Flow Separation Control Research Articles

The control of flow separation: Study of optimal open loop parameters

Effective parameters in active flow control (AFC) have been studied to clarify a perspective of AFC. In particular, NACA0015 airfoil has been chosen for this study. Fluidic devices such as a synthetic jet actuator (SJA) have been considered because these can enforce two main control parameters of AFC, frequency and momentum ratio, but fluidic excitation with complete blowing and suction cycle was applied with the harmonic excitation as the boundary condition. Flows around the NACA0015 airfoil were simulated for a range of operating conditions. Attention was paid to the active open loop control for the flow separation of the airfoil with SJA at different angles of attack and flap angles. A large number of simulations using unsteady Reynolds averaged Navier-Stokes and large eddy simulation models were performed to study the effects of momentum ratio (Cμ) in the range of 0% to 11% and non-dimensional frequency, F+, in the range of 0–2 for the control of flow separation at the various angles of attack. The optimum value of Cμ as well as F+ was evaluated and discussed. The computational model predictions showed good agreement with several experimental data available such as NASA Langley, Texas A&M, and Clarkson University. It was observed that different angles of attack and flap angles have different requirements for the minimum value of the momentum coefficient, Cμ, for the SJA to be effective for control of separation. It was also found that the variation of F+ noticeably affects the lift and drag forces acting on the airfoil. An intuitive analogy was introduced to evaluate the optimum momentum ratio in AFC, and the results of the optimum momentum ratio have been presented in comparison with available literature data.

Flow Separation Control of Backward-Facing Step Airfoil NACA0015 by Blowing Technique

The aim of this paper is to control the flow separation above backward-facing step (BFS) airfoil type NACA 0015 by blowing method. The flow field over airfoil has been studied both experimentally and computationally. The study was divided into two parts: a practical study through which NACA 0015 type with a backward -facing step (located at 44.4% c from leading edge) on the upper surface containing blowing holes parallel to the airfoil chord was used. The tests were done over two-dimensional airfoil in an open circuit suction subsonic wind tunnel with flow velocity 25m/s to obtain the pressure distribution coefficients. A numerical study was done by using ANSYS Fluent software version 16.0 on three models of NACA 0015, the first one has backward-facing step without blowing, the second with single blowing holes and the third have multi blowing holes technique. Both studies (experimental and numerical) were done at low Reynolds number (Re=4.4x105) and all models have chord length 0.27m.The experimental investigations and CFD simulations have been performed on the same geometry dimensions, it has been observed that the flow separation on the airfoil can be delayed by using velocity blowing (30m/s) on the upper surface. The multi blowing holes with velocity improved the aerodynamics properties.The multi blowing holes and single blowing hole thesame effect onpressure distribution coefficients

A comparative study of alternating current and nanosecond plasma actuators in flow separation control

A combined numerical and experimental study is performed to elucidate the difference in flow separation control mechanisms between alternating current (ac) and nanosecond (ns) plasma actuators. The focus is on the visualization of detailed flow actuation process and establishment of relationship between flow control authority and characteristics of different actuators. Time-resolved particle image velocimetry (PIV) system is used to capture the transient separated flow over a stalled NACA 0015 airfoil at low Reynolds number (Re) of 6.3 × 104 and response of the flow to plasma actuation. Reynolds-averaged two-dimensional computation is performed for both low and high Reynolds number cases at Re = (0.063, 0.75) × 106 to complement experiment. Utilizing the elaborate Reynolds stress turbulence model in conjunction with an empirical body force model for surface ac plasma discharge and a sophisticated self-similar formulation for ns pulsed discharge, we have been able to numerically reproduce the experimentally observed intricate mechanisms, e.g. discretization of unstable shear layer into individual vortices, formation of large-scale spanwise vortex, strong entrainment effect of vortices and flow regularization procedure. Moreover, the information from simulation allows us to make a number of important observations which have no experimental counterparts. The detailed simulation shows that the convection of residual heat from ns discharge with cross flow makes ns plasma actuator more flexible in perturbing far downstream fluid in comparison with other actuators, and makes it superior to other devices in some high Re number flows. Also observed is the occurrence of strong flow resonance and frequency lock-in at a proper range of forcing frequency. In addition, the continuous ac plasma actuator appears to have complex control mechanisms, depending on flow conditions.

