Recent developments in DBD plasma flow control

Progress in Aerospace Sciences 62 (2013) 52–78

Contents lists available at ScienceDirect

Progress in Aerospace Sciences journal homepage: www.elsevier.com/locate/paerosci

Recent developments in DBD plasma flow control Jin-Jun Wang a,n, Kwing-So Choi b, Li-Hao Feng a, Timothy N. Jukes b, Richard D. Whalley b a b

Fluid Mechanics Key Laboratory of Education Ministry, Beijing University of Aeronautics and Astronautics, Beijing 100191, China Faculty of Engineering, University of Nottingham, Nottingham NG7 2RD, UK

art ic l e i nf o

a b s t r a c t

Available online 8 July 2013

Flow control using DBD (dielectric-barrier-discharge) plasma actuators is a relatively new, but rapidly expanding area of research. There are a number of review papers available on this subject, but few discuss on their latest developments. The purpose of the present article is to “fill the gap” by reviewing the recent trend of plasma actuator design and to summarise aerodynamic control techniques. Here, we review new plasma actuators, such as plasma synthetic jet actuators, plasma spark jet actuators, threedimensional plasma actuators and plasma vortex generators, which can induce three-dimensional flows away from the wall. We also review the starting vortex that leads to formation of a plasma wall jet. This is an important subject not only for a better understanding of the flow induced by DBD plasma actuators, but also as a database that can be used to calibrate the numerical models for plasma flow control. Design of DBD plasma actuators to obtain turbulent skin-friction reduction is shown and the modifications to nearwall turbulence structures are summarised. Novel applications of DBD plasma actuators for aerodynamic control are then discussed, including pitch and roll control, plasma jet vectoring, circulation control and plasma flap, showing a potential of DBD plasma actuators for replacing movable, aircraft control surfaces. Finally, vortex shedding control techniques by a number of different plasma actuators are surveyed. & 2013 Elsevier Ltd. All rights reserved.

Keywords: DBD plasma actuator Flow control Aerodynamic control

Contents 1. 2.

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53 Flow induced by DBD plasma . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53 2.1. Starting vortex . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53 2.2. Wall jet . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55 3. Novel plasma actuators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56 3.1. Plasma synthetic jet actuators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56 3.2. Plasma spark jet actuators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58 3.3. Three-dimensional plasma actuators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58 3.4. Plasma vortex generators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61 4. Novel plasma flow control techniques . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62 4.1. Aerodynamic control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62 4.1.1. Pitch and roll control. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62 4.1.2. Plasma jet vectoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63 4.1.3. Circulation control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64 4.1.4. Plasma flap. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64 4.2. Skin-friction reduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65 4.2.1. Plasma spanwise flow oscillation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67 4.2.2. Plasma spanwise travelling waves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68 4.3. Vortex-shedding control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68 5. Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75 Acknowledgements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75 References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75

n

Corresponding author. Tel.: +86 10 82339304. E-mail address: [email protected] (J.J. Wang).

0376-0421/$ - see front matter & 2013 Elsevier Ltd. All rights reserved. http://dx.doi.org/10.1016/j.paerosci.2013.05.003

J.J. Wang et al. / Progress in Aerospace Sciences 62 (2013) 52–78

1. Introduction Dielectric-barrier-discharge (DBD) plasma actuators are all-electric devices without the need of pneumatics, hydraulics or moving parts. They are light and fast reacting, ideal for active control of flow over aerodynamic bodies. Moreover, DBD plasma actuators are very thin, which can be either attached on or integrated into the aerodynamic body. DBD plasma actuators are designed to produce glow discharge plasma on one side (asymmetric actuator) or both sides of the exposed electrode (symmetric actuator), with either an offset or continuous lower electrode. An example of the asymmetric actuator is shown schematically in Fig. 1. Here, plasma is created when sufficient voltage is applied between the electrodes to ionise the surrounding air, which spreads out from the edge of the exposed electrode to the side under which the lower electrode is placed. The discharge is self-limiting, whereby charge emitted from the exposed electrodes gradually accumulates on the dielectric surface and reduces the potential difference. This quenches further plasma formation unless the exposed electrode potential continually increases. These actuators are usually excited with AC voltage to allow continuous plasma generation (albeit intermittently throughout the AC cycle). Momentum is transferred from the plasma discharge to the ambient air through a collision of ions, which results in a net body force transmitted to the air flow. This accelerates the air and induces a flow away from the exposed electrode, directed to the right in Fig. 1. DBD plasma actuators are therefore simple, yet very powerful devices that make novel aerodynamic applications possible. For example, pitch and roll control of aircraft can be made without a movable control surface on the wing; streamwise vortices can be generated without mechanical vortex generators; wall-normal jets can be generated without a pneumatic source, which can be vectored in any direction on demand; lift can be augmented as and when required without slats or flaps attached to the wing. The purpose of the present article is to review these latest DBD plasma actuator designs and novel aerodynamic control techniques. Our review also includes a discussion on the starting vortex that leads to a formation of plasma wall jet. This is an important subject not only for a better understanding of the flow induced by DBD plasma actuators, but also as a database that can be used to calibrate the numerical models for plasma flow control. Design of DBD plasma actuators that were used to obtain turbulent skin-friction reduction is shown and the modifications to near-wall turbulence structures are summarised. In the present review article, we have tried not to duplicate the materials that are already available. For example, we have not reviewed the basic principles of DBD plasma actuators, such as the chemistry and physics of plasma generation, dielectric materials and body force generation. We have also omitted many other flow control applications as well as plasma modelling and other types of plasma for flow control. We have not discussed on the implementation of plasma actuators through closed-loop control under unsteady aerodynamic conditions, either. For those, readers should consult other review articles [1–6], and references therein. Table 1 shows the subjects covered by the present and previous review papers on DBD plasma actuators and flow control.

Fig. 1. Schematic representation of a DBD plasma actuator used by Whalley and Choi [8].

53

Table 1 Subjects covered by the present and previous review papers on DBD plasma actuators and flow control. Here, [P] Present paper; [1] Moreau (2007); [2] Corke et al. (2007); [3] Corke et al. (2010); [4] Jayaraman and Shyy (2008); [5] Touchard (2008); [6] Cho and Shyy (2010). Subjects covered Principle of plasma actuators Plasma chemistry and physics Thrust generation Electrical properties Actuator optimisation Plasma modelling Plasma flow simulation Adaptive plasma flow control Induced flow Starting vortex Wall jet Plasma actuators Single DBD actuators Sliding discharge actuators Travelling wave actuators OAUGDP Plasma synthetic jet actuators Plasma spark jet actuators Bi-directional jet actuators 3-D plasma actuators Plasma vortex generators Aerodynamic flow control Pitch and roll control Plasma jet vectoring Plasma circulation control Plasma flap Flow separation control Skin-friction reduction Vortex shedding control

[1]

[1] [1] [1]

[P] [P]

[1] [1] [1]

[2] [2] [2] [2] [2] [2] [2]

[P] [P] [P] [P] [P] [P] [P] [P] [P] [P] [P] [P]

[4] [4] [4]

[3] [3] [3]

[4] [4]

[6] [6] [6]

[2] [2] [2] [2]

[1] [1] [1]

[3] [3] [3]

[3] [3]

[5] [5] [5] [5] [5] [5] [5]

[5] [1] [1] [1]

[5]

2. Flow induced by DBD plasma 2.1. Starting vortex On initiation of DBD plasma in quiescent air a shear layer is created, which rolls up to form a vortical structure. This starting vortex induced by DBD plasma was first observed by Post [7], who used phase-locked particle image velocimetry (PIV) to observe the transient flow field. Under continual plasma forcing, the starting vortex moved along and away from the wall. Fig. 2a is the smokeflow visualisation carried out by Whalley and Choi [8], showing tightly compacted spirals of smoke circling around the centre of the vortex core. The corresponding PIV vorticity field is given in Fig. 2b. Their proposed mechanism of starting vortex formation and development (see Fig. 3) suggests that the laterally ejected jet flow was replenished by entrainment of fluid from directly above the plasma actuator, which formed a starting vortex. To preserve the no-slip boundary condition, secondary vorticity was generated along the wall, as illustrated in the figure. The secondary vorticity wrapped around the starting vortex and steered its trajectory away from the wall [8–10]. It is also shown that the plasma momentum increased linearly with time once the plasma actuator had reached a steady state velocity U0. This suggests that the DBD plasma actuator entrained and accelerated the surrounding fluid with a constant force. Through high-speed smoke-flow visualisations Whalley and Choi [11] showed that Kelvin–Helmholtz instabilities can be generated in the shear layer of the plasma wall jet, which were entrained and ingested into the core of the starting vortex. Jukes et al. [12] also showed that a symmetric plasma actuator created a pair of counter-rotating starting vortices which moved away from the actuator in opposite directions. Measurements of the temperature distribution of the transient flow field were also conducted with an infrared camera and cold wire anemometry to

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Fig. 2. The starting vortex induced by DBD plasma. (a) Smoke-flow visualisation and (b) PIV vorticity field at t* ¼ 2700 . The white dots indicate the centre of the vortex core in the flow visualisation and PIV data. From Whalley and Choi [8].

Fig. 3. Entrainment by the plasma actuator and the development of the starting vortex: (a) plasma initiation, (b) vortex formation and (c) generation of secondary vorticity. From Whalley and Choi [8].

Fig. 4. Vortex core scalings: (a) xnc versus t* and (b) ync versus t.* From Whalley and Choi [8].

