An Experimental Study of Synthetic Jet Actuators with Application in Airfoil LCO Control

An experimental study on the development and implementation of Synthetic Jet Actuators (SJAs) is conducted for eliminating aeroelastic phenomenon such as Limit Cycle Oscillations (LCO). One of the biggest challenges involved in the design of UAVs operating in unsteady atmosphere conditions is the susceptibility of the airframe to aeroelastic instabilities, such as flutter or LCO. Suppression of such instabilities can be achieved through the implementation of Active Flow Control (AFC) techniques, however to this day, a limited amount of experimental studies exist. Thus, the focus of this work is to develop a new AFC method consisting of an actuator that is directly instrumented in the internal volume of the airfoil. Due to the complex geometry of airfoil/actuator integration, advanced manufacturing technique has been employed for rapid manufacturing of these complex parts. In addition, a newly designed experimental test facility is fabricated to study the effect of the developed actuator on aerodynamic performance. Parametric analysis are conducted to investigate the effect of actuator along the airfoil surface, Reynolds number, and angle of attack. Results of this study demonstrated the actuator effectiveness on overall aerodynamic performance and show consistent trends with high-order Computational Fluid Dynamics (CFD)

Lift Enhancement for Low-Aspect-Ratio Wings with Periodic Excitation

In an effort to enhance lift on low-aspect-ratio rectangular flat-plate wings in low-Reynolds-number post-stall flows, periodic injection of momentum is considered along the trailing edge in this numerical study. The purpose of actuation is not to reattach the flow but to change the dynamics of the wake vortices such that the resulting lift force is increased. Periodic forcing is observed to be effective in increasing lift for various aspect ratios and angles of attack, achieving a similar lift enhancement attained by steady forcing with less momentum input. Through the investigation on the influence of the actuation frequency, it is also found that there exists a frequency at which the flow locks on to a time-periodic high-lift state

Experimental Investigation on a 3D Wing Section Hosting Multiple SJAs for Stall Control Purpose

Flow control over aerodynamic shapes in order to achieve performance enhancements has been a lively research area for last two decades. Synthetic Jet Actuators (SJAs) are devices able to interact actively with the flow around their hosting structure by providing ejection and suction of fluid from the enclosed cavity containing a piezo-electric oscillating membrane through dedicated orifices. The research presented in this paper concerns the implementation of zero-net-mass-flux SJAs airflow control system on a NACA0015, low aspect ratio wing section prototype. Two arrays with each 10 custom-made SJAs, installed at 10% and 65% of the chord length, make up the actuation system. The sensing system consists of eleven acoustic pressure transducers distributed in the wing upper surface and on the flap, an accelerometer placed in proximity of the wing c.g. and a six-axis force balance for integral load measurement. A dSPACE™ hardware connected to the software environment Matlab/Simulink® and dSPACE Control-Desk® complete the test architecture. Wind tunnel experiments, on the uncontrolled wing (actuators off), are primarily performed for system identification purpose. The open-loop control operation (actuators on but no feedback) of the wing is implemented and tested, obtaining a stall delay of about 2.8 degrees of AOA. Furthermore, a closed-loop strategy, based on the wing upper surface mean pressure chord-wise distributions signature is adopted to characterize the forthcoming boundary layer detachment. This allows for triggering the controller in stall proximity only, for energy saving purpose. Pertinent results and discussion are provided along with concluding remarks and prospects for future research

Experimental study in near-and far-field of trailing vortices and their active control