Flow separation prevention around a NACA0012 profile through multivariable feedback controlled plasma actuators

In the present paper, the problem of controlling the unsteady flow separation over an aerofoil, using plasma actuators, is addressed. Although a plasma flow is essentially a two-phase flow, it has been proved that a local force term may be successfully considered within a flow solver in order to mimic the effect of a local acceleration (see [63]). By exploiting the Chimera Overlapping Grids technique to allow the inclusion of one or more actuators, accurate numerical simulations have been carried out with the in-house Finite Volume solver Xnavis. The effects of varying both number and position of the actuator/sensor pairs along the aerofoil are investigated by focusing on the dynamic control of the boundary layer separation. The capabilities of the proposed control strategy are also examined by showing how the practical flow separation control problem can be formulated as a simple output regulation problem, so that a simple control strategy may be used. Several numerical simulations of incompressible flows around a pitching NACA0012 at Reynolds Re=20,000 illustrate the effectiveness of the proposed approach, in the presence of time-varying angles of attack and complex non-linear dynamics, which are neglected in the control design. The present work extends the former paper of Broglia et al. [2], where the control algorithm was applied to the simple case of single input/single output regulation and implemented in a Finite Elements numerical code; in that work the boundary layer control of an airfoil at fixed angle of attack was addressed. In the present paper a multi input/multi output approach is theoretically developed and numerically implemented. Robust and fast flow reattachment is achieved, along with both stabilisation and increase/reduction of the lift/drag, respectively. A major advantage of the presented method is that the chosen controlled outputs can be easily measured in realistic applications.

Flow Separation Control in a Curved Diffuser with Rigid Traveling Wave Wall and Its Mechanism

Traveling wave wall is a useful method to suppress flow separation. However, the interaction between the traveling wave wall and unsteady separation flow is complex, which causes difficulty in discovering the corresponding mechanism. To reveal the mechanism of traveling wave wall control, numerical simulation of a separated curved diffuser using rigid traveling wave wall flow control is performed, which shows some unique characteristics. Then, a nonlinear simplified model is used to explain this phenomenon in flow control in consideration of nonlinear dynamics and order of degree. Flow field data from the numerical simulation are further analyzed using fast Fourier transform analysis, linear stability theory of free shear layers, and the nonlinear simplified model to reveal the control mechanism of traveling wave wall.

Numerical Simulation of Horizontal Axis Wind Turbines with Vortex Generators

In the present study, a simulation about the effects of vortex generators on horizontal axis wind turbine rotor blade was numerically conducted using a static coupled CFD–CSD method. A Navier–Stokes CFD flow solver based on unstructured meshes was used to obtain the blade aerodynamic loads. A FEM-based CSD solver employing a nonlinear coupled flap-lag-torsion beam theory was utilized to calculate the blade elastic deformation. The coupling of the CFD and CSD solvers was accomplished in a loosely coupled manner by exchanging the information between the two solvers at infrequent intervals. The static coupled CFD–CSD method was applied to the NREL 5 MW reference wind turbine rotor under steady axial flow conditions. Triangular counter-rotating vortex generators were adopted to control flow separation and radial flow in the inboard section of the NREL 5 MW reference rotor blades. They were installed on the inboard part of the blade from 0.2 to 0.4 R. As a result of the flow analysis considering the counter-rotating vortex generators, strong vortices were generated by counter-rotating vortex generators. It can be seen that the regions where flow separation and radial flow occur in the inboard sections were reduced compared to the baseline wind turbine. For this reason, the maximum power improvement due to counter-rotating vortex generators was 1.04% at the rated wind speed.

Active Stall Control of an Airfoil Using Synthetic Jets

Flow separation is mostly an undesirable phenomenon and flow separation control is very important on airfoils. The synthetic jet is one of the effective techniques for active flow control. The synthetic jet actuator is composed of a speaker or a piezoelectric diaphragm. The purpose of this study is to develop the active stall control system of NACA0012 airfoil’s upper surface. The system detects the flow conditions (stall or unstalled) by two static pressure data alone on the airfoil. When the system detects the stall, synthetic jets automatically operate to control the stall. The maximum lift increases 16 ％ and the stall angle can improved at 3 ° by using synthetic jet actuator at Re = 1.5×105 based on the chord length of the airfoil. Our proposed system can enhance the aerodynamic performance of the airfoil and separation control is always attained with no stall for all flow fields produced by changing the angle of attack examined in the present study.

