Content uploaded by Carlos Xisto
Author content
All content in this area was uploaded by Carlos Xisto on Jul 29, 2014
Content may be subject to copyright.
PECyT - Plasma Enhanced Cycloidal Thruster
Carlos M. Xisto,∗Jos´e C. P´ascoa,†M. Abdollahzadeh‡and Jakson A. Leger§
Universidade da Beira Interior, Covilh˜a, 6201-001, Portugal
Pierangelo Masarati¶and Louis Gagnon k
Politecnico di Milano, 20156 Milano, Italy
Meinhard Schwaiger∗∗and David Wills ††
IAT21-Innovative Aeronautics Technologies GmbH, A-4050 Traun, Austria
In the following paper we introduce PECyT system for enhancing the aerodynamic
efficiency of cycloidal rotors. For that purpose the incorporation of Dielectric Barrier
Discharge plasma actuators for active flow control on a pitching airfoil, under deep-stall
conditions, will be assessed using a numerical tool. Two different arrangements of DBD
actuators will be analysed, namely single- and multi-DBDs configurations. For the single-
DBD plasma actuator the effect of different modes of actuation on the lift coefficient will
also be studied. We will show that the multi-DBD actuator, in a steady-actuation mode,
could delay stall and allows for a faster reattachment of the flow. However during the
downstroke phase of the pitching cycle the unsteady operation of a single-DBD gives us
the best results in terms of lift coefficient.
I. Introduction
Acyclorotor, also known as a cyclocopter or a cyclogiro, is a rotating wing machine where the axis of
rotation is parallel to the blade span, see Fig. 1-a). It consists of a set of blades that are connected to
a mechanical system with the ability to define a periodic pitching schedule for the wings as they preform
one revolution around the horizontal axis. This allows us to define a periodic variation of the angle of
attack (AoA) on each blade and to control the direction and magnitude of the resultant thrust vector almost
instantly in 360 radial direction, thus allowing for a substantial increase in flight control.
Cyclorotors were studied in the 1920’s by Kirsten with the cooperation of Boeing1as a mean of propulsion
for aerial and marine crafts. In the mid 30’s they were deeply study, under contracts for the National
Advisory Committee for Aeronautics, by Strandgren2who developed an analytical model able to predict the
amount of lift and propulsion that could be generated by a cyclorotor, and by Wheatley3,4who conducted
experimental activities and also theoretical studies on the efficiency of cyclorotors in hovering and forward
flight conditions. Wheatley tested a four-blade rotor with a span and diameter of 8 feet (2.4384m) and chord
length equal to 0.312 feet (0.0950976m) in a 20-foot (6.09600m) wind tunnel, showing that the tested rotor is
able to take-off vertically and hover in air, fly horizontally and auto-rotate in the case of main engine failure.
Wheatley concluded that cyclorotors could have a practical application into aeronautical propulsion, but
that the probable achievable performance is very poor when compared with the available screw propellers of
that period. However, only one rotor geometry was analysed and several aspects related to the flow around
pitching airfoils were not fully comprehended during that time.
∗Research Fellow, Dep. de Eng. Electromecˆanica, Rua Mqs D’ ´
Avila e Bolama, xisto@ubi.pt, AIAA Member
†Professor Auxiliar, Dep. de Eng. Electromecˆanica, Rua Mqs D’ ´
Avila e Bolama, pascoa@ubi.pt
‡Ph.D. Student, Dep. de Eng. Electromecˆanica, Rua Mqs D’ ´
Avila e Bolama, mm.abdollahzadeh@yahoo.com
§Ph.D. Student, Dep. de Eng. Electromecˆanica, Rua Mqs D’ ´
Avila e Bolama, leger mec@hotmail.com
¶Associate Professor, Dip. di Scienze e Tecnologie Aerospaziali, via La Masa 34, pierangelo.masarati@polimi, AIAA Member
kResearch Associate, Dip. di Scienze e Tecnologie Aerospaziali, via La Masa 34, louis.gagnon@polimi.it
∗∗General Manager, IAT21, Langholzstraße 16, meinhard.schwaiger@iat21.at
††Business Director, IAT21, Langholzstraße 16, davidwills@fastmail.fm
1 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference
July 28-30, 2014, Cleveland, OH
AIAA 2014-3854
Copyright © 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
Propulsion and Energy Forum
(a) (b)
Figure 1. a) 3D representation of a cyclorotor with six NACA0012 blades and a maximum pitch angle of 40
degree. b) IAT21 D-DALUS concept uses four contra-rotating cycloidal rotors for propulsion.