Spatialaveraged two-dimensional PIV velocity profiles are compared for ����=7×103 by using direct numerical simulations (DNS) up to eleven chords from the wing. Once we validate our results, we fit the theoretical parameters as function of ����. Five theoretical parameters are given from computational and experimental results: two corresponding to Batchelor’s model and three regarding Moore & Saffman’s model. Two critical Reynolds numbers were found. Our DNS computations verify that the onset of instability of the flow around the wing at the first threshold ������1 ≈1.3×103 captures the change in the trend of theoretical parameters. In addition, the theoretical parameters appear to become constant experimentally for a second critical Reynolds number ������2 greater than 10-20×103 as our results are compared with those given by other authors. Consequently, Reynolds number plays an important role in the stability analysis for trailing vortices not only taking into account viscous terms but also determining the input parameters for theoretical models. Finally, we have carried out a study of the blowing effect of continuous jets that are perpendicular to the moving direction, and blowing from the tip of a NACA0012 airfoil. We analyze three Reynolds numbers ���� and four jet-to-crossflow blowing ratios ��������. We show how these jets are good candidates to reduce the strength of the wingtip vortices at the lowest Reynolds numbers considered, e.g. ���� = 7×103. For higher Reynolds numbers up to ����=20×103, the forcing has a weak influence on the vortex strength in the near-field once the rolling-up process has already finished, and especially at axial distances greater than 7 chords behind the wing. The reason for the presence of two different strength decays depending on the Reynolds number is explained by the ability of the continuous jet to break the vorticity sheet creating a counter-rotating vortex or co-rotating vortex at low or high values of ����, respectively. This mechanism makes the wingtip vortex to decrease or remain its vortex strength as we apply different blowing ratios ��������. This effect is evident at the lowest Reynolds number at which we observe a strong vortex decay. Conversely, the continuous jet changes the characteristics of the vortex flow in the formation and the near-field evolution of the wingtip at high Reynolds numbers, but there is not a appreciable effect on the vortex strength and how downstream evolution.In order to predict the axial development of the wingtip vortices strength, an accurate theoretical model is required. Several experimental techniques have been used to that end, e.g. PIV or hotwire anemometry, but they imply a significant cost and effort. For this reason, we have carried out experiments using the smokewire technique to visualize smoke streaks in six planes perpendicular to the main stream flow direction. Using this visualization technique, we obtained quantitative information regarding the vortex velocity field by means of Batchelor’s model for two chord based Reynolds numbers, ������ = 3.33 ⋅ 104 and 105. Therefore, this theoretical vortex model has been introduced in the integration of ordinary differential equations which describe the temporal evolution of streak lines as a function of two parameters: the swirl number, ��, and the virtual axial origin, ��0. We have applied two different procedures to minimize the distance between experimental and theoretical flow patterns: individual curve fitting at six different control planes in the streamwise direction as well as the global curve fitting which corresponds to all the control planes simultaneously. Both sets of results have been compared with those provided by del Pino et al. [2011a], finding good agreement. Finally, we have observed a weak influence of the Reynolds number on the values �� and ��0 at low-to-moderate ������. This experimental technique is proposed as a low cost alternative to characterize wingtip vortices based on flow visualizations. Secondly, we present a detailed analysis of experimental and numerical results for the flow of wingtip vortices behind a NACA0012 airfoil. Particular attention is paid to a specific value of the angle of attack, ��=9∘, and ultra-low and low chord-based Reynolds numbers ranging from ����=0.3×103 to 20×103

Active and Adaptive Flow Control of Twin-Tail Buffet and Applications

Modern fighter aircraft with dual vertical tails are operated at high angles of attack. The vortex generated by leading edge extension (LEX) breaks down before reaching the two vertical tails. The wake of highly unsteady, turbulent flow causes unbalanced broadband aerodynamic loading on the tails and may produce severe buffet on the tails and lead to tail fatigue failure. Flow suction along the vortex cores (FSVC) is investigated as an active control method for tail-buffet alleviation. Suction tubes have been tilted at different angles to study the control effectiveness of suction tubes orientation. Flow field response, aerodynamic loading and aeroelastic results are compared with the no-control case. These flow modifications produce lower tip bending and rotation angle deflections and accelerations. Moreover, the root bending and twisting moments are reduced in comparison with the no-control case. However, there was no shift in the frequencies at which the peaks of the power spectral density (PSD) responses occurred. The primary effect of the FSVC methods is the amplitude reduction of the aeroelastic responses up to 30%. A parametric investigation is conducted and the best control effectiveness is obtained with the suction tubes tilted at −10°. Next, the twin-tail buffet alleviation is addressed by using adaptive flow control, and an adaptive active control method is developed. Control ports, whose locations are determined according to the locations of a range of high-pressure difference, are placed within a small area on the tail surfaces. Flow suction and blowing are applied through these control ports in order to equalize the pressures on the two surfaces of the tail. Mass flow rate through each port is proportional to the pressure difference across the tail at the location of this port. Comparing the flow field and aeroelastic response with the no-control case, the normal-force and twisting-moment distributions are substantially decreased along with the damping of their amplitudes of variation. The bending-deflection and rotation-angle responses have not changed their sign. The PSD of the root bending moment and root twisting moment have shown substantial decreases of more than 70%. The tail tip acceleration responses have shown similar decreases too. Next, a parallel high-order compact-scheme code (PHCC) is developed to investigate flow control more accurately and more efficiently. The validation cases are presented and compared with theoretical results, experimental results and other computational results. The PHCC results show good accuracy and high efficiency. Flow computational simulations of Jet and Vortex Actuator (JaVA) or synthetic jet have been investigated. The computational results show good agreement with the experimental data and other computational results. Simplified 2D models, which include an airfoil under the effect of JaVAs and synthetic jet actuators, are developed and investigated for control effectiveness. Simulation results show: with properly selected parameters, the oscillating amplitude of pressure difference and normal force acting on airfoil can be reduced, the peak of the normal force PSD can be reduced and the frequencies at which the peaks of the pressure difference PSD responses occurred can be shifted to higher frequency levels. Too low or too high exciting frequencies have no effect or adverse effect. Low exciting velocity may not produce enough disturbances to suppress the pressure oscillation