Fig. 5. Self-similarity of the starting vortex at (a) t* ¼ 3000, (b) t* ¼ 4000 and (c) t* ¼ 5000. A black dot in each indicates the location of the centre of the vortex core. From Whalley and Choi [8].

reveal that the starting vortices entrained heated fluid over the upper electrodes into their cores. By operating plasma actuators in an unsteady mode [13,14], a train of vortices were agglomerated to form a single vortex downstream of the plasma actuator [15]. Whalley and Choi [8] showed that the location of the vortex core in the x (wall-parallel) and y (wall-normal) directions scaled with t0.71 as it travelled at 311 to the wall (see Fig. 4). By approximating the measured power law exponent to a close rational number 2/3, they formulated self-similarity variables for

the starting vortex of the form ξ


x 1=3

ν1=3 U 0 t 2=3

;

η


y 1=3

ν1=3 U 0 t 2=3

;

τ


U 1=3

ν1=3 U 0 t −1=3

:

ð1Þ

This showed that the length scale of the starting vortex scaled with t2/3, the velocity scaled with t−1/3 and correspondingly, circulation scaled with t1/3. The similarity laws were verified by plotting the experimental data over a large range of non-dimensional

J.J. Wang et al. / Progress in Aerospace Sciences 62 (2013) 52–78

time t*¼ tU0/ν, as illustrated in Fig. 4. The validity of the selfsimilarity variables in Eq. (1) is clearly demonstrated in Fig. 5 as the location of the centre of the vortex core remained nearly constant at ðξ; ηÞ ¼ ð0:70; 0:52Þ. Mertz and Corke [16] modelled the DBD plasma actuator using the lumped-element circuit model developed by Orlov [17]. This simulated the space–time variation of a DBD plasma actuator by modelling the plasma, surrounding air and the dielectric barrier by a series of diodes, capacitors and resistors, showing that the x-location of the vortex core scaled with t2/3 as found experimentally by Post [7]. In addition, Elam [18] recently performed a direct numerical simulation (DNS) of the DBD plasma actuator in quiescent air to confirm that the length scale of the starting vortex scaled with t2/3 as it travelled at an angle of 321 to the wall. Here, the linear plasma body-force model [4] produced a downwash towards the actuator, thereby correctly representing the entrainment process observed in the experimental investigations. Sattari et al. [19] studied the temporal development of the starting vortex induced by DBD plasma at the trailing edge of a flat plate using time-resolved PIV. Without the influence of the wall, both the location of the vortex core and the strength of the vortex scaled with t1/2. This is in contrast to the results obtained by Whalley and Choi [8]. The difference can be, however, accounted for by the lack of secondary vorticity in this study, which influences the development of the starting vortex. These scaling laws were found to be a function of the thickness and velocity of the developing shear layer.

55

the local maximum jet velocity, Umax, and the jet half-width, δ½ (defined as the wall normal location at which the velocity has reduced to 0.5Umax). This self-similar, steady wall jet was created after some initiation transient during which starting vortices were formed [12,13,28]. The downstream development of the wall jet was quite similar to that predicted by the laminar wall-jet theory [29,30], for which the maximum jet velocity varies as x−1/2 and the jet thickness varies as x3/4. The jet velocity for a single plasma actuator is usually less than 7 m/s at a distance of about 0.5 mm from the wall [1,31]. This limits a typical jet Reynolds number to below 1000, which is low enough to remain laminar [32], although there is evidence of instability vortices within [9,11]. This suggests that natural transition into a turbulent wall jet may occur downstream of an actuator with intense plasma forcing. However, the DBD plasma induced wall jet differs from a classic laminar wall jet because there is no mass added to the flow. Here, the plasma actuator draws ambient fluid towards the wall and then ejects this fluid tangentially away from the electrode (see Fig. 3). This can be seen clearly in the time-averaged PIV of Fig. 7, where fluid is induced into the DBD from just above and upstream of the plasma, and then accelerated laterally along the wall by the body force to form the laminar wall jet. The wall-ward flow is a direct result of continuity since the plasma is a source of momentum, not mass. In this respect, the DBD actuators are similar to synthetic jets [33,34]. Numerous experimental studies were undertaken at the start of this millennium into understanding the relationship between the plasma-induced jet velocity and the plasma excitation parameters

2.2. Wall jet Some of the first measurements of the plasma wall jet induced by a single asymmetric actuator in quiescent air were conducted by Roth et al. [20,21] using a Pitot tube downstream of a plasma actuator. A jet velocity of up to 3.4 m/s was measured within the first few mm of the wall. Johnson and Scott [22] also presented similar jet velocity profiles at several stations downstream of a single symmetric actuator using hot-wire anemometry. Wilkinson [23] described a decelerating, broadening velocity profile in hot-wire measurements, whilst Corke et al. [24] and Post and Corke [25,26] used time-averaged PIV to measure the high-speed jet which was initially parallel to the wall. Both Roth et al. [27] and Jukes et al. [28] compared the induced velocity profile to the theoretical solution of a laminar wall jet [29]. Fig. 6 shows that hot-wire measurements of the velocity profile of the plasma wall jet collapse onto the laminar walljet profile for a number of stations to either side of a single, symmetric actuator. Here, the velocity profile is normalised with

Fig. 7. Time-averaged velocity field induced by a single asymmetric DBD actuator in quiescent air. From Jukes and Choi [35].

Fig. 8. Schematic of a synthetic jet actuator. From Smith and Glezer [43].

Fig. 6. Non-dimensional velocity profile to either side of a symmetric DBD actuator. Measured with a hot-wire probe on both sides of a symmetric actuator. From Jukes et al. [28].

Fig. 9. Schematic of (a) annular and (b) linear plasma synthetic jet actuators. From Santhanakrishnan et al. [47].

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(i.e. voltage, frequency, waveform, etc.) and geometric parameters (i.e. electrode height, length, gap, etc.). Detailed parametric studies have been performed using PIV by Enloe et al. [36], Pitot tubes by Roth and Dai [37], and using Pitot tubes and LDV by Forte et al. [31], where the latter authors indicated that jet velocities could be up to 7 m/s from a single DBD plasma actuator through optimised electrode geometry. Enhanced force production was also possible using thick dielectric material [38]. More recently, the plasmainduced wall jet has been measured using high spatial–temporal resolution PIV. This has enabled the spatial distribution of the body force to be estimated throughout the AC cycle by decomposition of the Navier–Stokes equations [39–42].

3. Novel plasma actuators 3.1. Plasma synthetic jet actuators Fig. 8 shows a mechanical synthetic jet actuator, consisting of a diaphragm (or a piston), a cavity and an orifice. During the

forwards motion of the diaphragm, fluid is ejected from the cavity, forming a vortical structure as shown in the figure. The vortical structure moves further downstream during the backwards motion, and is therefore unaffected by the suction of the diaphragm. Thus, the synthetic jet produces vortex structures periodically during each phase of actuation. Here, the net mass flux from the orifice during one period of actuation is zero, while the net momentum flux is nonzero. The round orifice results in a vortex ring [44,45], while the slot orifice generates a vortex pair [43,46]. Santhanakrishnan and his co-worker originally developed the novel concept of plasma synthetic jets that combined the features of both plasma actuators and synthetic jets [48]. They have conducted a series of experimental and numerical investigations on the characteristics of plasma synthetic jets. Similar to the conventional synthetic jet actuators, there are also two representative types of plasma synthetic jet actuators, namely annular and linear actuators, as shown in Fig. 9a and b, respectively. Annular plasma synthetic jet actuators consist of an annular exposed electrode, embedded electrode and a dielectric sheet, while linear actuators contain two linear exposed electrodes, one or two linear

Fig. 10. Phase-locked vorticity contours overlaid with streamlines for excitation frequency of 1 Hz at (a) t¼ 25 ms and (b) t¼ 94 ms. From Santhanakrishnan and Jacob [48].

Fig. 11. Mean axial velocity distribution of a linear plasma synthetic jet operated with steady actuation in (a) and (c), and with unsteady actuation at 10 Hz in (b) and (d). The x-axis is normalised by the embedded electrode width in (a) and (b) (Santhanakrishnan and Jacob [53]), and by the jet half-width in (c) and (d) (Liu et al. [54]).

J.J. Wang et al. / Progress in Aerospace Sciences 62 (2013) 52–78

embedded electrodes and a dielectric sheet [47,49–53]. For both actuator types, the exposed electrode is located beside the embedded electrode, with little or no gap. The actuator configuration can be easily reversed to act as a suction device [48]. When actuated, the plasma synthetic jet pulls in residual fluid adjacent to the surface and ejects it in the form of a jet. Here, the

Stage 1: Energy Deposition

57

vortical structure is created due to turning of the fluid after suction [48]. Fig. 10 shows the vortex ring evolution during one period of excitation at a frequency of fe ¼1 Hz by using phase-locked PIV. Here, the diameters of the exposed and embedded electrodes were 25.4 mm and 12.7 mm, respectively. The plasma actuator was powered by an AC square wave input with a voltage of 5 kV and

Stage 2: Discharge

Stage 3: Recovery

Fig. 12. Stages of the plasma spark-jet actuator operation cycle. From Grossman et al. [58].

Fig. 13. Time evolution of a plasma spark jet shown by the velocity vector (left) and by the spanwise vorticity (right). From Cybyk et al. [59].

