There were two defined models for analysis and this analysis has been done in CFX post-processor. First model without any boundary layer control surfaces has been analysed and then model with combination of suction and blowing boundary layer control surfaces. Three dimensional simulations have been done by using simulation software ANSYS CFX. In post processing results have been analysed in the form of velocity vectors. Velocity vectors has been checked not at various section of span but also at two models with changed length units and rpm.
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A review is given to the major developments in the boundary layer control by means of micro vortex generator (MVG). The appearance of MVG dates back to about three decades from the present review and this class of device rapidly spreads as an efficient and robust boundary layer control strategy for flows with separation in particular. Substantial successful applications of MVG have been achieved in subsonic flows and its application into supersonic flow is under progressive development, which requires the current status and capability of MVG to be reviewed. The fundamental aspects of MVG are first summarized, including its working principle and various geometries. Understanding of the fluid mechanics involved in the wake of a MVG device is of uppermost importance: the time- averaged and instantaneous wakes are discussed respectively. In the former, effects from the MVG geometry parameters are presented; while in the latter, focus is placed on the phenomenon associated with flow instability inside the wake. MVG’s applications in the subsonic regime are discussed, which covers the fundamental flow separation investigations and studies to tackle several practical problems. The ongoing studies in supersonic regime dealing with shock wave boundary layer interaction (SWBLI) induced separation are reviewed later, where the presentation is organized according to different types of SWBLI with emphasis on the interaction region.
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Abstract. Improving the aerodynamic performance of an airfoil is one of the primary interests of the Aerodynamicists. Such performance improvement can be achieved using passive or active flow control devices. One of such passive devices having a compact size along with an effective performance is the Micro Vortex Generators (MVGs). A special type of MVGs, which has been recently introduced in the aerospace industry, is “Triangular Shape” MVGs and its impact on aerodynamic characteristics is the main interest of this study. This study will compare the effects of various configurations through which delay of the flow separation using boundary layer control will be analysed by experimental and theoretical approach. The experimental investigations have been conducted using subsonic wind tunnel and the theoretical analysis using ANSYS® 13.0 FLUENT of which the final results are compared with each other.
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Simulation results presented in this study are in line with the experimental results in almost all cases. That is, the results validate the experimental data. At all velocities maximum lift coefficient increases with the rpm. The same does not hold for the drag coefficients since their variation is irregular. The drag coefficient remains between 0.13 and 0.24 in most cases. At 20 ms −1 stall angle increases at all rpm at a constant rate, except for the zero rpm case. More specifically, the value of the lift coefficient is driven by the delay in flow separation at higher angles of attack in view of momentum injection by the leading edge cylinder. The Magnus effect of the leading edge cylinder is a second major contributor to the improvement of the overall aerofoil coefficient of lift. The leading edge rotating cylinder acts like the vane less compressor, compressing the air in the channel and supplying the high pressure air to the bottom surface and the front of an aerofoil, which in turn contributes marginally to the improvement in aerofoil performance. This preliminary study indicates the feasibility of practical implementation of leading edge cylinders for thick aerofoil in aviation, wind turbine, and other applications. The use of the above-mentioned modifications to the aerofoil would add further layer of complexity for the structural health monitoring (SHM) system due to the addition of cylinder-driven vibrations [41–51].
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As indicated previously the ‘cone sheet’ was manufactured by rolling two sheets of metal (one perforated, one solid) to form two longitudinal halves of a cone. The 26 gauge (0.45 mm) steel used for the perforated and solid sheets came from the same stock in order to insure that no spurious effects were introduced. The intention was that these two halves would be welded together along the longitudinal seam to form the ‘cone sheet’. This was a significant challenge since the small diameter (the front end) of the cone was only 25.2 mm (0.993 in). The smallest rollers used by precision sheet metal manufacturers are 50.8 mm (2.0 in); therefore, conventional rolling procedures were not possible. The only feasible method for forming the desired cone was by manually using a press brake which required considerable operator skill and expertise. In particular, making the two separate halves match exactly along the seam while still maintaining a reasonable tolerance on the overall half-angle of the cone was quite difficult. Another issue was that the laser drilling process resulted in significant residual stresses in the sheet that needed to be relieved. For this reason, a ‘rough cut’ was first performed to allow the material to relieve before performing the final precision cut to provide the required shape to accurately form the cone half. A local company, Lane & Roderick, succeeded in ‘rolling’ the final cone halves and benefited greatly from the two prototyping iterations. The subsequent welding of the cone halves was also not trivial since conventional welding of such thin materials typically results in significant warping and distortion. A specialty shop, California Lasers, performed a laser fusion weld that did not use filler material and essentially kept the sheet at room temperature except in the immediate vicinity of the weld. This technique minimized any surface rippling and resulted in a very fine weld that minimized the surface area that would affect the boundary layer. Refer to Figure C.4 for a micrograph of the final weld.