Such system was never accepted in the aeronautical industry due to the lack of technology and due to
the fact that old days analytical tools were unable to analyse the complex flow that was associated with
the interaction between the pitching rotating blades. New advancements in technology are now opening
the possibility of introducing the cyclorotor concept in the aeronautical market as a propulsion system for
airships5,6,7and into disruptive aircrafts, see Fig. 1-b).
During the last decade several research teams are working on the applicability of cyclorotors in MAV-
scale (Micro Air Vehicles) concepts. Iosilevskii and Levy8performed experimental and CFD studies on
the efficiency of a MAV-scale four-blade cyclorotor with span and chord length equal to 0.110 and 0.022 m,
respectively. They found out that in terms of performance the cyclorotor is comparable with an heavy-loaded
helicopter rotor. Sirohi et al.9analyzed the hoover performance of a cyclorotor with a diameter and span
of 6 inches (0.1524 m). In their research they have also included an analytical model able to predict rotor
performance, where vertical axis wind turbine theory was adapted in the context of aeronautical propulsion.
A comparison between the tested rotor and a conventional rotor, of the same diameter, showed that the
cyclorotor was able to achieve higher power loading. Hwang et al.10,11 developed a four-rotor cyclocopter
with elliptic blades that produces a spanwise uniform distribution of the induced velocity, that could open
the possibly of exploiting aeroelastic tailoring in such a way that the aerodynamic loads themselves induce
twist changes, which will improve the distribution of induced velocity for different operating conditions. A
throughout study on the understanding of cyclorotors was taken by Benedict et al.12,13,14,15 They have
study12 the influence of the number of blades on the performance of a 0.152 m in diameter/span cyclorotor
composed by NACA0010 airfoils with a uniform chord of 0.0254 m. Their main conclusion was that, for
small scale cyclorotors with a number of blades up to five and a maximum pitching amplitude of 40 deg,
the power loading increases with the increasing number of blades. Different rotor parameters, like airfoil
section, blade flexibility, blade camber, rotor radius, blade span, rotor aspect-ratio, rotor solidity, blade
planform and blade kinematics were also analysed by Benedict et al.13,15 They found out that for low-
Reynolds numbers the airfoil section does not plays a significant role in the aerodynamic performance of a
cyclorotor, since the inverted NACA0010 airfoil produces similar values of efficiency when compared with
the baseline NACA0010, cambered blades resulted in lower efficiency when compared with symmetrical ones.
They also concluded that the virtual camber effect and the chord-to-radius ratio play a significant role in the
aerodynamic efficiency. Regarding the pitching schedule, they observed that using an asymmetric pitching
could be beneficial if one selects an higher pitching amplitude on the top blade.