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a frequency of 2.8 kHz at 50% duty cycle. The starting vortex ring is seen centred around x ¼6 mm at t¼25 ms, which is then moved downstream to x ¼30 mm at t ¼94 ms. Fig. 11a and b shows the mean axial velocity distributions of a linear plasma synthetic jet experimentally obtained by Santhanakrishnan and Jacob [53]. Here, the widths of the exposed and embedded electrode were 6.35 mm and 12.7 mm, respectively, and the spanwise length was about 90 mm. The power supply was the same as that previously used (see Fig. 10). Corresponding numerical results were obtained by Liu et al. [54] (see Fig. 11c and d) by solving the Reynolds-averaged Navier–Stokes equations with a phenomenological model for the plasma-induced body force. No fluid suction is observed in the near field of a plasma synthetic jet in a steady operation mode (see Fig. 11a). For an unsteady mode of plasma operation, on the other hand, negative axial velocity is observed near the orifice (Fig. 11b). When the radial location is normalised by the jet half-width b0.5, as in Fig. 11c and d, the axial velocity profiles show a good self-similarity for both steady and unsteady actuations. The normalised velocity distributions are well described by the hyperbolic cosine function u/umax ¼cosh−2(Ay/b0.5), agreeing well with the results of mechanical synthetic jets [43]. Other characteristics of plasma synthetic jet, such as the time evolution of the maximum axial velocity and the jet half-width, as well as the trajectory of the centre of the vortex core have also been studied [47–50,52,53,55]. All of them were found to be similar to those of mechanical synthetic jet actuators [53]. The influence of excitation frequency on the induced flow characteristics was investigated by Zhang et al. [55], while Santhanakrishnan and Jacob [56] studied the effect of plasma morphology on the jet characteristics. Bolitho and Jacob [57] used several plasma synthetic jet actuators in various configurations to determine their flow control potential, demonstrating that the plasma induced velocity and momentum could be increased by combining plasma actuators in series. 3.2. Plasma spark jet actuators The plasma spark-jet actuators were developed over the last decade by a Johns Hopkins University group [58–61]. Following this lead, several other groups, including ONERA and University of Texas at Austin, started to investigate the basic characteristics of plasma spark-jet actuators and its applications [62–67]. Although plasma spark jets are not DBD actuators, these were included in this review due to their importance in actuator design and potential for high-speed flow control. Three electrodes are fashioned in a typical plasma spark jet actuator: an anode, a sharp cathode and a grid. Main discharge current flows from the cathode to the anode, which is initiated by a small cathode-to-grid discharge. The plasma spark-jet operation cycle consists of three stages: energy deposition, discharge and recovery (see Fig. 12). The energy deposition causes a plasma discharge inside the small cavity, which rapidly heats the gas, causing an increase in pressure. The pressurised gas is expelled out of the cavity through a small orifice to produce a jet flow. Fluid is entrained into the cavity during the recovery phase of the cycle to

preserve the conservation of mass. This cycle is repeated to produce a sustained jet. Hence, the plasma spark jet is similar to a synthetic jet, and holds promise of controlling high-speed flows without the need of movable control surfaces. High-resolution PIV measurements were made by Cybyk et al. [59] to examine the characteristics of the plasma spark jet (see Fig. 13). Here, the plasma spark-jet actuator was operated at 0.2 Hz. It had an orifice diameter of 0.33 mm and a chamber volume of 42.28 mm3. The capacitor potential to initiate the spark discharge was 1000 V. At t¼0, the spark is ready to discharge and no flow motion is detected. At time t¼25 μs, well-organised vortical structures were seen near the orifice, which are similar to those induced by a synthetic-jet actuator. Here, the jet velocity of at least 100 m/s was observed, although the actual speed at the jet core was thought to be much higher. In a similar investigation by Emerick et al. [61], the jet front velocity induced by the plasma spark-jet actuator was nearly 310 m/s. As the time evolved to t¼150 μs, more complex vortical structures were convected downstream as the mixing between the jet and the ambient air continued. Caruana et al. [62] used a plasma spark-jet actuator for exhaust noise control at a Mach number between 0.3 and 0.9. The noise control system was made of 6 spark jets directed towards the exhaust jet centreline, impacting an angle of 451. Acoustic measurements showed little influence on noise, since the number of actuators was not sufficient. Since the plasma spark-jet actuator can induce a jet with high velocity close to the speed of sound, it has a great potential for supersonic flow control [61,65,66]. An example is given in Fig. 14, where a plasma spark-jet actuator was tested in a Mach 1.5 crossflow. Here, a Z-type focusing shadowgraph system with two parabolic and two 451 mirrors was used in a wind tunnel for flow visualisation. Fig. 14a is the baseline image without control. The Mach waves can be observed in the images, which are inclined at approximately 431. When the plasma spark-jet actuator was operated in a single-shot mode, a stronger oblique shock was observed at the spark-jet orifice accompanied by an expansion fan. It then developed and propagated along the test section, as shown in Fig. 14b and c. The maximum oblique shock angle reached 481 at around 275 μs. Even when the actuator was operated in a burst mode, the flow features of the plasma spark jet were similar to those in a single-shot mode [61]. Narayanaswamy et al. [67] used a plasma spark-jet actuator to control the shock/boundary-layer interaction generated by a compression ramp at Mach 3. The maximum downstream displacement of the extrapolated separation shock-foot location was about one-quarter of the boundary-layer thickness. Based on the experiments, they suggested that the plasma spark-jet actuator array might be an effective device to shift the shedding frequency of the separated flow to a band that was less likely to cause structural-panel resonance. 3.3. Three-dimensional plasma actuators Roy and his co-workers [68,69] developed new types of plasma actuators with a view to introduce three-dimensional effects on

Fig. 14. Shadowgraph images of the plasma spark-jet array (single shot mode) at Mach 1.5: without control (a), and with control at t ¼150 μs (b) and at t ¼275 μs (c). From Emerick et al. [61].

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the flow field. Schematics of these plasma actuators are shown in Fig. 15a, giving an oblique view of triangular, serpentine and square actuators. Fig. 15b shows the discharges from a serpentine actuator, displaying ionised air along the curved electrode. The pinching and spreading effects of actuators can be observed, showing a three-dimensional swirling flow in the vicinity of the plasma region. The flow fields induced by three-dimensional plasma actuators were numerically studied by Wang et al. [68] by solving the threedimensional drift–diffusion plasma governing equations as well as Navier–Stokes equations. Fig. 16 shows the instantaneous vortical structures that were created by four different types of plasma actuators (see Fig. 15) after 16 ms of operation. While the linear actuator produces very little vorticity, three-dimensional plasma actuators, the serpentine and square actuators in particular, can generate large vorticity downstream through three-dimensional mixing. Based on the stereo-PIV measurements of the induced flow fields, Durscher and Roy [70] concluded that the serpentine configuration combines the characteristics of a plasma synthetic jet actuator and a DBD plasma actuator.

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Fig. 17 shows the effect of a serpentine actuator over a flat plate, where the directions of incoming flow and plasma forcing are indicated by arrows. There was a 2 mm horizontal gap between the exposed and embedded electrodes. The radius of the inner curve was 2 mm, while that of the outer curve was 6 mm (see Fig. 15b). The linear length and the width of the actuator were 38 mm and 2 mm, respectively. Two cases, namely S1 and S2, with different polarities are shown in Fig. 17. The streamwise velocity Vy in the x−y plane in Fig. 17a shows a rather complex flow pattern. Here, the streamwise velocity in case S1 is higher than that in case S2. Fig. 17b shows the distributions of the streamwise velocity Vy on the y−z plane (x¼ 0), while Fig. 17c shows the distribution of wall-normal velocity Vz in the vertical mid-plane (y¼0). These figures show that streamwise vortices created by the serpentine actuators are similar to those produced by plasma vortex generators [71]. Other types of three-dimensional plasma actuators have also been investigated and show similar results to those of the serpentine actuator [69]. Three-dimensional plasma actuators have other useful applications where an increased mixing is desired. For example, Wang

Fig. 15. Schematics of different types of three-dimensional plasma actuators (a) (Wang et al. [68]) and the discharges for a serpentine actuator (b) (Roy and Wang [69]).

Fig. 16. Vortical structures induced by three-dimensional plasma actuators after 16 ms of operation. Red and blue colours denote the positive and negative vorticities with 7 1000 s−1. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.) From Wang et al. [68].

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Fig. 17. Velocity contours of flow induced by the serpentine actuators in a freestream. (a) Streamwise velocity Vy at z ¼1 mm; (b) streamwise velocity Vy at x ¼0; and (c) wall-normal velocity Vz at y¼0. From Roy and Wang [69].