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For a flow at 50 m/s in a wind tunnel with no boundary layer control where the vehicle is 10 meters downstream from the start of the test section δ ≈ 116 mm. In the case of a motorsport vehicle that has a splitter or wing with a ground clearance of 10 mm, the boundary layer in this example would envelope a key aerodynamic device. This is of great concern in motorsport applications, as these vehicles routinely operate at ground clearances of 10 mm. In motorsport applications, in order to provide the necessary boundary layer control, wind tunnel tests are done in a wind tunnel with a moving belt floor and with a vehicle whose wheels rotate. Unfortunately, in this type of wind tunnel, it is very difficult to measure forces, potentially leading to a loss of resolution. For a wind tunnel simulation to model aerodynamic forces accurately, the pressure distribution must be precisely accounted for. Pressure is related to velocity, in that the pressure at every point on the vehicle is proportional to the square of the local velocity. Even a minor error in the velocity distribution can have a significant impact on the test results. Therefore, for both the wind tunnel and a CFD simulation, achieving a near-flat velocity gradient upstream of the vehicle is desired.
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Abstract G¨ ortler vortices develop along concave walls as a result of the im- balance between the centrifugal force and radial pressure gradient. In this study, we introduce a simple control strategy aimed at reducing the growth rate of G¨ ortler vortices by locally modifying the surface geometry in span- wise and streamwise directions. Such wall deformations are accounted in the boundary region equations (BRE) by using a Prandtl transform of dependent and independent variables. The vortex energy is then controlled via a classical proportional control algorithm for which either the wall-normal velocity or the wall shear stress serves as the control variable. Our numerical results indicate that the control algorithm is quite effective in minimizing the wall shear stress. Keywords Boundary layer control · Gortler vortices
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boundary layer thickness increases with increasing n for shear thinning fluid (n<1), but in the shear thickening (n>1), boundary layer thickness decreases with increasing n beyond ξ=1. Flow behavior index n, surface temperature exponents a 5 and a 6 and Prandtl number have profound on local Nusselt
The present numerical investigation focuses to active separation control using vortex generator jets. VGJs constitute an AFC strategy in which fluid is injected into the boundary layer through an array of small holes. Initially, most of the research regarding VGJs was focused on the application of VGJs on the turbulent boundary layers [17,18]. The use of VGJs for laminar separation control is motivated by the Wind tunnel experiments at the Air Force Research Laboratory (AFRL) at Wright-Patterson Air Force Base [7,9,19,20]. Low pressure turbines (LPTs) are important components of many modern jet engines. The experimental investigations of Bons et al. [19,20] and Sondergaard et al.  at AFRL showed that laminar separation from LPT blades for chord Reynolds numbers between 25,000 and 100,000 can be drastically controlled by vortex generator jets. They reported that both steady and pulsed actuation to be effective in reducing separation losses. Whereas pulsed blowing requires a fraction of the momentum compared to the steady VGJs, pulsed forcing was shown to be more efficient than steady case.
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In aerodynamic wind tunnel testing, it is sometimes necessary to artificially trigger the acceleration of the turbulence in those cases where the measurement results are significantly affected by the existing laminar boundary layer. Very often, the strips with roughness elements are used at varying distances at the leading edge of testing airfoil or wings when there is a need to fix the boundary layer transition.