The EU-FP7 funded CROP (Cycloidal Rotor Optimized for Propulsion) project strives to design a
cycloidal propulsion system that is suitable for manned and unmanned aerial vehicles. In order to asses
the applicability of cyclorotors in large scale, several studies have been taken using pure analytical16,17
and CFD18,19 approximations. It was concluded that CROP could introduce into the aeronautical market
several advantages in comparison with traditional VTOL, fixed wing air vehicles and hybrid aircrafts. It uses
2 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
common surfaces to achieve lift and thrust along the full range of flight speeds. The problem of retreating
blade stall of helicopters is not so severe in cycloidal rotors, meaning that they can achieve higher subsonic
speeds. The use of a wing rotating around the axial axis creates lift, and thrust, when the blades move
backward in relation to the vehicle direction of flight. This makes possible to use the intermittent, but very
high, lift value generated by the unsteady pitching of the blades. Further, each blade of the cycloidal rotor
operates at similar conditions (angle of attack, velocity, Reynolds number) so, in principle, the blades are
easier to optimize in terms of aerodynamic performance.20
The unsteady flow mechanism that is generated by the pitching movement of the blades also plays a
significant role in the aerodynamic efficiency of the cyclorotor, since it can delay blade stall, thus increasing
the amount of lift that can be produced by each blade. This phenomenon is known as dynamic stall and
its predominant feature is the creation and convection of a vortex disturbance on the suction surface of the
airfoil that will resist to the adverse pressure gradients that are responsible for flow separation. This effect
as been extensively studied using experimental techniques21,22 and CFD tools.23,24,25,26
Since the main aerodynamic component of a cycloidal rotor comprises a pitching airfoil, any flow control
techniques that could increase the aerodynamic efficiency dealing with stalled flow over the airfoil could be
beneficial. Several active and passive flow control mechanisms had been reported in literature for controlling
the flow around pitching airfoils, among them Dielectric Barrier Discharge (DBD) plasma actuators are
considered a promising technique.27,28,29,30 DBD plasma actuators31,32,33 are novel means of controlling
flow which imparts momentum to the fluid by creating an ionic-wind beside the control surface. These
actuators have several advantages in comparison to other flow control systems, which includes being fully
electronic, having a fast response time, less complexity and easy integration into the system. In the following
paper we intend to study the idea of possible inclusion of DBD actuators for active flow control in cycloidal
rotors, where here after we call it PECyT (Plasma Enhanced Cycloidal Thruster). We will show that the
combined effect of the leading edge vortex (LEV) with the PECyt system could delay separation at high
angles of attack. Using numerical models for gas dynamic34,35 with a suitable multiphysic CFD models,36,37
numerical analysis of such complex flows was performed. For that purpose we will consider an individual
airfoil performing a sinusoidal pitching movement (the most simple aerodynamic equivalent system). The
results will provide information for design an optimization of DBD plasma actuators that could control the
flow for cycloidal rotors.
II. Numerical Model
To model the effect of plasma actuation, we will consider a simple phenomenological model, which will
consider the effect of the actuator as a body force. Thus, the Unsteady Reynolds Averaged Navier-Stokes
(URANS) equations under the influence of a body force are given by:
∂ρ
∂t +∇ · (ρU) =0
∂ρU
∂t +∇ · (ρUU) = − ∇p+∇ · τ+F
(1)
Where Frepresents the external eletrodynamic body force generated by the plasma actuator. In Eq. (1)U
is the velocity field vector, pis the static pressure, ρrepresents the density of air and τis the viscous stress
tensor. The body force is here added under the assumption that the plasma formation and flow response
are decoupled from each other. Such approximation can be applied due to the big disparities between the
characteristic velocities associated to each physical process.32
The external body force in the momentum equation is given by:
F=ρcE,(2)
where ρcis the net charge density and Eis the electric field vector. The Maxwell equations for a quasi-steady
plasma, with the assumption of negligible time variation of the magnetic field, are the following:
∇ · (ε0εrE) = ρc,∇ × E= 0.(3)
Where εrand ε0are the relative permittivity of the medium and permittivity of free space, respectively.
The electric field can be derived from the gradient of a scalar potential,
E=−∇Φ,(4)
3 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
and
∇ · (εr∇Φ) = −ρc
ε0
.(5)
Since the gas particles are weekly ionized, we can decouple the potential in two parts,38
Φ = ϕ+φ, (6)
where φis the potential due to external electric field and ¯ϕis related to the potential due to net charge
density in the plasma. For the potential due to the external field,
∇ · (εr∇φ) = 0.(7)
The charge density could be related to the electrical potential with the Debye length,
ρc
ε0≈ − 1
λ2
d
ϕ, (8)
where λdis de Debye length. An equation for the charge density can now be written by using Eq. 9, 7 and
6,
∇ · (εr∇ρc) = ρc
λ2
d
.(9)
The value for the Debye length is estimated with the relation proposed by,39
λd= 0.20.15 ×10−6Vpp −7.42 ×10−4(10)
where Vpp is the applied peak-to-peak voltage.
The above described model assumed a Gaussian distribution of charge density in accordance to experi-
mental observations,
ρc,w (t) = ρmax
cG(x)f(t) (11)
G(x) = exp "−(x−µ)2
2σ2#(12)
where µis a location parameter that indicates the position of the function maximum, and σrepresents the
rate of decay of charge density. We note that Eq. 11 as a significant effect on the accuracy of the numerical
simulations. In Suzen et al.38 a value for ρmax
cwas selected accordingly with experimental data. In the
present work we apply the modification of Abdollahzadeh et al.33 that has scaled ρmax
cand the electric field
with the following functions:
ρmax
c= 2fCeq.