Fig. 18. Counter-rotating (left) and co-rotating (right) plasma vortex generator arrays and application on the suction surface of an aerofoil (flow direction upwards from line LE). From Jukes et al. [76].

and Roy [72] used the serpentine actuators for combustion stabilisation. Here, two different configurations of plasma actuators were employed. One was a counter-flow serpentine actuator,

where plasma forcing was opposite to the incoming flow direction. The other was a co-flow serpentine actuator, whose forcing was in the same direction as that of the incoming flow. Numerical

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simulation results indicated that both co- and counter-flow serpentine actuators could significantly influence the combustion process by creating large recirculation zones, which enhanced the mixing of fuel with the surrounding air. 3.4. Plasma vortex generators When yawed to the incoming flow at some angle, DBD plasma actuators create streamwise vortices which can aid mixing between the boundary layer and the freestream [71]. These DBD plasma actuators are called plasma vortex generators. In contrast to conventional DBD actuators, plasma vortex generators re-energise the boundary layer by using three-dimensional vortex action to induce high momentum fluid from the outer flow into the near-wall region. This concept is identical to that of mechanical vortex generators [73,74], which have proven particularly useful for flow separation control because streamwise vortices have remarkable longevity. While mechanical vortex generators have an associated pressure drag penalty, plasma vortex generators can be simply energised to form streamwise vortices as and when required without adding device drag. Plasma vortex generators typically consist of rows of actuators placed along the freestream direction, creating rows of streamwise vortices. Each actuator can be of asymmetric geometry, so that the plasma jets are all in the same direction to produce co-rotating (CoR) streamwise vortices. Counter-rotating (CtR) vortices can be created by an array of symmetric plasma actuators with a common lower electrode. These geometries and the vortex structures are illustrated in Fig. 18. The plasma vortex generator concept can be traced back to studies by Roth et al. [20,21], where symmetric-type DBD actuators were oriented along the flow direction of a laminar boundary layer. Smoke-wire flow visualisation near the wall at U¼4 m/s revealed unstable streamwise vortical structure, which broke down close to the actuator tips. These studies were aimed at reducing skin-friction drag and the streamwise actuators were counterproductive to this goal due to transition. One of the first studies of plasma vortex generators for flow separation control was conducted by Okita et al. [75]. Here, a single, short (l/c¼ 0.25) asymmetric DBD actuator was set at several

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yaw angles at x/c¼ 0.3 on the suction surface of a NACA 0024 at Re¼104. Flow visualisation indicated that a streamwise vortex was created for all yaw angles except when the electrode was perpendicular to the flow (co-flow forcing). PIV showed that this actuator was very effective in suppressing flow separation when placed at 30 degree yaw angle close to the separation point. A detailed study of plasma vortex generators was conducted by Jukes and Choi [35], who used cross-stream PIV to study a single asymmetric plasma actuator in a flat plate laminar boundary layer at Rex ¼9−70  104. The vortex development is shown in Fig. 19, where the plasma actuator created a spanwise directed wall jet whose outer layer rolls up through interaction with the cross-flow to create a stable streamwise vortex structure. A parametric study demonstrated that the vortex circulation increased with the plasma jet-to-freestream velocity ratio (Wp/U∞) and actuator length, and reached a maximum when the plasma actuator was streamwise oriented (901 yaw angle). Optimal co- and counter-rotating spanwise spacing was suggested by minimising the distance between vortices but maintaining sufficient spacing so that they did not interact. This, however, depended on Wp/ U∞, which is a key parameter for these devices as it dictates the vortex size and strength. Plasma vortex generators were tested for wind turbine application by Jukes et al. [71,76]. Here, co- and counter-rotating DBD-VG arrays were implemented using rows of asymmetric and symmetric streamwise-oriented plasma actuators, respectively, and attached to the fore section of the suction surface of the aerofoil. Flow reattachment was demonstrated at both Re¼ 3.5  104 and 9.5  104 with Wp/U∞ ¼ 0.14 and 0.09, respectively, where drag was reduced by up to 63% and lift increased by 67%. Flow separation control can be clearly seen in Fig. 20. Counter-rotating plasma vortex generators produced the largest effect along the line of the plasma actuator (downwash region), although co-rotating vortex generators produced more homogeneous flow control across the span due to vortex translation in the z direction. A similar result was also documented over a backwards facing ramp model by Jukes and Choi [35], as shown in Fig. 21, where flow separation control was still possible with Wp/U∞ as low as 0.07. Furthermore, Jukes et al. [71,76] demonstrate that the

Fig. 19. Orthographic projection and 3D representation of the streamwise vortex generated by a single, streamwise oriented asymmetric plasma actuator in a laminar boundary layer at Rex ¼ 90  103 and Wp/U∞ ¼ 0.85. From Jukes and Choi [35] at Rex ¼ 90  103 and Wp/U∞ ¼0.85.

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Fig. 20. Flow separation control on a NACA 4418 with counter-rotating plasma vortex generators at Re¼9.5  104 and Wp/U∞ ¼ 0.09. From Jukes et al. [71].

Fig. 22. Schematic of roll moment produced by operating plasma actuators. From Vorobiev et al. [86].

Fig. 21. Flow control over a backwards facing ramp using co-rotating (top) and counter-rotating (bottom) plasma vortex generators at Rex ¼ 9.0  104 and Wp/U∞ ¼ 0.53. From Jukes and Choi [35].

plasma vortex generators on the aerofoil suction surface in the region 5%ox/co34% could outperform a pair of co-flow actuators operated simultaneously and with identical power input at x/c¼ 5% and 34%. Huang et al. [77], Chan et al. [78] and Huang and Zhang [79] used plasma vortex generators to control aerodynamic noise in a l/d ¼1 cavity similar to a landing gear or weapons bay at U¼20 m/ s, Re¼7.2  104, Wp/U∞≈0.1. Using microphone and PIV measurements they concluded that plasma vortex generators placed just upstream of the cavity were more effective than the conventional plasma actuators for noise attenuation. This was because threedimensional vortical structures were produced in the shear layer which impeded the development of organised structures in the cavity [79]. This disrupted the feedback mechanism that was necessary to sustain the fluid–acoustic process.

He et al. [80] carried out turbulent boundary-layer separation control over a ramp model using counter-rotating (CtR) plasma vortex generators at U¼10 m/s, Rec ¼ 2.88  105 (unknown Wp/U∞). Smoke visualisation revealed vortical structures but the performance of plasma vortex generators was not as good as that of the conventional plasma actuators, presumably due to unoptimised actuator location and forcing strength. In contrast, Schatzman and Thomas [81] observed that plasma vortex generators could be just as effective as the conventional plasma actuators at reducing turbulent flow separation in a similar configuration. Spanwise variations in the boundary layer thickness using PIV and LDV at U¼5 m/s, Wp/U∞≈0.7 indicated that vortical motions were created within the boundary layer which promoted mixing between high and low momentum fluid. Grundmann et al. [82] used a plasma vortex generator to modify the pressure recovery in a 3D diffuser at Re ¼8−15  103, Wp/U∞≈0.3 to obtain a better result than by a conventional plasma actuator. Streamwise vortices in the inlet section of the diffuser interacted with the secondary flows and redistributed the velocities in the inlet plane. The performance of plasma vortex generator could be either improved or degraded depending on the actuator modulation frequency, duty cycle, and drive voltage. Plasma vortex generators have also been used to control boundary layer transition. Hanson et al. [83,84] used symmetric plasma actuators to generate counter-rotating vortices in a laminar boundary layer with cylindrical roughness elements upstream. Hot-wire measurements at U¼5 m/s demonstrated that plasma vortex generators were able to control the transient growth of the disturbances to delay bypass transition.

4. Novel plasma flow control techniques 4.1. Aerodynamic control 4.1.1. Pitch and roll control Plasma actuators have been used to control flow around aerofoils, wings and aircrafts to improve their aerodynamic performance.

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Fig. 23. Lift and roll control by a combination of plasma actuators. From Vorobiev et al. [86].

When aerodynamic forces were controlled by plasma actuators, the pitch and roll were also affected. Thus, Post and Corke [85] were able to show the changing pitching moment of an oscillating aerofoil by using plasma actuators placed at the leading edge. Plasma actuators can also be used for roll control of aircrafts, as demonstrated by Vorobiev et al. [86]. Here, two plasma actuators were recessed into the suction side of the wing, as shown in Fig. 22, without steps or gaps. The actuators were situated at x/ c ¼0.75 with a streamwise extent of approximately 40 mm, which were driven by a positive saw-tooth signal with a peak voltage of 16 kV at a frequency of 2.9 kHz. Roll control was achieved by operating one of the actuators. A typical change in roll moment coefficient on the wing is given in Fig. 23, showing that each plasma actuator produced an equal but opposite roll moment coefficient when operated individually. When both actuators were turned on, only the lift coefficient was increased. It was found that the equivalent flap angle can be as much as 31 at a freestream velocity of 2 m/s. Similarly, Wei et al. [87] placed plasma actuators on both wings for aerodynamic control, showing that even a small plasmainduced roll moment could satisfactorily achieve manoeuvring tasks. He et al. [88] examined a different plasma configuration, where one actuator was located on the suction side while the other was on the pressure side. When the plasma actuators were placed along the half wing span, the roll moment coefficient produced was equivalent to that by an aileron with 2.51 deflection. When the plasma actuators were placed along the full wing span, the equivalent aileron deflection increased to nearly 91. The leading-edge vortex is a source of lift for delta wings [89]; thus controlling vortical structures over delta wings can modify aerodynamic forces and moments. Zhang et al. [90] showed that leadingedge vortex breakdown of a delta wing can be moved further downstream by a plasma actuator, resulting in an increase in both lift and lift-to-drag ratio. A potential application of plasma actuators for roll control of delta wings was proposed by Budovsky et al. [91], who conducted an experiment where only one side of the leadingedge vortex was controlled. Patel et al. [92] used plasma actuators to control a 471 1303 unmanned air vehicle (UAV). Here, the plasma control was implemented at the leading edge of the wing to provide longitudinal control. Based on the experimental results, they deduced the technical feasibility of a plasma flap concept for hingeless flight control of air vehicles, in particular for vehicles with highly swept wings and at high angles of attack flight conditions where conventional flaps and ailerons are ineffective.

Fig. 24. Schematic of a UAV model mounted on a sting to demonstrate roll control by plasma actuators. From Nelson et al. [93].