Flow separation control over a rounded ramp with spanwise alternating wall actuation

An implicit large-eddy simulation is carried out to study turbulent boundary-layer separation from a backward-facing rounded ramp with active wall actuation control. This method, called spanwise alternating distributed strips control, is imposed onto the flat plate surface upstream of a rounded ramp by alternatively applying out-of-phase control and in-phase control to the wall-normal velocity component in the spanwise direction. As a result, the local turbulence intensity is alternatively suppressed and enhanced, leading to the creation of vertical shear-layers, which is responsible for the presence of large-scale streamwise vortices. These vortices exert a predominant influence on the suppression of the flow separation. The interaction between the large-scale vortices and the downstream recirculation zone and free shear-layer is studied by examining flow statistics. It is found that in comparison with the non-controlled case, the flow separation is delayed, the reattachment point is shifted upstream, and the length of the mean recirculation zone is reduced up to 8.49%. The optimal control case is achieved with narrow in-phase control strips. An in-depth analysis shows that the delay of the flow separation is attributed to the activation of the near-wall turbulence by the in-phase control strips and the improvement of the reattachment location is mainly due to the large-scale streamwise vortices, which enhance the momentum transport between the main flow and separated region.

Influence on Drag by Flow Separation Control around Car Door Mirror Using a String-type Plasma Actuator

Influence on Drag by Flow Separation Control around Car Door Mirror Using a String-type Plasma Actuator

Designing of Internet of Things Sensor Based Information Gateway Using SDN Concept

In order to make the gateway of the Internet of things access to the public network effectively, using a Software-Defined Network (SDN) concept, a sensor network was studied. When performing hierarchical design, the gateway of the Internet of things was layered. It was divided into the policy layer and the forwarding layer. The policy layer distributed the policy, to receive instructions from the upper server. At the same time, it managed the nodes and matches the forwarding rules. The forwarding layer is only responsible for forwarding data packets. This control and data flow separation structure could effectively improve the system transmission efficiency. At the same time, the software structure design and hardware design of the sensor gateway of the Internet of things were given. The design and function of the gateway was analyzed and studied. A comprehensive solution of the gateway system was proposed. The results showed that this method could further improve the performance of the sensor network gateway. Therefore, the design of the Internet of things sensing information gateway based on the concept of SDN has a high practical value.

Plasma Actuator Application on a Full-Scale Aircraft Tail

Pulsed dielectric barrier discharge (DBD) plasma actuators were evaluated for the control of flow separation from the V-tail panels of a well-known unmanned air vehicle, using lightweight high-voltage generators. Following risk-reduction airfoil experiments, full-scale experiments were performed on an unswept tapered V-tail panel at typical takeoff speeds (). Using less than 10 W of input power, significant poststall lift coefficient improvements (up to 0.6 or 100%) were observed, actuation was effective to , bistable behavior (hysteresis) was eliminated, and a mild sensitivity to pulsation frequency and duty cycle was observed. Takeoff performance calculations indicated a 30 m reduction in ground roll and a 65% increase in allowable crosswind speeds. The physical system proved to be robust in that the silicone rubber dielectric material did not oxidize or degrade during several hundred hours of continuous operation. In summary, pulsed DBD plasma proved to be a viable and practical solution to the problem of flow separation from the panels during crosswind takeoff. Before flight tests, future wind-tunnel experiments will employ high-thrust, fully-integrated actuators operable on both surfaces of the panels, and the effect of mild sweep, encountered during takeoff and landing, will be investigated.

Dual-position excitation technique in flow control over an airfoil at low speeds

PurposeThe purpose of this paper is to control flow separation over a NACA 4415 airfoil by applying unsteady forces to the separated shear layers using dielectric barrier discharge (DBD) plasma actuators. This novel flow control method is studied under conditions which the airfoil angle of attack is 18°, and Reynolds number based on chord length is 5.5 × 105.Design/methodology/approachLarge eddy simulation of the turbulent flow is used to capture vortical structures through the airfoil wake. Power spectral density analysis of the baseline flow indicates dominant natural frequencies associated with “shear layer mode” and “wake mode.” The wake mode frequency is used simultaneously to excite separated shear layers at both the upper surface and the trailing edge of the airfoil (dual-position excitation), and it is also used singly to excite the upper surface shear layer (single-position excitation).FindingsBased on the results, actuations manipulate the shear layers instabilities and change the wake patterns considerably. It is revealed that in the single-position excitation case, the vortices shed from the upper surface shear layer are more coherent than the dual-position excitation case. The maximum value of lift coefficient and lift-to-drag ratio is achieved, respectively, by single-position excitation as well as dual-position excitation.Originality/valueThe paper contributes to the understanding and progress of DBD plasma actuators for flow control applications. Further, this research could be a beneficial solution for the promising design of advanced low speed flying vehicles.

Numerical Investigation of Steady and Harmonic Vortex Generator Jets Flow Separation Control

Active flow control of canonical laminar separation bubbles by steady and harmonic vortex generator jets (VGJs) was investigated using direct numerical simulations. Both control strategies were found to be effective in controlling the laminar boundary-layer separation. However, the present results indicate that using the same blowing amplitude, harmonic VGJs were more effective and efficient at reducing the separated region than the steady VGJs considering the fact that the harmonic VGJs use less momentum than the steady case. For steady VGJs, longitudinal structures forming immediately downstream of the injection location led to the formation of hairpin-type vortices, causing an earlier transition to turbulence. Symmetric hairpin vortices were shown to develop downstream of the forcing location for the harmonic VGJs, as well. However, the increased control effectiveness for harmonic VGJs’ flow control strategy is attributed to the fact that the shear-layer instability mechanism was exploited. As a result, disturbances introduced by VGJs were strongly amplified, leading to the development of large-scale coherent structures, which are very effective at increasing the momentum exchange, thus limiting the separated region.

Practical Computational Methods for Airplanes with Flow-Control Systems

This survey paper consists of a review of flow-control applications using numerical simulations, where the goal is to provide guidance to the developers of flow-control approaches. The computational methods are described with the focus on practical simulations of flow control for flight vehicles. One approach uses the unsteady Reynolds-averaged Navier–Stokes formulation, and the other method is based on the lattice Boltzmann method. Special boundary conditions used to model a set of flow-control techniques and to guide the development of new actuation methods are highlighted. Results of studies aimed at solving a broad set of aerodynamic and propulsion flow problems through computational simulations are presented. The flow-control techniques are used for the control of flow separation as well as manipulating vortical flow structures for achieving a desired objective. The results include flow-control implementations for aircraft and rotorcraft targeting commercial and military applications. In cases where experimental data are available, selected flow-control scenarios are used to illustrate the validity of the computational methods in predicting realistic flows.

Optimal performance of adaptive flap on flow separation control

The stall caused by flow separation of airfoil seriously compromises the efficiency and safety of wind turbines. Inspired by bird feathers, a passive flow control strategy was investigated by configuring adaptive flaps on the suction surface of the airfoil to delay flow separation. The Computational Fluid Dynamics (CFD) software ANSYS Fluent 17.0 using SST k–ω turbulence model was performed to study the performance of the adaptive flap with different length and location. The performance was evaluated by the comparison of total lift coefficient between the flap airfoil and clean airfoil. In addition, the performance of the double flaps configuration was investigated by Fluid-Solid Interaction (FSI) method. The results show that when the separation point is at the upstream of flap hinge point, the total lift coefficient increases with the flap rising, and when the flap reaches the optimal flap angle, the lift coefficient begins to decrease. When the flap operates in optimal angle, the closer the flap is to the trailing edge of the airfoil, the more the lift enhancement for moderate flow separation, and the closer the flap is to the leading edge of the airfoil, the more the lift enhancement for serious flow separation. For the flaps with the same location, the longer flap can block more backflow resulting in a better performance, but when the flap is long enough to block the fluid, the shorter flap has better performance. Considering the variations of the flap aerodynamic moment with the flap angle, an external hinge moment was recommended to improve the swinging flap performance.

Three dimensional simulation of the effects of plasma actuator with radial arrangement of electrodes

The purpose of this article is using plasma actuator with special arrangement of electrodes to control flow separation. For this reason, a computational study is carried out on an unsteady, turbulent, and developing flow within a circular duct. The three phenomena including the cross-sectional increase, the developing flow and both tangential and radial simultaneous flow movement determine the physics of the problem. Due to the geometry, the electrodes for the first time are radially arranged.The results showed that presence of a plasma actuator transferred the separation point from r = 246.6 mm to r = 239.2 mm in radial direction. In addition, it was shown that the arrangement of electrodes in the flow direction does not play a role in separation point in radial direction.

Numerical Study of Unsteady Pulsed Suction through Endwall Bleed Holes in a Highly Loaded Compressor Cascade

Abstract Unsteady pulsed suction (UPS) was applied as an unsteady flow control (UFC) technique in a certain highly loaded compressor cascade to control the flow separations. Only two bleed holes symmetrically mounted on the endwalls (one on the upper endwall and another on the lower endwall) were set up to achieve steady constant suction (SCS) and UPS. The improvements in aerodynamic performance by SCS and UPS under different time-averaged suction flow rates are firstly investigated and compared. The related unsteady aerodynamic parameters of UPS such as excitation frequency, excitation location, pitch angle, and skew angle are discussed and analyzed in detail. The results show that UPS can provide a better flow control effect than SCS with the same time-averaged dimensionless suction flow rate in the control of flow separation. The aerodynamic performance of compressor cascades can be significantly enhanced by UPS when unsteady aerodynamic parameters are in their optimum ranges. Based on the optimum parameters for UPS, the total pressure loss coefficient is reduced by 19 % only with the time-averaged dimensionless suction flow rate ms =0.4 %.