Vpa −Vbd
fcorr λd
, fcorr =1
2√2πσ
lp"erf 1
2
µ√2
σ!+ erf 1
2
√2 (lp−µ)
σ!#; (13)
E=E0E∗, E0=(Vpa −Vbd)
lp
.(14)
In Eq. 13 fis the AC voltage frequency, Vpa =Vpp/2 is the applied voltage peak amplitude, Ceq is the
equivalent capacitance of the DBD actuator, Vbd represents the critical breakdown voltage where ionization
occurs and lpis an estimation of the length of the plasma region. These scaling functions were proposed by
Abdollahzadeh et al.33 in order to reduce the dependence on empirical fitting parameters and to improve
the applicability of the model into different practical applications.
We note that Eqs.(7) and (9) do not have any temporal derivative term, only Eq. 11 is time dependent.
We can then solve them in terms on normalized variables,
ρ∗
c=ρc
ρmax
c
, φ∗=φ
φmaxf(t)E∗=∇φ∗=lp∂φ∗
∂x i+∂φ∗
∂y j,(15)
where f(t) is the shape of the applied voltage. We can return to the dimensional variables by multiplying
the electric field and charge density distributions by the corresponding normalization factor.
4 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
III. Results and Discussion
In the following section we apply the numerical model of section II into the analysis of DBD plasma
actuation in the flow of a pitching airfoil. We start by validating the numerical model in the computation
of a purely gas dynamic flow (no-actuation) under deep-stall conditions. Afterwards in section Bthe effect
of plasma actuation will be assessed in terms of number of DBD actuators and operation mode.
A. Validation
We start by validating the gas dynamic model without the inclusion of plasma actuation. For that purpose
we selected the pitching NACA0012 airfoil experimental test case of Lee and Gerontakos21 in deep-stall
conditions. This experiment was preformed in a suction type low-speed wind tunnel with 0.9 m×1.2 m×2.7 m
operating at a turbulent intensity of I= 0.08% at U∞= 35 m/s. The aluminium NACA0012 airfoil had
a chord length, c, equal to 0.150 m and a span length, s, equal to 0.375 m, the pitching axis was located
at a quarter of the chord length (0.25c). In order to minimize the three-dimensional flow effects at the
measurement location (airfoil midspan), two circular end plates with 0.3 m of diameter were fitted into the
experiment, the gap between the endplates and the airfoil tip was less than 1 mm. By using a 5 µm hot-wire
probe the two-dimensional uniformity of the flow was measured, and was found that the 2D non-uniformity
was around ±4% of the free-stream value, the wind tunnel maximum blockage was around 5%. The pitching
movement was controlled by a four-bar-linkage and flywheel mechanism that is capable to oscillate the airfoil
at several amplitudes and frequencies, the instantaneous pitching amplitude,
α(t) = αm+ ∆αsin (ωt) (16)
is a function of time defined by a mean pitching angle, αm, an oscillating amplitude, ∆α, and a circular
frequency, ω= 2πf. Here the deep-stall condition was used for validating the gas dynamic numerical model:
αm= 10 deg; ∆α= 15 deg; and ω= 18.67 rad/s for a reduced frequency of k=ωc/(2U∞) = 0.1. The mean
free-stream velocity was U∞= 14 m/s corresponding to a chord Reynolds number, Re = U∞c/ν = 1.35×105.
Figure 2shows an overall and partial views of the numerical domain. It has a circular shape with R= 50c
and it is composed by an hybrid mesh with a O-type structured mesh in the boundary layer region and an
unstructured mesh in the remanding domain. The structure mesh was used in order to achieve a desirable
value for the y+, in our case less than one. Three different regions were created: an outer-region, a refined
region for the wake and a blade region. The blade region is a circular shaped domain with center located
at a quarter of the chord length, prescribing the movement that is given by Eq. 16. This region exchanges
information with the outer and wake regions through a numerical interface. Regarding boundary conditions,
in the outer boundary we have specified an inlet-outlet BC where we impose U∞= 14 m/s in the xdirection,
a turbulent intensity of I= 0.08% and a turbulent length scale of 0.02c.