Nelson et al. [93] conducted a series of experiments to evaluate the effectiveness of plasma actuators for controlling the 1303 UAV model at high angles of attack. The model was mounted on a sting that was held with low-friction roller bearings to allow roll motion, as shown in Fig. 24. The plasma actuators were aligned parallel to the leading edge by covering 90% of the wing span, with the exposed electrode on the lower side of the model wing. The voltage amplitude was 7.5 kV and the excitation frequency was 5 kHz. Fig. 25 shows the rolling angle of the model when the plasma actuators were turned on. When both actuators were simultaneously turned on or off, the model did not roll. When only one actuator was turned on, the model rolled rapidly to a trim angle at −401 or +401. Based on the force measurements, it was found that the plasma actuators were as effective as the conventional aileron for roll control.

4.1.2. Plasma jet vectoring Traditional plasma actuators affect only the near-wall region of the boundary-layer flow; thus their application may be limited to flow control along the wall surface. Vectoring of plasma jet at an angle to the wall can be achieved, however, by a pair of plasma actuators similar to the linear plasma synthetic jet actuator, as shown in Fig. 9b. Here, the input signal, such as the operating voltage, frequency and duty cycle, for each of the two exposed electrodes should be differentiated [94–96]. Flow fields of vectored plasma jet as a result of differential voltage, frequency and duty cycle are shown in Fig. 26a, b and c, respectively. Here, the straight line denotes the centreline of the vectoring jet. By adjusting the input parameter of either of exposed electrodes, the vectoring angle could be adjusted. Here, even a full 1801 of vectoring angle was possible. The relationship between the angle of the vectoring jet and the voltage, frequency and duty cycle differential was approximately linear. Possible applications of jet vectoring in cross-flow include virtual aero-shaping and vortex generators.

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Fig. 25. Response of free-to-roll model to the left and the right plasma actuators at Re¼ 4.12  105 and a¼ 201. From Nelson et al. [93].

Benard et al. [97] used plasma jet vectoring to carry out flow separation control of a NACA 0015 aerofoil at a freestream velocity of 15 m/s (see Fig. 27). The aerofoil had a 200 mm chord with a 296 mm span, where the plasma actuator was flush mounted. The upstream exposed electrode, 10 mm wide, was placed at 2.5% chord from the leading edge, while the downstream exposed electrode was placed at 24.5% chord. The plasma jet vectoring was achieved by adjusting the difference in the voltage supply to these electrodes. The natural stall angle of the aerofoil was increased by 2.51 by a wall-normal jet, as shown in Fig. 27a. A similar increase in the stall angle was observed by a jet with 181 vectoring angle, Fig. 27b, and with −181 vectoring angle, Fig. 27c. A tangential co-flow plasma actuation increased the lift performance at the most by 3.01, as shown in Fig. 27d, while the control using a counter-flow mode had little effect (Fig. 27e). Therefore, the directional plasma jet was no more effective than a conventional co-flow plasma actuator in a quasi-steady mode. However, Benard et al. [97] argued that an increase in the directional plasma jet performances was expected in an unsteady actuation mode because it can produce counter-rotating vortices to enhance the momentum mixing significantly.

4.1.3. Circulation control The Coanda effect in circulation control keeps the tangential jet attached over a curved surface [99], as shown in Fig. 28a. Here, the

lift performance of aerofoils and wings can be enhanced by a simple addition of a wall jet. However, only two circulation control aircrafts have ever been built and flight tested due to a number of technical issues associated with the blowing air supply [100]. Indeed, the pneumatic power required for circulation control increases as the square of the take-off velocity [101]. In an attempt to overcome some of these problems in conventional circulation control technique, Zhang et al. [98] proposed a novel plasma circulation control concept. In their numerical simulation, a DBD plasma actuator was attached to the blunt trailing edge of an NCCR 1510-7067N aerofoil, as shown in Fig. 28b. The 15% thick aerofoil with an elliptic section had 1% camber. The aerofoil chord length and the freestream velocity were 1 m and 10 m/s, respectively, to give the chord Reynolds number of 6.84  105. The plasma actuator was simulated with a phenomenological model, by adding the body force term to the conservation equations for momentum and energy. The plasma wall jet remained attached to the trailing edge of the aerofoil due to the Coanda effect, drawing the air flow over the trailing edge towards the lower surface, as shown in Fig. 29. This created suction peaks at the leading and trailing edges of the aerofoil, while the pressure over the lower surface was increased, as shown in Fig. 30. Thus, the pressure difference between the upper and lower surfaces of the aerofoil was increased, resulting in an increase in the lift coefficient. Such a lift increment mechanism is similar to conventional circulation control techniques. Fig. 31 shows the aerodynamic characteristics of an NCCR 15107067N aerofoil with and without plasma circulation control. When the plasma was actuated, aerodynamic performance was dramatically enhanced for all angles of attack tested. For example, the lift coefficient increased from 0.48 to 1.30, while the maximum lift-todrag ratio increased from 29.7 to 41.2, giving an increment of 38.7% at an angle of attack α¼ 41. Here, the lift augmentation efficiency of this plasma circulation control was CL/Cμ ¼81.8. After the optimisation of the plasma actuator location, the lift-to-drag ratio increased to CL/Cμ ¼134.9, which was much higher than the maximum lift augmentation efficiency of 80 for the conventional circulation control [102].

4.1.4. Plasma flap He et al. [88] used plasma actuators to control flow separation over a NACA 0015 aerofoil in a manner that could potentially replace the leading-edge slat and trailing-edge flap. The plasma actuator uniformly increased the lift coefficient over all angles of attack when it was placed on the suction surface near the trailing edge. Similar lift enhancement was also observed by Corke et al. [24] and Little and Samimy [103]. When the plasma actuator was placed near the leading edge, however, the lift coefficient was increased only after stall [1,2,104,105]. Such behaviour was similar to that of a mechanical trailing-edge flap [106]; thus the conception of a “plasma flap” was proposed [88]. The plasma flap gave the same amount of shift in zero-angle lift coefficient as that of a mechanical flap with a deflection angle of 1.51. This is within the range of aileron deflections at cruise. The Gurney flap is a small and simple device that can increase the lift coefficient of aerofoils and wings, thereby shortening the take-off distance of aircraft. It is easy to attach onto the pressure side of an aerofoil near its trailing edge, as shown in Fig. 32a. The size of a Gurney flap is typically 0.5–2% of the chord length of the aerofoil [106]. Zhang et al. [107] carried out a numerical simulation of a plasma-based Gurney flap. Unlike the mechanical Gurney flap, the plasma Gurney flap does not incur a form drag; therefore there will be no drag penalty during cruise. The aerofoil chord length and the freestream velocity were 1 m and 10 m/s, respectively, giving the chord Reynolds number of 6.84  105. The trailing edge

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Fig. 26. Velocity distribution of vectoring jet due to (a) 0.38 kV voltage difference (Porter et al. [95]); (b) 750 Hz frequency difference, and (c) 25% duty cycle difference (Fleming et al. [96]).

of the aerofoil was 3 mm thick, where a plasma actuator was attached to form a wall jet, as shown in Fig. 32b. The pressure distribution over the aerofoil at angle of attack α¼ 81 with and without a plasma Gurney flap is shown in Fig. 33. It shows that the pressure over the upper surface was reduced by the plasma control, while it was increased over the lower surface. Thus, the pressure difference between the upper and lower surfaces increased with the deployment of a plasma Gurney flap. The change in the pressure distribution over the aerofoil with a plasma Gurney flap was similar to that with a mechanical Gurney flap. Fig. 34 shows the aerodynamic characteristics of a NACA 0012 aerofoil with and without plasma Gurney flap. Fig. 34a indicates that the linear part of the increased lift curve was parallel to that of the clean aerofoil. Meanwhile, Fig. 34b demonstrates that the drag coefficient with a plasma Gurney flap was less than that of a clean aerofoil at the same lift coefficient before stall. The lift-to-drag ratio and nose-down pitching moment were also increased by plasma control, as shown in Figs. 34c and 34d, respectively. Thus, it was suggested that a plasma Gurney flap has the same functionality as that of a mechanical Gurney flap but without drag penalty. Feng et al. [108] attached a plasma actuator on the downstream surface of a mechanical Gurney flap with a view to further enhance the lift force. The test model used was a NACA 0012 aerofoil with a 100 mm chord and a 250 mm span. The freestream velocity was 3.0 m/s, corresponding to the chord Reynolds number of Re¼ 2  104. Three different mechanical Gurney flaps of height 3.0%c, 4.5%c and 7.0%c incorporating plasma actuators were tested with corresponding momentum coefficients of Cμ ¼ 0.11%, 1.15%, and 1.39%, respectively. A dynamic force balance was used to measure the lift and drag forces, while a time-resolved PIV system was employed to obtain the velocity field in the near-wake region. Fig. 35a shows that the lift coefficient was increased by increasing the mechanical Gurney flap height. The plasma actuators

further increased the lift coefficient over the mechanical Gurney flaps alone during the entire angles of attack. Feng et al. [108] concluded that the plasma forcing with Cμ ¼1% had an effective Gurney flap height increment of h/c¼ 1%. The drag coefficient was increased by the mechanical Gurney flap, however, which was further increased by plasma control, as shown in Fig. 35b. The time-averaged streamwise velocity distribution is shown in Fig. 36. It indicates that the mechanical Gurney flap shifted the wake region downwards, while the plasma forcing on the Gurney flap shifted the wake region further downwards. This suggests that the plasma actuator acted to increase the virtual height of the Gurney flap without physically extending the flap. Analysis of flow physics in the aerofoil wake based on the PIV measurements indicated that the recirculation region behind the Gurney flap became shorter and narrower with plasma forcing, leading to stronger wake vortices. The enhanced entrainment due to the additional plasma forcing between upper and lower wake vortices can be clearly seen in Fig. 37. It indicates that with plasma control both upper and lower wake shear layers have been entrained towards the plasma actuator, and they come much closer to the Gurney flap. During the formation process, the wake vortices with plasma control moved to a much lower position, suggesting that the plasma forcing increased the equivalent camber of the aerofoil in a similar manner to that of the mechanical Gurney flap.