“Cognition enhanced” or “Self-Optimizing” have been used synonymously for control systems in the recent literature. Their basic principle has been introduced by Skogestad in 2000 for the application of chemical processes . Later on, the idea has been put into application in  as part of the Cluster of Excellence 128 “Integrative Production Technology for High-Wage Countries” for the area of manufacturing as well as in  within the Collaborative Research Center 614 “Self-Optimizing Concepts and Structures in Mechanical Engineering” for the area of mechatronic systems: A classic control loop is enhanced with a system that is able to change the control parameters and strategy as well as the set value. Through sensors, it determines the current status of the controlled systems, the overall goals as well as boundary and environmental conditions and deducts the new control
counteracting the G¨ ortler vortices. The kinetic energy associated with the vortices was shown to decrease by almost three orders of magnitude from the original amplitude in some cases, with a commensurate reduction in the wall shear stress. Our results clearly indicate that the controlled surface deformations are more effective than wall transpiration in altering the stream-wise development of the G¨ ortler vortices for a given roughness height and span-wise wavelength. Potential explanations for this difference were given, and the physical mechanism behind the control was described. It was found that for control based on wall deformation, a positive displacement of the wall accelerates fluid particles in the stream-wise direction, while negative displacement decelerates them, which means that the fluid momentum at the wall is increased, weakening the ’lift-up’ effect. The opposite was shown to occur at negative displacement locations, where the momentum of fluid particles at the wall is decreased. On the other hand, for control based on wall transpiration, blowing is applied in regions where the vertical velocity in the flow is negative, and suction is applied where the vertical velocity in the flow is positive (this is similar to the ’opposition control’ mechanism that was introduced in the framework of control of turbulent channel flows, Choi et al. (1994)), which is the main mechanism behind the reduction of the energy of the G¨ ortler vortices. The fact that wall deformation-based control was found to be more effective than wall transpiration in our formulation of the control problem is clear in the discussion and pictorial representation of velocity field and streamlines of both types of control mechanisms (see figures 18-21). While the boundary layer was excited by a row of roughness elements as derived in Goldstein et al. (2010), the algorithm can be easily extended to include freestream disturbances or turbulence at the upstream boundary. The proposed control algorithm has potential application to a number of realistic problems including aircraft wing design, to mitigate induced drag, or in supersonic wind tunnels, where it is known that flow instabilities in the test section are triggered by G¨ ortler vortices developing on the concave portion of the walls. Of course, the latter application will necessitate the extension of the algorithm in the compressible regime, which is a matter of future analysis.
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it was purely a study on understanding the effect of activation energy, wall boundary condition, and wall temperature on ignition. Sano and Yamashita (1994) performed thorough two-dimensional simulations of ignition in a methane-air laminar boundary layer over a horizontal hot plate. The reaction mechanism for methane-air consisted of 18 species and 61 elementary reactions. The plate consisted of a hot section placed between two cold sections. The study focused on the effects of wall temperature, flow velocity, hot plate length, and species boundary condition at the wall, i.e. (a) species concentration vanishing at the wall (reactive) and (b) zero species diffusion to the wall (inert), on ignition. Sano and Yamashita (1994) used an ignition criterion based on the peak of the methane consumption rate to determine an ignition delay time. Changing the flow velocity from 1 to 100 cm/s did not affect the ignition delay time since ignition occurred close to the wall. In addition, they found that decreasing the hot surface length led to an exponential increase in the delay time whereas increasing the length resulted in a constant delay time. Finally, to test the effect of a reactive wall, several species mass fractions were set to zero at the wall, e.g. H, O, OH, HO 2 ,
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The location of laminar to turbulent transition of the boundary layer, as estimated from the surface pressure measurements, appears to shift significantly between the lowest and highest Reynolds numbers tested. In a few cases the flow visualisation also provides indications of boundary layer transition. More definitive measurements of the boundary layer state are de- sirable given the importance of the boundary layer state to the overall flow. To perform these measurements a three-dimensional traversing system and a fast response total pressure probe (FRTPP) were designed and manufactured. In addition, methods were implemented to enable both accurate determination of the model position and examination of the boundary layer state.