For solving the incompressible Unsteady Reynolds Averaged Navier-Stokes equations a pressure-based
coupled solver was used, where a Courant number of 100 was selected within each time step. The desirable
time and space accuracy were achieved by using the second-order implicit and second-order linear-upwind
interpolation schemes, respectively. For turbulence modelling, since the experimental data21 was showing
clear signs of transition the k-ω-SST transition model40,41,42 was selected. It was shown by Wang et al.25
that URANS with k-ω-SST transition is capable to capture major flow features of dynamic stall, namely the
formation and development of the LEV and the hysteresis loop, the dynamic loading acting on the airfoil
and also the unsteady behaviour of the lift coefficient in post-stall conditions.
A grid sensitivity study was taken in order to achieve a final solution that is independent from the
mesh dimension. Four different grids were build with increasing size: Grid-0 with ∼89,000 cells; Grid-1
with ∼180,000 cells; Grid-2 with ∼582,000 cells; and Grid-3 with ∼873,000 cells. In Fig. 3-a) we show the
experimental results for the lift coefficient (CL) and the computed values obtained with the four different
grids as a function of the pitching angle. For all the cases we have computed 8 complete cycles (only the
curves of the 8th cycle are displayed), although a time-converged solution was found after 4 cycles. The
time-step that was used in the grid independence study was dt2 = Tc/1000, where Tc= 2π/ω is the cycle
period. We can see that in the upstroke phase the curves of CLalmost overlap with each other in the entire
half-cycle. The biggest discrepancy for the upstroke phase is located in the peak region, see Fig. 3-b).
Here the results obtained with Grid-2 and Grid-3 show a good agreement, while the results of Grid-0 and
Grid-1 are slightly different. In the downstroke phase the flow is significantly unsteady, thus we have a more
pronounced discrepancy between the curves. Nevertheless the results obtained with Grid-2 and Grid-3 are
5 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
= 50
Wake Region: 300,000 cells
Blade Region: 200,000 cells
Interface
Grid-2: 582,381 cells
∞
y
x
Figure 2. Numerical grid (grid-2) used for the computation of Lee and Gerontakos21 pitching airfoil in deep-
stall conditions.
very similar in terms of average values. Based on these observations we assume that the results obtained
with Grid-2 are acceptable and we will use this mesh for the remaining part of this paper. For all the four
grids a time-step independence study was also preformed. In Fig. 3-c) we show the results obtained for
CLin grid-1 for three consecutively refined time-steps: dt1 = Tc/500; dt2 = Tc/1000; dt3 = Tc/2000. We
can observed that there isn’t a significant difference in the plots dt2 and dt3, thus, and following the same
conclusions of the grid independence study, we will use the dt2 time-step for the remaining part of this paper.
Let’s now compare the experimental data with the numerical results obtained with grid-2 and dt2. Figure
3-a) shows that the numerical model is able to predict with some accuracy the time-dependent values of CL
even in the downstroke phase of the cycle, when the flow is completely separated. The biggest discrepancy is
found in the peak region, where the numerical model is predicting the stall onset to occur sooner. This effect
should be related to a non-well resolved LEV in terms of size and intensity. The formation and convection of
this vortex structure induces pressure fluctuations that keep the flow attached at high angles of attack, when
the LEV leaves the trailing edge and goes into the wake the flow completely separates and there is an abrupt
decrease in Lift and an increase in Drag coefficients. Figure 4-a) shows the instantaneous non-dimensional
vorticity (ω∗
z=ωzc/U∞) contour plots for the upstroke and downstroke phase of the cycle. We can see that
the current model is able to capture the formation and convection of the LEV. When the LEV leaves the
top surface of the blade the flow becomes completely separated and it will remain in stall conditions for
almost the entire downstroke phase until reatchament is achieve at 5.4 deg. Although over-predicting Drag
in several pitching positions, the computed result of drag coefficient CDwith the transition-SST model is
showing in general a good agrement with experiments, see Fig. 3-d).
B. DBD plasma actuators for active flow control
Dielectric Barrier Discharge plasma actuator is a very promising technology for active flow control at reduced
cost and weight. Plasma-based devices exploit the momentum coupling between the surrounding gas and
plasma to manipulate the flow. When compared with other flow control techniques, these devices require low
power consumption and do not require any moving mechanical parts. They are suitable for incorporation
in thrust vectoring propulsion systems43 since they have a very fast frequency response, that allows a real-
time control of the fluid flow. Figure 5-b) shows a single-DBD, it consists in two electrodes separated by
a dielectric material usually glass, Kapton or Teflon. One electrode is exposed to the air and the other
is fully covered by a dielectric barrier. The exposed electrode is loaded by an high AC voltage signal, of
sufficient amplitude (5-40kV) and frequency (1-20 kHz), whereas the covered electrode is grounded. The
intense electric field partially ionizes the surrounding gas producing non-thermal plasma on the dielectric
surface. The ionized air (plasma) in the presence of the electric field produces an attraction/repulsion force
on the charged particles. The ionized particles are then accelerated and transmit their momentum, through
collisions, to the neutral air particles in the plasma region over the covered electrode.44 The result is the
6 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
-1
-0.5
0
0.5
1
1.5
2
2.5
3
-5 0 5 10 15 20 25
CL
α (º)
Experimental
Grid-0
Grid-1
Grid-2
Grid-3
(a)
1.7
1.8
1.9
2
2.1
2.2
2.3
2.4
2.5
20 21 22 23 24 25
CL
α (º)
Experimental
Grid-0
Grid-1
Grid-2
Grid-3
(b)
-1
-0.5
0
0.5
1
1.5
2
2.5
3
-5 0 5 10 15 20 25
CL
α (º)
Experimental
Grid-1-dt1
Grid-1-dt2
Grid-1-dt3
(c)
0
0.2
0.4
0.6
0.8
1
1.2
-5 0 5 10 15 20 25
CD
α (º)
Experimental
SST-Transition
(d)
Figure 3. a) Lift coefficient as a function of the pitching angle computed on several grids. b) Detail of CL
in the upstroke phase (20-25 degs). c) Time-step independence study computed on grid-1. d) Comparison
between experimental data and computed values (grid-2) for the drag coefficient.
creation of an induced velocity (ionic wind) in the proximity of the dielectric surface.
At critical values of pitching amplitude, the cyclorotor blades could experience flow separation. In stall,
lift coefficient rapidly decreases and drag coefficient increases. The main goal of PECyT is to delay stall
at high pitch amplitudes thus increasing the lift and the efficiency of the rotor. The PECyT-DBD induced
flow can play a significant role in delaying boundary layer separation. However, for an efficient flow control,
the location of the actuator needs to be optimized based on the chord position where separation occurs.
Such position will vary with several flow parameters, such as the pitch angle and rotation speed. For an
efficient control of the aircraft the pitch angle and rotation speed of the rotor will be continuously modified,
so the rotor blade could experience stall at different chord locations. Because of this, the usage of a single-
DBD might not be enough for providing an efficient mean of flow control. In Fig. 5-a) we show a possible
configuration of a rotor blade with a multi-DBD plasma actuator (Fig. 5-c). Such system could be adjusted
using a suitable control-loop to activate a single or a combination of DBD actuators, thus controlling the
boundary in the most optimum manner.
7 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
1.2 deg
12.3 deg
22.1 deg
23.4 deg
24.8 deg
25 deg
/∞
10 deg
5.4 deg
18.8 deg
20.6 deg
24.8 deg
23.4 deg
22.1 deg
20.6 deg
18.8 deg
1.2 deg
LEV formation
LEV
Separated flow
LEV
LEV
Reattachment
(a)
1.2 deg
12.3 deg
22.1 deg
23.4 deg
24.8 deg
25 deg
/∞
10 deg
5.4 deg
18.8 deg
20.6 deg
24.8 deg
23.4 deg
22.1 deg
20.6 deg
18.8 deg
1.2 deg
(b)
Figure 4. Instantaneous vorticity contour plots for a pitching cycle obtained with: a) the baseline case (without
plasma actuation); b) unsteady actuation St= 0.5with leading edge actuation.