4.2. Skin-friction reduction Spanwise flow oscillation is one of the most effective techniques in turbulent boundary-layer control [109], where as much as 45% reductions in skin-friction drag can be achieved from this open-loop technique. This can be achieved by a number of different techniques, such as spanwise-wall oscillation [110], electromagnetic oscillation [111] and local oscillatory blowing

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Fig. 27. Lift coefficient of NACA 0015 aerofoil, where arrows indicate the beginning of stall. (a) Wall-normal mode of plasma jet operation (20–20 kV applied to both of the exposed electrodes); (b) directional mode with jet vectoring angle of 181 to the vertical axis (20–16 kV); (c) jet vectoring angle of −181 (16–20 kV); (d) tangential co-flow wall jet (20–0 kV); and (e) tangential counter-flow wall jet (0–20 kV). From Benard et al. [97].

Fig. 28. Circulation control of an aerofoil by wall jet (a) and by plasma actuator (b). From Zhang et al. [98].

[112]. All of these methods result in a similar amount of drag reduction, suggesting that the drag reduction mechanism may be associated with the interaction of near-wall turbulence structures with the Stokes layer being created by the spanwise oscillation

(see Fig. 38a). Further discussions on the spanwise flow oscillation technique can be found in other papers [113–115]. Similarly, spanwise travelling waves can be applied to turbulent wall flows to achieve skin-friction reductions up to 30%. In their

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direct numerical simulation (DNS), Du and Karniadakis [116] applied a Lorentz force of the following form:   2π 2π t ; ð2Þ F z ¼ Ie−y=Δ sin z− λz T where I is the forcing amplitude, Δ is the penetration depth, λz is the spanwise wavelength, and T is the period of the travelling wave. Spanwise flow oscillation is a special case of spanwise travelling waves as λz-∞. Du et al. [117] found that the application of the spanwise travelling waves led to a large change in the nearwall turbulence structure. Here, the spanwise travelling wave amalgamates the low-speed streaks into a wide ribbon of lowspeed fluid, which is spread over the wall (see Fig. 38b). Spanwise travelling waves can be realised by Lorentz forcing [118,119] and out-of-plane motion of a flexible wall [120–122].

Fig. 29. Velocity vectors near the aerofoil trailing edge with plasma circulation control. From Zhang et al. [98].

Fig. 30. Pressure distribution over an NCCR 1510-7067N aerofoil with and without plasma circulation control at α¼ 01. From Zhang et al. [98].

4.2.1. Plasma spanwise flow oscillation Wilkinson [23] used DBD plasma actuators to generate oscillatory flow over a surface in quiescent air with an intention of producing a spanwise oscillation to reduce skin-friction drag. A single upper electrode was flanked by two lower electrodes, and plasma was created on either side of the upper electrode by frequency modulating the input signals to the lower electrodes. However, this approach caused problems because there were regions within the oscillation cycle where both sides of the electrodes formed plasma (one side increasing in strength whilst the other side diminishing). Hot-wire measurements showed that this created a flow away from the wall between adjacent actuators, therefore unsuitable for generating spanwise oscillation. Jukes et al. [14] used two sets of interlaced asymmetric actuators and a two channel power supply to create the oscillatory flow. The actuators are drawn schematically in Fig. 39, where two upper electrode sets are located on either side of a common grounded electrode. Alternately activating each electrode sets produced oscillatory plasma jets near the wall but unlike Wilkinson's [23], this strategy avoided the formation of wall-normal jets due to the discreet nature of the excitation. However, the actuator configuration was similar to co-rotating plasma vortex generators, so that rows of co-rotating streamwise vortices were created near the wall, although the vortex location and rotational direction changed in each oscillation cycle, as illustrated in Fig. 40. Jukes et al. [14] and Jukes [15] experimentally studied spanwise oscillation plasma actuators in a turbulent boundary layer at Reτ ¼δuτ/ν ¼400. Hot-wire and cold-wire anemometry was used to measure the boundary layer structure, where the change in near-wall velocity gradient suggested a drag reduction of up to 45% for the oscillation period, T+ ¼Tuτ2/ν¼16, plasma jet velocity, W+ ¼W/uτ≈10 and electrode spacing s+ ¼suτ/ν¼20. Large modifications in the boundary layer structure associated with the spanwise flow oscillation were also reported, where the

Fig. 31. Aerodynamic forces on an NCCR 1510-7067N aerofoil with and without plasma circulation control. (a) Lift coefficient and (b) lift-to-drag ratio. From Zhang et al. [98].

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Fig. 32. Sketch of the mechanical Gurney flap (a) and the plasma Gurney flap (b). From Zhang et al. [107].

Fig. 33. Pressure distribution over a NACA 0012 aerofoil with and without a plasma Gurney flap at α¼ 81. From Zhang et al. [107].

turbulence intensity reduced near the wall and there was a significant decrease in intensity and duration of sweep events [123]. It was shown that a spanwise-oscillating wall jet was created in addition to the streamwise vortices, as illustrated in Fig. 41. The jets appeared to be disrupting the near wall turbulence activity and reducing the velocity gradient in the viscous sublayer. The relative importance and relationship between the wall jets and streamwise vortices was not clear. Elam [18] numerically verified this approach using DNS at Reτ ¼ 200, and also observed a drag reduction of more than 40%. Like Jukes [15], the drag reduction occurred only when the plasma actuators had small spanwise spacing (s+≤20), so that it appears that this parameter is crucial for obtaining skin-friction reduction.

4.2.2. Plasma spanwise travelling waves Whalley and Choi [124] performed an investigation into the use of DBD plasma actuators to generate spanwise travelling waves as an attempt to obtain skin-friction drag reduction in turbulent wall flows. A schematic representation of the travelling-wave array used in their experiments is shown in Fig. 42a, which consists of four plasma actuators with wavelength λ ¼100 mm (λ+ ¼500). To create the travelling waves the plasma actuators were operated in sequence from phase (i) to phase (iv) for a duration of T/4: see Fig. 42b. Over the four phases of actuation the plasma forcing effectively moved fluid across a wavelength of the travelling wave. The boundary-layer measurements were performed at a freestream velocity of U∞ ¼1.8 m/s, in a boundary layer with thickness of δ¼90 mm, giving a Reynolds number of Reτ ¼ δuτ/ν ¼435. Changes in the near-wall turbulence structure with plasma spanwise travelling waves are shown in Fig. 43a. These data have been obtained with particle image velocimetry (PIV) in the x−z plane at y+ ¼ 5 and were phase-averaged over 51 forcing periods.

The location of plasma forcing is z+ ¼225 (shown on the side of the image), which generated a wide ribbon of low-speed streamwise velocity spanning 275 oz+ o375. The generation of a wide ribbon of low-speed streamwise velocity with the plasma spanwise travelling wave was similar to the observation of Du et al. [116], Fig. 38. However, a noticeable difference was the high-speed streamwise velocity region around z+ ¼200, Fig. 43a, which was formed due to the downwash associated with plasma actuators [9,124]. It was also observed that quasi-streamwise vortices in the nearwall region were amalgamated by spanwise-travelling waves to form a single streamwise vortex, which transported low-speed fluid in the spanwise direction [124]. Fig. 44a shows how this streamwise vortex generates the Reynolds stress within the turbulent boundary layer. Here, the data in the x–z plane at y+ ¼5 have been obtained with 2D PIV, while the data in the z–y plane have been obtained with stereoscopic PIV. The streamwise vortex can be seen to the right of the image by the circling arrows of V- and W-components of velocity, where the location of the operating plasma actuator is shown under the image. The streamwise vortex collected the low-speed fluid within the near-wall region and formed a wide band of low-speed fluid, as illustrated by the −uw+ Reynolds stress in the x–z plane of Fig. 44a. At the same time, the streamwise vortex lifted the low-speed fluid outwards into the outer-boundary layer by vortex induction. The upwash side of the streamwise vortex caused an ejection event (a Q-II event), which created the −uv+ Reynolds stress in the z−y plane of the boundary layer. The downwash side of the streamwise vortex created an inward-wall event (Q-III event) as the low-speed fluid lifted around the periphery of the streamwise vortex was entrained back into the wall. Close to the location of the plasma actuator, fluid is entrained into the wall, which caused a sweep event (Q-IV event) and produced a negative band of the −uw+ Reynolds stress. There are strong similarities between the spanwise travelling waves created by plasma and Lorentz forcing. To illustrate these similarities, the production of Reynolds stress by spanwise travelling waves generated by Lorentz forcing is shown in Fig. 44b. This has been obtained by DNS in a turbulent channel flow [125]. The distributions of the −uv+ Reynolds stress and fluctuating streamwise velocity show remarkable resemblance with those by plasma spanwise travelling waves (see Fig. 44a). The streamwise vortex generated by Lorentz forcing also created Q-II and Q-III events, which contributed to the formation of the wide ribbons of lowspeed fluid within the viscous sublayer. 4.3. Vortex-shedding control The wake behind bluff bodies offers excellent opportunities to study flow control actuators since they offer a range of fluid dynamics phenomena. The flow around circular cylinders is perhaps the most fundamental configuration which exhibits boundary layer transition, flow separation, free shear layers and vortex shedding. It is no wonder that the circular cylinder has become a very useful test bed, especially since the fluid mechanics

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Fig. 34. Aerodynamic forces and moment on a NACA 0012 aerofoil with and without a plasma Gurney flap. (a) Lift coefficient; (b) lift coefficient versus drag coefficient; (c) lift-to-drag ratio; and (d) nose-down pitching moment coefficient. From Zhang et al. [107].

Fig. 35. Aerodynamic forces of the aerofoil installed Gurney flap with and without plasma control. (a) Lift coefficient and (b) drag coefficient. From Feng et al. [108].

phenomenon is well documented for a wide range of the Reynolds number [126–128]. Flow control is usually applied in order to enhance or eliminate the vortex shedding, so that wake mixing, fluctuating forces and aerodynamic noise are controlled. Asghar and Jumper [129] used asymmetric co-flow plasma actuators to synchronise the vortex shedding between two side by side cylinders at Re¼8−41  103 with a spanwise spacing of 4 diameters. Their interest was to introduce periodic disturbances in the flow upstream of a linear compressor cascade, which simulated unsteady flow into a turbofan engine. However, the vortex shedding between side-by-side cylinders upstream was often out of phase, but this could be rectified by using plasma actuators at the sides of the cylinder. Cross-correlation of hot-wire anemometry signals in the wake showed phase synchronization of the vortex shedding when using phase-locked plasma excitation on both cylinders. However, little or no synchronization could be achieved when the actuators were placed away from the sides. McLaughlin et al. [130] took recommendations from Asghar and Jumper [129] and mounted a pair of DBD plasma actuators at 7901 from the front stagnation point on a single circular cylinder at Re¼7.4  103. Both actuators produced forcing with the flow and spanned the length of the cylinder. Anti-phase forcing was

applied between opposite sides of the cylinder by phasemodulating the power supply and an array of hot-film probes was used to study the near wake. The vortex shedding could lock on at the plasma forcing frequency, provided the forcing frequency was similar to the natural shedding frequency (i.e. f+ ¼ fd/U∞≈0.21). Here, f is the forcing frequency, d is the cylinder diameter, and U∞ is the cross-flow velocity. The frequency band for which lock-on occurred increased with applied voltage (equivalent to increased momentum) up to 0.1≤f+≤0.3. This study was expanded by Munska and McLaughlin [131], where hot-film probes were used three diameters downstream of the cylinder to show that the vortex shedding had increased two-dimensionality with DBD actuators. Furthermore they found that impulsive, burst forcing was more effective than phase-modulated forcing, and suggested that suddenly activating the plasma would provide an effective trigger to start vortex shedding. Vortex shedding lock-on was possible up to Re ¼8.8  104 in their study. Sung et al. [132] used flow visualisation and PIV with steady DBD plasma forcing simultaneously at 7901, 71201 and 7 1501 on a cylinder at Re¼1.8  104. This showed that the flow could be reattached nearly to the rear stagnation point with co-flow actuators whilst flow separation could be advanced to about

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Fig. 36. Distribution of the time-averaged streamwise velocity at angle of attack α¼ 21 for a clean aerofoil and the aerofoil with a 7.0%c height Gurney flap, with and without plasma control. From Feng et al. [108].

Fig. 37. Phase-averaged spanwise vorticity field ωzc/U∞ superposed with velocity vector (with 0.8U∞ subtracted from streamwise velocity) for the aerofoil with a 7.0%c height Gurney flap, without (a) and with (b) plasma control at α¼ 21. From Feng et al. [108].

7601 with counter-flow actuators. Steady co-flow forcing at 7901, only, could narrow the wake and reduce the wake momentum deficit. Several studies have been undertaken by Thomas et al. [133– 135] and Kozlov and Thomas [136,137], with an intent of reducing the vortex shedding around aircraft landing gear to reduce aerodynamic noise. In Thomas et al. [135], plasma actuators were placed at 7901 and 71351 from the front stagnation point and studied with smoke visualisation, hot-wire wake measurements and PIV at Re¼3.3  104. The actuators were activated continuously, in symmetric pulsed mode (top and bottom activated in unison), and in asymmetric pulsed mode (top and bottom activated in anti-phase). Here, the momentum coefficient was estimated to be Cμ ¼F/((½)ρU∞2d) ¼1.1% and 0.3% for steady and unsteady cases, respectively, where unsteady forcing was applied with 25% duty cycle. Large-amplitude vortex shedding was observed when excited near the natural shedding frequency (lock-on), but these authors concentrated on higher excitation

frequency with optimum f+ ¼ 1 (i.e. 5 times the natural vortex shedding frequency). At these frequencies, the vortex shedding was suppressed and the wake became much thinner and more streamlined. Reductions in turbulence levels in the wake of up to 80% were reported. Phase-locked PIV showed discrete shedding vortices from either side of the cylinder at the plasma forcing frequency (shed either symmetrically or asymmetrically depending on forcing mode), which propagated to and combined at the wake centreline. Both steady forcing and unsteady forcing eliminated the vortex street, reduced the wake turbulence and reduced the noise. Unsteady forcing produced additional noise at the forcing frequency but achieved flow control with only 25% of the power input. This study was extended to higher Reynolds number in Kozlov and Thomas [136], where co-flow, steady forcing was used at 7 901, only, at Re¼8.5  104. Using LDA and flow visualisation, they observed similar vortex shedding suppression, wake thinning and reductions in turbulence intensity to their low Reynolds number study. Both pulsed and continuous mode plasma

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Fig. 38. Near-wall velocity profile in a turbulent channel flow to show low-speed streaks (blue) and high-speed streaks (yellow-red). (a) Streamwise velocity contours near a wall controlled by a spanwise oscillatory force with T+ ¼100 and Lz+ ¼ 840 (Karniadakis and Choi [109]), (b) travelling-wave excitation of near-wall velocity with I¼ 1, λz+ ¼ 840 and T+ ¼50 and (c) no control (Du and Karniadakis [116]). (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)

Upper HV Electrodes Plasma Regions

HV RF Dielectric Barrier

U

Lower Ground Electrodes

Fig. 39. Spanwise oscillation plasma actuator. From Jukes et al. [14].

Fig. 40. Spanwise oscillation plasma actuators in quiescent air. From Jukes et al. [14].

Fig. 42. Cross section of the plasma actuators (a) and plasma timing (b) of spanwise travelling-wave excitation with λ+ ¼500 (100 mm) and T+ ¼ 82 (208 ms). From Choi et al. [123]. Fig. 41. Near-wall phase-averaged velocity vectors throughout the spanwise oscillation cycle. Solid line at the base of the plot indicates direction and timing of plasma force. From Choi et al. [123].

have been studied computationally by Kim and Wang [138] and Rizzetta and Visbal [139] using large-eddy simulation (LES) at Re¼3.0  104 and 104, respectively, who observed similar reductions

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in drag and suppression of the vortex shedding as those of the experimental works. Also motivated by landing gear noise reduction, Li et al. [140] used plasma actuators on a cylinder plus an oblique strut. Acoustic noise measurements showed that counter-flow oriented DBDs at 7901 reduced the noise level at high Reynolds number Re¼2.4  105, because flow separation was promoted, as observed

Fig. 43. Streamwise velocity distribution in a turbulent boundary layer with travelling-wave excitation (a) with λz+ ¼ 500 and T+ ¼ 82 and without control (b). From Whalley and Choi [124].

at Re¼1.8  104 by Sung et al. [132]. A similar study was conducted on a cylinder with a downstream torque link by Huang et al. [141] at Re¼2.1  105. Acoustic far-field sound measurements in an anechoic chamber and near-field PIV showed that steady co-flow and counter-flow plasma forcing at 7801 could either reduce or increase the noise and wake. These studies showed that upstream DBD forcing reduced the noise more effectively because this widened the wake and reduced the flow impingement on the downstream obstructions. There have been several studies of vortex shedding control of circular cylinders by plasma actuators by Jukes and Choi ([142– 145]). The first study at Re¼6.5  103 [142] used co-flow DBD plasma actuators at either 7701, 7891, 71001 or 71301 in pulsed mode with frequency 0.2≤f+≤4 and force coefficient Cμ ¼4.1%. Time-resolved PIV in the near wake revealed wake amplification and vortex shedding lock-on when forcing was applied at similar frequency to the natural shedding frequency (f+≈0.2), similar to that of Munska and McLaughlin [131]. Meanwhile for 0.8≤f+≤2 the vortex shedding was completely suppressed and small-scale vortices were shed from the upper and lower cylinder surface at the forcing frequency which mutually annihilated

Fig. 44. Distribution of Reynolds stresses with travelling-wave excitation. (a) Plasma forcing with λz+ ¼ 500 and T+ ¼ 82 in a turbulent boundary layer and (b) Lorentz forcing in a turbulent channel flow with I ¼1, λz+ ¼ 377 and T+ ¼40, From Whalley [9] for (a) and Huang et al. [125] for (b).

Fig. 45. Drag reduction (left) and lift fluctuation reduction (right) by plasma actuators. From Jukes and Choi [143].

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Fig. 46. Shedding modes with forcing frequency fp+ and force coefficient Cp. From Jukes and Choi [143].

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at the cylinder centreline, consistent with Thomas et al. [135]. However, for 2≤f+≤4 these vortices were very small and did not combine at the wake centreline. These merely travelled parallel to the freestream and vortex shedding still occurred, but slightly further downstream than without excitation. DBD plasma actuators at 7891 were most effective at this Reynolds number, although actuators at 71001 were also effective. However, actuators at 71301 had very little effect because they were too far downstream of separation, whilst upstream actuators at 7701 could promote only lock-on phenomenon. This study was extended to Re¼1.5  104 by Jukes and Choi [143], using time-resolved PIV with simultaneous direct measurements of the dynamic lift and drag forces. A wide range of forcing frequency and force coefficient were used (0.1≤f+≤2, 0.04%≤Cμ≤0.39%, note that Cμ was changed by varying the duty cycle from 1% to 40%), with actuators pulsed in unison at 7891. As shown in Fig. 45, large modifications to the lift and drag could be achieved depending on f+ and Cμ, where both increases and decreases of the aerodynamic forces could be achieved simply by changing the plasma actuation frequency and duration. A change in flow response was observed around f+ ¼0.6, where for f+ 40.6 the vortex shedding was suppressed (no-shedding regime), resulting in up to 32% reduction in drag and 72% reduction in lift fluctuations, whilst for f+ o0.6

Fig. 47. Time-averaged velocity (left), turbulence intensity (middle) and instantaneous vorticity filed (right) for flow around a circular cylinder at Re¼1.5  104 without control (top) and with co-flow unsteady plasma actuation at 7 891, f+ ¼ 1, Cμ≤0.17% (bottom). From Jukes and Choi [143].

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Fig. 48. (a) Vorticity contour and force vector (shown by arrow) and (b) flow visualisation around upper surface of the cylinder. From Jukes and Choi [145].

the vortex shedding locked on to a harmonic of the forcing frequency (lock-on regime), resulting in up to 8% drag increase and 87% increase in lift fluctuations. The shedding regimes together with sketches of the wake from PIV data are shown in Fig. 46. It was also found that pulsed forcing produced more effective flow control than continuous forcing. The no-shedding regime is characterised by small-scale vortices created in the separating shear layers at the plasma forcing frequency. Comparison between normal vortex shedding and with forcing at f+ ¼ 1 can be seen in Fig. 47, where the plasma forcing vastly reduces the wake turbulence and narrows the wake.

Symmetric pairs of vortices can be seen emanating from the upper and lower sides of the cylinder with control (lower right). Provided these are created at high enough frequency (so that they are small-scale), these inhibit the growth of shear-layer roll-up which prevents entrainment from the opposite shear layer so that the vortex shedding process cannot be maintained [143]. The mechanism of creating these vortices was studied further by Jukes and Choi [144,145] using PIV and dynamic force measurements at Re¼1.5  104. A single plasma actuator was placed on the upper surface of the cylinder, which was rotated to be placed at +601 to +1001 in 51 increments. The optimum actuator location was at +751 from the front stagnation point, which was just upstream of the phase-averaged separation point at 81.81. The vortex formation process is illustrated in Fig. 48. Frame (i) shows the flow just before the plasma was turned on, where shear-layer transition vortices can be seen just downstream of separation (x/d 40.4, y/d40.6). When the plasma was activated, momentum was added to the boundary layer so that flow separation was moved downstream (ii). This caused a break to occur between the reattached boundary layer and the shear layer that has already separated. The tail of the separated shear layer then rolled up and lifted away as it moved downstream ((ii): x/d ¼ 0.2, y/d ¼0.6 and (iii): x/d ¼0.6, y/d ¼0.7). Meanwhile, the re-energised boundary layer had sufficient momentum to remain attached to at least 1201 (iii). When the plasma was turned off (iv), the separation point retreated back upstream and the re-energised fluid rolled up into a vortex and detached from the upper surface (x/d ¼ 0.6, y/d ¼0.2– 0.3). This process was repeated each time the plasma was activated, regardless of timing relative to the natural vortex shedding frequency, so that a chain of small-scale vortices were shed from the cylinder at the forcing frequency. Jukes and Choi [145] were able to use a single pulse of DBD plasma near the separation point to make long lasting modifications to the vortex shedding cycle. Depending on when the vortex was released into the cylinder wake, the drag and lift fluctuation could be increased by 22% and 50% or reduced by 8% and 40%, respectively. This demonstrated flow control for a timescale of over 150 times the pulse duration with a power saving ratio, defined as the power saved to drag reduction to fluidic power introduced by plasma, of over 1000. More recent work has focussed on novel plasma actuators for vortex shedding control. Kozlov and Thomas [136] used plasma vortex generators over the forwards part of the cylinder (−901 to +901) with spanwise wavelength equal to half the cylinder diameter. Using LDA and flow visualisation at Re¼8.5  104, they could suppress the vortex shedding by introducing streamwise vorticity into the shear layers, which delayed flow separation due to cross-stream momentum transfer. However, there was a higher level of wake turbulence than with co-flow plasma actuators at 7901 due to enhanced three-dimensionality in the wake. Yamada et al. [146] use symmetric plasma vortex generators across the separation region (80–1441) with wavelength equal to the cylinder diameter at Re¼103. These plasma vortex generators were used in unsteady mode at f+ ¼0.22, 1.0 and 2.0 with 30% duty cycle. Flow visualisation clearly showed streamwise vortices in the wake at x/ d¼ 1. X-wire wake measurements showed that vortex shedding could be suppressed in a similar manner to co-flow plasma forcing, but again there was enhanced three-dimensionality in the wake. No attempt was made to optimise the spanwise electrode spacing in either study. Three-dimensional forcing has also been attempted by Gregory et al. [147], who used sinusoidal induced velocity pattern to alternate wall-normal and wall-tangential blowing at Re¼6.5  103. Bhattacharya and Gregory [148] used a square wave induced velocity pattern. Their goal was to break up the spanwise coherency in the vortex street, thereby promoting phase cancellation

1

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Fig. 49. Time-averaged streamlines of the flow around a circular cylinder: (a) without control and (b) DBD plasma control with fe/fo ¼ 1, Cμ ¼ 0.052 (duty cycle λP ¼ 80%). From Feng et al. [150].

between spanwise portions of the wake structure. In this study, co-flow steady forcing was applied at +801 but with intermittent pattern along the span (wavelength of 4d). X-wire wake measurements showed that this configuration caused greater drag reduction than continuous spanwise forcing, although the exact mechanisms are still to be identified. Sliding discharges have also been used around circular cylinders. These actuators consisted of a single asymmetric plasma actuator with a third electrode on the exposed surface to which a DC voltage was applied. This effectively stretched the plasma to allow creation of a large-volume discharge. Sosa et al. [149] directly measured a reduction in the drag coefficient on a circular cylinder by up to 25% at up to Re¼1.2  104 using a sliding discharge from 7901 to the rear stagnation point. This was a similar drag reduction by standard actuators at 7 901, but the sliding discharge required lower power so that the flow control efficiency was increased. Li et al. [140] also showed slight improvement of sliding discharges over conventional DBDs for landing gear noise control at Re¼2.4  105 using acoustic and PIV measurements. These actuators were used only in steady mode. Feng et al. [150] carried out a vortex shedding control of a circular cylinder by actuating a plasma synthetic jet at the natural shedding frequency from the rear stagnation point at Re¼5  103. Time-resolved PIV indicated that the plasma actuator introduced the high-momentum fluid into the recirculation region. This reduced the velocity defect and changed the flow topology in the near wake region. An additional recirculation region was introduced just upstream of the original recirculation region by a periodic plasma forcing, which can be clearly seen in Fig. 49. Such control pattern is similar with the previous finding by using the piston-actuated synthetic jet [151–153].

Novel applications of DBD plasma actuators for aerodynamic control have also been reviewed in this article, including pitch and roll control, plasma jet vectoring, circulation control and plasma flap. There, we showed DBD plasma actuators that can replace movable control surfaces of aircraft. For example, some DBD plasma actuators can act like slats and flaps when they are placed at the leading or trailing edge of an aerofoil. Plasma actuators can also be combined with a Gurney flap to further enhance the lift on aerofoils as and when required. It was also shown that DBD plasma actuators can replace ailerons for roll control. It now looks possible to design a new generation of UAVs without movable control surface based on DBD plasma actuators. We have also discussed on the starting vortex that leads to a formation of plasma wall jet. This is an important subject not only for a better understanding of the flow induced by DBD plasma actuators, but also as a database that can be used to calibrate the numerical models for plasma flow control. Design of DBD plasma actuators to obtain turbulent skin-friction reduction was also shown and the modifications to near-wall turbulence structures were summarised. Finally, vortex shedding control techniques for drag and lift suppression were surveyed, demonstrating a true potential of plasma actuators for active flow control.

5. Conclusions

References

Recent developments in DBD plasma actuator design have been reviewed, including plasma synthetic jet actuators, plasma spark jet actuators, three-dimensional plasma actuators and plasma vortex generators. As compared to conventional plasma actuators which can induce only two-dimensional wall jets close to the wall, these actuators are so designed that three-dimensional flows can be induced either parallel or normal to wall. For example, plasma vortex generators can generate streamwise vortices when the actuators are placed in the freestream at a yaw angle, while plasma synthetic jet actuators can issue wall normal jets with or without cross-stream. It should be emphasised here that the material and construction of these novel plasma actuators are identical to those of conventional DBD plasma actuators. Only the design and positioning of actuators are different, which enable them to induce three-dimensional flows. As such, the mixing capabilities of these plasma actuators are naturally greater, enhancing the flow control performance in flow separation, fluid mixing and combustion.

Acknowledgements This work was supported in part by the National Natural Science Foundation of China (Grant no. 11202015). TNJ and RDW acknowledge the support in part by the Nottingham Advance Research Fellowship and the EPSRC PhD Plus Fellowship, respectively, to carry out this work.

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