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characteristics of this particular turbulent boundary layer, the tunnel itself, and the atmospheric boundary layer modelling can be found in report . It has to be mentioned here, that the turbulence intensity falls from 30% close to the ground to 5% at 900 mm altitude. To complete the description of BLWT, the dimensions of the rectangular measurement area were 1.8 x 1.5 meters. For completeness, the second tunnel used for comparison was Low Speed Wind Tunnel of Eiffel Type with the low turbulence level (up to 1%).
The dynamic modulation can be varied by changing the modulation function profile, signal amplitude and frequency. To isolate the effect of dynamic modulation, all other control conditions were maintained the same. In this test, 8 mm sized air- films were sustained over a large dimple surface (Design 4 dimple surface), i.e., 12 air-films were sustained along the streamwise direction in TBL with Re θ equal to 1200. In the assembled FOV (with a streamwise span from 0 x 0 + to 3900 x + 0 ), the air-films were between 760 x + 0 and 3200 x 0 + . The air-film was initially set in bulged configuration with a fixed maximum height of 2 mm, although it may be changed by dynamic modulation. The dynamic modulation output power was set at two levels, with the higher-level output power four times larger than the lower-level output power. The modulation frequencies were selected within the range between 5 Hz and 50 Hz. Both sinusoidal and saw-tooth shaped modulation profiles were tested. The tested parameters and corresponding air-film oscillation information are listed in Table 9.1.
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Irregularity - Turbulent flows are irregular and chaotic. They are the product of a broad spectrum of different eddy-sizes/length scales. It is however, contrary to popular misconceptions, deterministic and fully described by the Navier-Stokes equations. That is to say for two separate ‘numerical experiments’ with the same set of input conditions that identical results will be produced. However, the background acoustic noise that one may observe during physical experimentation may, on the other hand, essentially be random. In order to sufficiently resolve such issues a valid DNS approach may have to include the full wind tunnel and possibly the laboratory environment itself in the calculation, which would of course be exceptionally (and prohibitively) expensive. As such many of the inflow boundary conditions on numerical simulations, which solve some form of the Navier- Stokes equations, utilise random algorithms to superimpose turbulent structures on the incoming mean flow to represent freestream turbulence, see the synthetic-eddy-method of Jarrin et al. (2006) for example. Theoretically, with a well-designed synthetic inflow boundary condition, which is representative of the complementary experimental set-up, it would be possible to obtain convergence with the statistically steady mean properties. This would hold true if the Reynolds averaged results for solution time tending towards infinity and, likewise, an infinite number of ensemble averaged experiments/Reynolds averaged experiments over infinite sampling time.
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by Incropera & DeWitt (1996, p. 319) and Tritton (1995, p. 170) for small-Pr (Pr < 1) ﬂuids. However, for Pr < 1 this balance only exists in the near-wall region, where ∂ 2 v/∂x 2 is positive. In the outer region the ﬂow is driven in the same direction via viscous eﬀects and buoyancy; these quantities have the same sign and a balance will exist between them and inertia, providing a diﬀerent scaling. It is therefore not possible to obtain a single scaling for the velocity structure over the full velocity boundary-layer width. Nonetheless it is clear that for Pr < 1 the overall (outer) velocity boundary-layer thickness has the same scaling and approximately the same magnitude as the thermal boundary-layer thickness, supporting the statement of Gebhart et al. (1988).
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The research on the use of plasma actuators for flow control is carried out by many research centers around the world. Numerous modifica- tions are being introduced, among others; chang- es in the shape of electrodes , modification of the power supply system (frequency, voltage, power), number and placement of electrodes on the wing. Most of the experimental tests carried out are with an asymmetrical configuration of electrodes. Such configurations are presented in Figure 3.
Therefore, we obtain the boundary values of un- known function as detached. Now we could as- sume equation (2.1) with conditions (4.10) and (4.11) is a boundary value problem with local boundary conditions (Dirichlet conditions).For this problem we can easily apply classic meth- ods for recognizing formation or nonformation of boundary layers. If we let: