Swept-Wing Active Flow Control with a Streamwise Row of Vortex Generating Jets

A computational study was performed of a flow-excited Helmholtz resonator. This setup was used to model the phenomenon that occurs when a sunroof is open, or rear passenger compartment window lowered, in an automobile. The results of the computational study were validated through comparison to experimental and numerical results found in literature. After validating the setup of the computational study, the research then focused on using methods of active flow control to reduce the sound pressure levels and fluctuations within the resonator. Specifically, this was accomplished through the use of a small elongated jet placed upstream of the resonator opening. After seeing changes in the pressure levels using this active flow control technique, the velocity of the jet was altered to study its effect on the level of pressure reduction. The final results from this computational study show the desired outcome of a reduction in pressure and it was concluded that air from the jet that is added to the boundary layer produces an effect similar to that of running the simulation at a higher free stream velocity, but without active flow control. Thus, using the jet for active flow control will provide the ability to reduce pressure levels across a spectrum of free stream velocities.

Shock Wave-Boundary Layer Interaction (SBLI) is a phenomenon occurring in high-speed propulsion systems that is highly undesirable. Numerous methods have been tested to manipulate and control SBLI which includes both active and passive flow control techniques. To determine the improvements brought by the flow control techniques, advanced and state-of the-art flow diagnostics and experimental techniques are required, especially when it involves high-speed flows. In this study, a number of advanced flow diagnostics were employed to investigate the effect of micro-vortex generators in controlling SBLI in Mach 5 such as Pressure Sensitive Paints (PSP), Particle Image Velocimetry (PIV), schlieren photography and oil-flow visualization. The flow diagnostics successfully visualized the boundary layer separation and also the improvements brought by the micro-vortex generators.

As railway transportation advances toward higher speeds, traditional passive measures may struggle to meet the stringent aerodynamic criteria in tunnels, necessitating the exploration of novel active flow control techniques. This study employs three-dimensional, compressible, unsteady Reynolds-averaged Navier–Stokes simulations to investigate the aerodynamic effects of the suction and blowing slit area (S) positioned on the front and rear noses of the train. The results indicate that suction and blowing activation is particularly effective in alleviating pressure on the narrower side of the tunnel. Specifically, with a 4 m2 slit, the original 4.8% pressure difference between symmetrical points on the train body is fully eliminated. The influence of suction and blowing on the positive pressures is confined to the front and rear noses where the slits are located. Notably, only suction at the front nose mitigates pressure gradients, while blowing at the rear is unrelated. The peak-to-peak pressure (ΔP) on both the train surface and tunnel wall exhibits a linear decline, with reductions of 17.4% and 16.6%, respectively, as S increases from 0 to 4 m2. Similarly, the slipstreams on both sides of the tunnel decrease linearly with increasing slit area: with u/Umax = −0.008S + 0.24 for the near side, and u/Umin = 0.014S − 0.265 for the far side. Additionally, expanding the slit area further boosts the stability and safety of the train during tunnel exit by reducing lateral forces and rolling moments, while also decreasing overall drag, thereby partially compensating for the energy input. Although the maximum lift on the head car increases with slit area, the lift on the tail car initially rises and then decreases, helping to mitigate instability upon tunnel exit. Overall, the hybrid suction and blowing technique offers promising potential for enhancing the tunnel aerodynamics in the future.

Among the various active flow control techniques, Plasma Synthetic Jet (PSJ) actuators, or Sparkjets, represent a very promising technology, especially because of their high velocities and short response times. A practical tool, employed for design and manufacturing purposes, consists of the definition of a low-order model, lumped element model (LEM), which is able to predict the dynamic response of the actuator in a relatively quick way and with reasonable fidelity and accuracy. After a brief description of an innovative lumped model, this work faces the experimental investigation of a home-designed and manufactured PSJ actuator, for different frequencies and energy discharges. Particular attention has been taken in the power supply system design. A specific home-made Pitot tube has allowed the detection of velocity profiles along the jet radial direction, for various energy discharges, as well as the tuning of the lumped model with experimental data, where the total device efficiency has been assumed as a fitting parameter. The best fitting value not only contains information on the actual device efficiency, but includes some modeling and experimental uncertainties, related also to the used measurement technique.

This paper presents a control mechanism for video compression and transmission that relies on spatio-temporal image segmentation and non-uniform resolution representation. In the proposed scheme, a linked pyramid structure is built over each frame. Therefore, the video sequence is processed as a 3D data batch instead of a series of 2D images. The spatio-temporal segmentation is based on the adaptive link principle. The set of links defined between nodes of a single pyramid structure is extended to consecutive pyramids. According to the adaptive link principle, links between consecutive pyramids are rearranged so that a given node at a pyramid is linked to an homogeneous region of pixels at its base, but also to an homogeneous region of pixels at the base of the previous pyramid. Hence, resulting regions are coherent in time but also in space and any region may be extracted and tracked through the sequence by following the link structure. Then, if we assume that the most important regions of the frames are typically those changing more often, they can be represented at high resolution, while the least important regions can be represented at low resolution. This operation allows a severe data volume reduction when transmitting the video signal. Besides, regions presenting few changes from frame to frame do not need to be transmitted as often as moving areas because they will be very similar during a few frames.

