance of the code. Two Reynolds numbers have been selected. At a lower Reynolds number of 50000 based on the chord length and the freestream velocity, the flow is stable and dominated by a pair of leading edge primary vortices. At a higher Reynolds number of 196000, the small-scale vortex shedding is observed near the leading-edge of the delta wing. The computational results are compared with the experimental work of Riley & Lowson (1998). The periodic shedding of small-scale vortical structures near the leading-edge has been studied in detail, and the vortex shedding is found to be associated with the Kelvin-Helmholtz-type instability and the secondary vortex. The period of vortex shedding is obtained from the time series of the three velocity components recorded near the leading-edge. The time-averaged features of the vortical structures are also discussed.
The three-dimensional separated flow around a slender flat-plate delta wing with sharp leading-edge at a large angle of attack has been studied by direct numerical simulation (DNS). The numerical simulation is performed by solving the full compressible Navier- Stokes equations in the generalized curvilinear coordinates. The time integration is carried out by using the second-order LU-SGS implicit scheme, which requires no tridiagonal inversion. A fourth-order centered compact difference scheme is used for spatial derivatives. A sixth-order implicit filter is employed to reduce numerical oscillation. Non-reflecting boundary conditions are imposed at the far-field and outlet boundaries to avoid possible non-physical wave reflection. Two Reynolds numbers have been selected. At a lower Reynolds number of 5 x lo4 based on the chord length and the freestream velocity, the flow is stable and dominated by a pair of leading-edge primary vortices. At a higher Reynolds number of 1.96 x 105, the Kelvin- Helmholtz-type instability as well as vortex shedding are observed near the leading-edge of the delta wing. The computational results are compared with the experimental work of Riley & Lowson (1998). The periodic vortex shedding process near the leading-edge has been studied in detail, and the vortex shedding is found to be associated with the Kelvin-Helmholtz-type instability and the secondary vortex. The period of vortex shedding is obtained from the time series of the three velocity components recorded near the leadingedge. The mean velocity profiles above the upper surface are compared with those obtained from the experiments.
Direct numerical simulation is used to study leading edge receptivity to free-stream disturbance. The full compressible Navier-Stokes equations in generalized curvilinear coordinates are solved by LU-SGS implicit scheme, which requires no tridiagonal inversion and is capable of being completely vectorized and parallelized. A fourth-order centered compact scheme is used for spatial derivatives and second-order Euler Backward scheme is applied for temporal discretization. A sixth-order implicit filter is employed to reduce the numerical oscillation. The non-reflecting boundary conditions are imposed at the far-field and outflow boundaries to avoid possible non-physical wave reflection. The code is developed in a form of total flow and no base flow with perturbation assumption is needed. The code also does not need buffer or sponge domain for DNS. The leading-edge receptivities of the compressible boundary-layer over the flat plate and the Joukowsky airfoil to free-stream disturbances are simulated.
High-order non-reflecting boundary conditions in a generalized curvilinear coordinate system for solving the time-dependent Navier-Stokes equations in complex geometry have been developed based on the characteristic analysis and the modified Navier-Stokes equation. Viscous terms are taken into account to include the viscous effect near the wall. All boundary conditions are added implicitly to the equations of interior points to ensure the stability of this scheme. The computational results show that the non-reflecting boundary conditions are compatible to the sixth- or fourth-order compact central difference scheme and maintain a high-order accuracy of the global solution. The non-reflecting boundary conditions work surprisingly well without any artificial buffer or sponge. No visible reflected wave was found from either inflow, outflow, far-field, or solid surface. The computational solution is found quite accurate in comparison with valid data.
The application of pulsed vortex generator jets to control separation on the suction surface of a low pressure turbine blade is reported. Blade Reynolds numbers in the experimental, linear turbine cascade match those for high altitude aircraft engines and aft stages of industrial turbine engines with elevated turbine inlet temperatures. The vortex generator jets have a 30 degree pitch and a 90 degree skew to the freestream direction. Jet flow oscillations up to 100 Hz are produced using a high frequency solenoid feed valve. Results are compared to steady blowing at jet blowing ratios less than 4 and at two chordwise positions upstream of the nominal separation zone. Results show that pulsed vortex generator jets produce a bulk flow effect comparable to that of steady jets with an order of magnitude less massflow. Boundary layer traverses and blade static pressure distributions show that separation is almost completely eliminated with the application of unsteady blowing. Reductions of over 50% in the wake loss profile of the controlled blade were measured. Experimental evidence suggests that the mechanism for unsteady control lies in the starting and ending transitions of the pulsing cycle rather than the injected jet stream itself. Boundary layer spectra support this conclusion and highlight significant differences between the steady and unsteady control techniques. The pulsed vortex generator jets are effective at both chordwise injection locations tested (45% and 63% axial chord) covering a substantial portion of the blade suction surface. This insensitivity to injection location bodes well for practical application of pulsed VGJ control where the separation location may not be accurately known a priori.