Airfoil flow separation control with plasma synthetic jets at moderate Reynolds number

An array of 26 plasma synthetic jet actuators (PSJA) is flush-mounted on a NACA-0015 airfoil model to control the leading-edge flow separation at moderate Reynolds number (Re_c=1.7times 10^5). The stall angle of this airfoil is postponed from 15.5^circ to approximately 22^circ, and the peak lift coefficient is increased by 21%. PSJAs exhibit distinctive separation control mechanisms depending on the relative location between actuation and separation and reduced frequency of actuation (F^*). At an angle of attack of alpha =15.5^circ, the non-actuated flow separates approximately 4% chord length downstream of the jet orifices . Plasma synthetic jets (PSJs) applied at F^*ge 0.5 can displace the separation point downstream to mid-chord position, as a result of the energizing of the incoming boundary layer through mixing enhancement. As a comparison, with actuation frequency of F^*le 0.25, the separation point at alpha =15.5^circ remains near the leading edge and the zero-velocity line is periodically swept towards the suction surface by the convecting spanwise vortices generated from PSJ actuation, leading to a reduction of time-averaged backflow area. For the case of separation control at alpha =22^circ, the separation point resides always upstream of the actuation position, regardless of actuation frequency. The peak lift coefficient is attained at F^*=1, and the decreasing lift at high actuation frequency (F^*=2) is ascribed to the severe interaction between adjacent spanwise vortices at short spacing.Graphical abstract

Flow-separation-control system operating in feedback closed loop

PurposeThe purpose of this study is to develop the concept of self-adapting system which would be able to control a flow on the wing-high-lift system and protect the flow against strong separation.Design/methodology/approachThe self-adapting system has been developed based on computational approach. The computational studies have been conducted using the URANS solver. The experimental investigations have been conducted to verify the computational results.FindingsThe developed solution is controlled by closed-loop-control (CLC) system. As flow actuators, the main-wing trailing-edge nozzles are proposed. Based on signals received from the pressure sensors located at the flap trailing edge, the CLC algorithm changes the amount of air blown from the nozzles. The results of computational simulations confirmed good effectiveness and reliability of the developed system. These results have been partially confirmed by experimental investigations.Research limitations/implicationsThe presented research on an improvement of the effectiveness of high-lift systems of modern aircraft was conducted on the relatively lower level of the technology readiness. However, despite this limitation, the results of presented studies can provide a basis for developing innovative self-adaptive aerodynamic systems that potentially may be implemented in future aircrafts.Practical implicationsThe studies on autonomous flow-separation control systems, operating in a closed feedback loop, are a great hope for significant advances in modern aeronautical engineering, also in the UAV area. The results of the presented studies can provide a basis for developing innovative self-adaptive aerodynamic systems at a higher level of technological readiness.Originality/valueThe presented approach is especially original and valuable in relation to the innovative concept of high-lift system supported by air-jets blown form the main-wing-trailing-edge nozzles; the effective and reliable flow sensors are the pressure sensors located at the flap trailing edge, and the effective and robust algorithm controlling the self-adapting aerodynamic system – original especially in respect to a strategy of deactivation of flow actuators.

Validating a Reduced-Order Model for Synthetic Jet Actuators Using CFD and Experimental Data

Synthetic jet actuators (SJA) are emerging in various engineering applications, from flow separation and noise control in aviation to thermal management of electronics. A SJA oscillates a flexible membrane inside a cavity connected to a nozzle producing vortices. A complex interaction between the cavity pressure field and the driving electronics can make it difficult to predict performance. A reduced-order model (ROM) has been developed to predict the performance of SJAs. This paper applies this model to a canonical configuration with applications in flow control and electronics cooling, consisting of a single SJA with a rectangular orifice, emanating perpendicular to the surface. The practical implementation of the ROM to estimate the relationship between cavity pressure and jet velocity, jet velocity and diaphragm deflection and applied driving voltage is explained in detail. Unsteady Reynolds-averaged Navier Stokes computational fluid dynamics (CFD) simulations are used to assess the reliability of the reduced-order model. The CFD model itself has been validated with experimental measurements. The effect of orifice aspect ratio on the ROM parameters has been discussed. Findings indicate that the ROM is capable of predicting the SJA performance for a wide range of operating conditions (in terms of frequency and amplitude).