Experimental Investigation of Active Control of Bluff Body Vortex Shedding

Mean and fluctuating forces acting on a body are strongly related to vortex shedding generated behind it. Therefore, it is possible to obtain substantial reductions of at least the unsteady forces if vortex shedding is controlled or its regularity is reduced. While conventional active flow control methods are mainly concerned with direct interaction with, and alteration of, the mean flow about a body, modern techniques involve altering existing flow instabilities using relatively small inputs to obtain large-scale changes of mean flows. Aerodynamic flow control may be intended to delay or suppress boundary layer separation through creation of a boundary layer downstream from the control input that is able to withstand adverse pressure gradients imposed by the outer (global) flow. In the present work, aerodynamic characteristics of a circular cylinder at Re=156,000 and an axisymmetric body (ogive cylinder) at Re=170,000 are first analyzed using a proposed phase averaging technique for the Particle Image Velocimetry (PIV) data. Later, the effect of plasma actuators on the aerodynamic characteristics of these bodies is investigated. When plasma actuators were placed 10° upstream of the separation point on the circular cylinder, momentum addition, and maybe the effect of local heating, modified the streamwise pressure gradient, leading to the establishment of a thinner boundary layer downstream. Phase synchronization of vortex shedding was also obtained for Re=156,000 for a narrow frequency band of the carrier signal of the actuators when they operated with a 90° phase shift To the knowledge of the author no other method has been shown to achieve vortex shedding control up to this high a Reynolds number. Effects of the different configurations of plasma actuators on the circumference, on the base, and in a streamwise direction were investigated for the ogive cylinder. It was observed that direct alteration of the mean flow about a body was not as effective as the boundary layer flow control where the flow instabilities are exploited. Also, the three dimensionalities in this flow made it significantly more complex to analyze

Computational study of a NACA4415 airfoil using synthetic jet control

textSynthetic jet actuators for flow control applications have been an active topic of experimental research since the 90’s. Numerical simulations have become an important complement of that experimental work, providing detailed information of the dynamics of the controlled flow. This study is part of the AVOCET (Adaptive VOrticity Control Enabled flighT) project and is intended to provide computational support for the design and evaluation of closed-loop flow control with synthetic jet actuators for small scale Unmanned Aerial Vehicles (UAVs). The main objective is to analyze active flow control of a NACA4415 airfoil with tangential synthetic jets via computational modeling. A hybrid Reynolds-Averaged Navier-Stokes/Large Eddy Simulation (RANS/LES) turbulent model (called Delayed Detached-Eddy Simulation-DDES) was implemented in CDP, a kinetic energy conserving Computational Fluid Dynamics (CFD) code. CDP is a parallel unstructured grid incompressible flow solver, developed at the Center for Integrated Turbulence Simulations (CITS) at Stanford University. Two models of synthetic jet actuators have been developed and validated. The first is a detailed model in which the flow in and out of the actuator cavity is modeled. A second less costly model (RSSJ) was also developed in which the Reynolds stress produced by the actuator is modeled, based on information from the detailed model. Several static validation test cases at different angle of attack with modified NACA 4415 and Dragon Eye airfoils were performed. Numerical results show the effects of the actuators on the vortical structure of the flow, as well as on the aerodynamic properties. The main effect of the actuation on the time averaged vorticity field is a bending of the separation shear layer from the actuator toward the airfoil surface, resulting in changes in the aerodynamic properties. Full actuation of the suction side actuator reduces the pitching moment and increases the lift force, while the pressure side actuator increases the pitching moment and reduces the lift force. These observations are in agreement with experimental results. The effectiveness of the actuator is measured by the change in the aerodynamic properties of the airfoil in particular the lift ([Delta]C[subscript t]) and moment ([Delta]C[subscript m]) coefficients. Computational results for the actuator effectiveness show very good agreement with the experimental values (over the range of −2° to 10°). While the actuation modifies the global pressure distribution, the most pronounced effects are near the trailing edge in which a spike in the pressure coefficient (C[subscript p]) is observed. The local reduction of C[subscript p], for both the suction side and pressure side actuators, at x/c = 0.96 (the position of the actuators) is about 0.9 with respect to the unactuated case. This local reduction of the pressure is associated with the trapped vorticity and flow acceleration close to the trailing edge. The RSSJ model is designed to capture the synthetic jet time averaged behavior so that the high actuation frequencies are eliminated. This allows the time step to be increased by a factor of 5. This ad hoc model is also tested in dynamic simulations, in which its capacity to capture the detail model average performance was demonstrated. Finally, the RSSJ model was extended to a different airfoil profile (Dragon Eye) with good results.Mechanical Engineerin