Transition arising from a separated region of flow is quite common and plays an
important role in engineering. It is difficult to predict using conventional models
and the transition mechanism is still not fully understood. We report the results of
a numerical simulation to study the physics of separated boundary-layer transition
induced by a change of curvature of the surface. The geometry is a flat plate with a
semicircular leading edge. The Reynolds number based on the uniform inlet velocity
and the leading-edge diameter is 3450. The simulated mean and turbulence quantities
compare well with the available experimental data.
Direct numerical simulation (DNS) for the flow separation and transition around a NACA 0012 airfoil with an attack angle of 4° and Reynolds number of 105 based on free-stream velocity and chord length is presented. The details of the flow separation, detached shear layer, vortex shedding, breakdown to turbulence, and re-attachment of the boundary layer are captured in the simulation. Though no external disturbances are introduced, the self-excited vortex shedding and self-sustained turbulent flow may be related to the backward effect of the disturbed flow on the separation region. The vortex shedding from the separated free shear layer is attributed to the Kelvin–Helmholtz instability.
The influence of a surface roughness element in the form of a two-dimensional hump on the transition location in a two-dimensional subsonic flow with a free-stream Mach number up to 0.8 is evaluated. Linear stability theory, coupled with the N-factor transition criterion, is used in the evaluation. The mean flow over the hump is calculated by solving the interacting boundary-layer equations; the viscous-inviscid coupling is taken into consideration, and the flow is solved within the separation bubble. The effects of hump height, length, location, and shape; unit Reynolds number; free-stream Mach number, continuous suction level; location of a suction strip; continuous cooling level; and location of a heating strip on the transition location are evaluated. The N-factor criterion predictions agree well with the experimental correlation of Fage; in addition, the N-factor criterion is more general and powerful than experimental correlations. The theoretically predicted effects of the hump's parameters and flow conditions on transition location are consistent and in agreement with both wind-tunnel and flight observations.
A three-dimensional numerical method based on the lower-upper symmetric-Gauss-Seidel implicit scheme in conjunction with the flux-limited dissipation model is developed for solving the compressible Navier-Stokes equations. A new computer code which is based on this method requires only 9 microsec per grid-point per iteration on a single processor of a Cray YMP computer and executes at the sustained rate of 170 MFLOPS. A reduction of 4 orders of magnitude in the residual for a high Reynolds number flow using 230 K grid points is obtained in 24 minutes. The computational results compare well with available experimental data.
Two new techniques for the study of the linear and nonlinear instability in growing boundary layers are presented. The first technique employs partial differential equations of parabolic type exploiting the slow change of the mean flow, disturbance velocity profiles, wavelengths, and growth rates in the streamwise direction. The second technique solves the Navier-Stokes equation for spatially evolving disturbances using buffer zones adjacent to the inflow and outflow boundaries. Results of both techniques are in excellent agreement. The linear and nonlinear development of Tollmien-Schlichting (TS) waves in the Blasius boundary layer is investigated with both techniques and with a local procedure based on a system of ordinary differential equations. The results are compared with previous work and the effects of non-parallelism and nonlinearity are clarified. The effect of nonparallelism is confirmed to be weak and, consequently, not responsible for the discrepancies between measurements and theoretical results for parallel flow.
Direct numerical simulation of boundary-layer receptivity for subsonic flow around airfoil. Recent Advances in DNS and LES
- L Jiang
- H Shan
- C Liu
Jiang, L. Shan, H., and Liu, C. 1999. Direct numerical simulation of boundary-layer receptivity
for subsonic flow around airfoil. Recent Advances in DNS and LES, Proceedings of the Second
AFOSR (Air Force Office of Scientific Research) International Conference. Rutgers, New
Jersey, June 7-9.
Non-reflecting boundary conditions for DNS in curvilinear coordinates. Recent Advances in DNS and LES
- L Jiang
- H Shan
- C Liu
Jiang, L. Shan, H., and Liu, C. 1999. Non-reflecting boundary conditions for DNS in curvilinear
coordinates. Recent Advances in DNS and LES, Proceedings of the Second AFOSR (Air Force
Office of Scientific Research) International Conference. Rutgers, New Jersey, June 7-9.
Numerical simulation of complex flow around a 85 delta wing
- H Shan
- L Jiang
- C Liu
Shan, H., Jiang, L., and Liu, C. 2001. Numerical simulation of complex flow around a 85 delta
wing. Proceedings of the Third AFOSR (Air Force Office of Scientific Office) International
Conference on DNS/LES. Arlington, Texas, August 5-9.