Turbulent boundary layers generate large wall pressure fluctuations and high skin friction drag, which in turn raise aerodynamic and acoustic penalties for aircraft, ships, and wind-energy devices. Active near-wall flow control techniques that target the dominant vortex structures have therefore attracted growing attention as a practical route to simultaneous noise and drag reduction. The impact of flow controls on zero-pressure gradient turbulent boundary layer flow at high Reynolds number was studied using Direct Numerical Simulation to investigate their influence on the altered vortex features and the resulting wall pressure fluctuations. Four distinct boundary layer flow control strategies, blowing (BW), suction (SC), spanwise wall oscillation (SWO), and streamwise-traveling wave with wall oscillation (STWO), were applied to examine their effects on enstrophy, vorticity, wall pressure, and their respective correlations. Turbulent statistics revealed that SC increased skin friction due to enhanced turbulence near the wall, while reductions were seen in BW, SWO, and STWO, with STWO showing the most significant decrease. Analysis of premultiplied energy spectra of the pressure fluctuations in SWO and STWO indicated notable suppression of wall pressure fluctuations and the associated vortex energies generated by the near-wall quasi-streamwise vortices. The correlation between wall pressure fluctuations and streamwise vorticity highlighted the effectiveness of these control methods, particularly with STWO, in redistributing energy and attenuating turbulence structures responsible for wall pressure fluctuations. Flow visualization and enstrophy analysis further confirmed that the control strategies substantially altered vortex structures, demonstrating the potential of SWO and STWO as effective approaches for reducing wall pressure fluctuations and drag mitigation.

[Abstract] This paper presents a low-order flow model for vortices over a delta wing suitable for feedback flow control design. It combines two fluid dynamics models based on conical flow assumptions for flow field over a delta wing  a theoretical wing rock model and an empirical vortex breakdown model, to develop a set of nonlinear state equations describing the dynamics of the flow field as well as the aircraft rolling motion through a series of mathematical transformations and simplifications. The current model assumes virtual control inputs that directly affect the rate (derivative) of the state variables, such as the rate of vortex strength, vortex core velocity and velocity of the vortex break-down point. This generic assumption allows a variety of active flow control effectors and techniques to be applied to the model with suitable (static) modeling of the mapping from the control effectors to the virtual control inputs. By virtue of the conical flow assumption, the model output variables are assumed to be pressure sensors equally spaced span-wise on both upper and lower surface of the wing at a chosen chord-wise (longitudinal) location. The state equation model presented here has been validated with the two existing theoretical and empirical models for the same delta wing configurations that used in the validation of the existing models.

We consider the application of active flow control to modify the radial pressure distribution of a single-phase wall-normal vortex. The present flow is based on the Burgers vortex model but with a no-slip boundary condition prescribed along its symmetry plane. The wall-normal vortex serves as a model for vortices that emerge upstream of turbomachinery, such as pumps. This study characterizes the baseline vortex unsteadiness through numerical simulation and dynamic mode decomposition. The insights gained from the baseline flow are used to develop an active flow control technique with rotating zero-net-mass blowing and suction with the objective of modifying the core-pressure distribution. The effectiveness of the control strategy is demonstrated by achieving a widened vortex core with increased pressure. This change in the flow field weakens the local strength of the wall-normal vortex core, potentially inhibiting the formation of hollow-core vortices, commonly encountered in liquids.

Experimental study on pulsed suction under different momentum coefficients in variable compressor cascade with annular gaps

The present contribution summarizes the efficiency of active flow control of the boundary layer of air around the hump. The synthetic jet generator with rectangular output part, i.e. slot, is actuated using a modulated signal. The modulated signal phase shift and its influence on the loss coefficient is investigated. The synthetic jet influences the character of the boundary layer and the size of the mixing area in a positive way. Decrease of the loss coefficient resulting from the change of the mixing size area (i.e. wake) is investigated. A velocity profile and a pressure profile were measured behind the hump by the CTA probe and the Kiel probe, respectively. Loss coefficients were obtained from total pressure measurements, each with a different measurement setup. These coefficients were compared with each other to show the benefits of the synthetic jet active flow control technique.

A new high-speed linear cascade has been developed for low-pressure turbine (LPT) studies at The Ohio State University. A compressible LPT profile is tested in the facility and its baseline performance at different operating conditions is assessed by means of isentropic Mach number distribution and wake total pressure losses. Active flow control is implemented through a spanwise row of vortex-generator jets (VGJs) located at 60% chord on the suction surface. The purpose of the study is to document the effectiveness of VGJ flow control in high-speed compressible flow. The effect on shock-induced separation is assessed by Mach number distribution, wake loss surveys and shadowgraph. Pressure Sensitive Paint is applied to understand the three dimensional flow and shock pattern developing from the interaction of the skewed jets and the main flow. Data show that with increasing blowing ratio the losses are first decreased due to separation reduction, but losses connected to compressibility effects become stronger due to increased passage shock strength and jet orifice choking; therefore the optimum blowing ratio is a tradeoff between these counteracting effects. The effect of added surface roughness on the uncontrolled flow and on flow control behavior is also investigated. At lower Mach number turbulent separation develops on the rough surface and a different flow control performance is observed. Steady VGJs appear to have control authority even on a turbulent separation but higher blowing ratios are required compared to incompressible flow experiments reported elsewhere. Overall, the results show a high sensitivity of steady VGJs control performance and optimum blowing ratio to compressibility and surface roughness.

This paper investigates the potential of using an active flow control technique to modify stall flutter oscillations of a NACA (National Advisory Committee for Aeronautics) 0015 wing section. Wind tunnel experiments have been performed with a test-rig that provides the elastic degree of freedom in pitch. Measurements of the clean airfoil are taken at preset angles of θ0=6°−12°, and for Reynolds numbers of Rec=6.2×104−1.25×105, which reveal the dependency of the stall flutter oscillations to Rec and θ0. Then, flow control experiments are carried out at θ0=10° and Rec=1.04×105. Two dielectric barrier discharge plasma actuators have been employed simultaneously to exert dual-point excitation to the baseline flow. It is shown that during the upstroke half-cycle, plasma actuation postpones the dynamic stall of the airfoil and increases the maximum pitch angle of the stall flutter cycle. On the downstroke, dual-point excitation effectively improves the rate of pitching moment recovery and leads to flow reattachment at a larger pitch angle. Normalized excitation frequencies Fex=fex/fw (where fw is the wake mode frequency of the stalled airfoil) ranging from 0.1 up to 3 have been examined. Among the controlled cases, excitation with Fex=0.6 and Fex=2.2 provides the largest and smallest pitch amplitude, respectively, and the case of Fex=3 demonstrates the most impact on flow reattachment. Finally, it has been concluded that the employed control strategy is effectively capable of modifying the dynamic stall process and associated pitching moment. However, a more sophisticated control strategy would be required to significantly mitigate the stall flutter oscillations.

To overcome the convective time delay issue for active closed-loop flow control, a model based predictive control algorithm is analyzed. From a controls perspective, forms of internal model control or model predictive control can be used to accommodate and minimize the effect of systems with pure time delays. Moreover, the Smith predictor is a commonly employed control technique to negate the pure time delay in a closed-loop system. This form of model predictive control is applied to the asymmetric vortex problem of an axisymmetric forebody (specifically a von Karman ogive with fineness ratio of 3.5). The full-order Navier-Stokes equations are numerically solved on the forebody at a high angle of attack and provide the plant process. Small port and starboard blowing patches are used to introduce fluidic disturbances at the nose of the ogive to augment the global flow state and produce a deterministic vortex state. As the active flow control technique exploits a convective instability, a convective time delay exists. Linear-time-invariant and non-linear time-invariant models are developed from the open-loop dynamics. A Smith predictor is employed within the full order CFD simulation. The results of the predictive control are compared to open-loop and model-free closed-loop behavior. It is shown that the predictive control developed in this paper while very suitable for control of this type of flow, is very sensitive to modeling uncertainties.

One of the strategies used to improve performance and increase the life-span of wind turbines is active flow control. It involves the modification of the aerodynamic characteristics of a wind turbine blade by means of moveable aerodynamic control surfaces. Trailing edge flaps are relatively small moveable control surfaces placed at the trailing edge of a blade's airfoil that modify the lift of a blade or airfoil section. An instrumented wind turbine test rig and rotor were specifically developed to enable a wide-range of experiments to investigate the potential of trailing edge flaps as an active control technique. A modular blade based on the S833 airfoil was designed to allow accurate instrumentation and customizable settings. The blade is 1.7 meters long, had a constant 178mm chord and a 6° pitch. The modular aerodynamic parts were 3D printed using plastic PC-ABS material. The blade design point was within the range of wind velocities in the available large test facility. The wind facility is a large open jet wind tunnel with a maximum velocity of 11m/s in the test area. The capability of the developed system was demonstrated through an initial study of the effect of stationary trailing edge flaps on blade load and performance. The investigation focused on measuring the changes in flapwise bending moment and power production for different trailing edge flap spanwise locations and deflection angles. The relationship between the load reduction and deflection angle was linear as expected from theory and the highest reduction was caused by the flap furthest from the rotor center. Overall, the experimental setup proved to be effective in measuring small changes in flapwise bending moment within the wind turbine blade and will provide insight when (active) flap control is targeted.

Active flow control of the airflow of a ship at yaw