1. Steady actuation
Firstly we will considerer the application of DBD actuators under a steady-actuation condition. In this case,
since the AC frequency is much more higher than fluid response frequency, the flow will sense a constant
body force. Two different arrangements of DBDs were selected for the steady-actuation analysis. In the first
configuration only one single-DBD located at the leading edge (x/c = 0) was activated. For the second case
a multi-DBD configuration was selected, which is composed by three pairs of electrodes that are placed 2
mm apart from each other in the suction surface of the airfoil. The idea behind the usage of multi-DBD
configurations is that one could achieve a multiplication factor that will increase the maximum induced
velocity per actuator and also cover a larger surface of the airfoil by actuator zone. Such effect is only
possible in steady-actuation conditions where a quasi-steady ionic wind is achievable, see Figure 6-c) and d).
Each one of the DBDs consisted in two electrodes, one being exposed to the air and the other being
covered with a dielectric material. The exposed electrode is 3 mm wide and the covered electrode as a
length equal to 4 mm. For the dielectric material we have assumed a thickness equal to td= 0.1×10−3m,
which is a common thickness for two layers of Kapton film,30 for a more detail picture of the dimensions
see Fig. 6-a). The signal supplied to the plasma actuator was a 7kHz sinusoidal wave with Vpp = 8 kV. For
the steady-actuation case this voltage was provided in a continuously way to the electrodes, allowing the
creation of a quasi-steady induced flow.
Figure 6-a) shows a detail of the grid near the electrodes. The mesh is very similar to grid-2 but it
further refined near the actuation zones that are given by the intersection point between the exposed and
cover electrode. This was done because the computational cells need to be small enough to properly compute
8 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
Multi-DBD Dielectric
Plasma
Exposed Electrode
AC voltage
Covered Electrode
(a) (b)
(c)
Figure 5. a)Possible arrangement of a multi-DBD plasma actuator on a NACA airfoil. Graphical representation
of a single- (b) and multi- (c) Dielectric Barrier Discharge plasma actuator.
the thickness of the plasma region (body force region) that, depending on the operating conditions, could
be some tens of λd.
In Fig. 6-b) we show the specified boundary conditions for the DBD model and the computed distribution
of the non-dimensional electric potential field. In the exposed electrode we specify the maximum value of
φ∗= 1 and in the covered electrode we impose a zero value for the electric potential. For the dielectric
surface we assumed a coupled boundary condition, by forcing the gradient of the electric potential to be
equal at the sides of the surface of the dielectric layer. In the remanning walls we assume that they are
perfect insulators of electricity. Regarding the boundary conditions for the charge density we specify a profile
for the discharge surface with Eq. 11 and a zero gradient boundary condition for the remanning boundaries.
In Fig. 7we show the CLresults with and without actuation. Figure 7-a) shows that for almost the
entire upstroke phase of the pitching cycle the inclusion of DBD actuators in steady actuation mode, with a
single- or multi-DBD arrangement, do not influence the lift coefficient profile as the pitching angle increases.
This was also observed experimentally by different authors.28,30 The effect of the plasma actuation is more
strong in the stall region where we can observe that the flow separation is slightly delayed. We note that
this effect is more pronounced with the multi-DBD configuration, see Fig. 7-b). However the maximum
lift is also slightly reduced. From Fig. 7we can observe that the intersection point between the upstroke
and downstroke points is occurring sooner for the actuation case, again this effect is more pronounced in
multi-DBD configuration, since more momentum is being applied to the flow. For the downstroke phase
it is visible that higher steady actuation is obtained with the multi-DBD, thus promoting separation and
reducing the generated lift force.
2. Unsteady actuation
The second case that we have analysed is related to the unsteady operation mode of the DBD. In this case the
AC signal that is supplied to the electrodes was cycled off and on with an unsteady period. The time that the
AC voltage is activated within that period is called duty-cycle and is normally given in terms of percentage
of the unsteady period, in our case the duty-cycle percentage was set at 10%. This unsteady actuation
promote the creation of non-stationary disturbances that, if generated near the separation point, will lead to
vortical structures. These vortices will increase the momentum on the suction surface of the blade, allowing
the flow to withstand the adverse pressure gradients, which are responsible for flow separation. We still need
to select the optimal forcing frequency that should be selected in a way that the DBD could generate and
convect 2-3 vortical structures over the separated zone of the airfoil.28 This could be attained if the Strouhal
number,
St=fc
U∞
,(17)
is equal to one. This empirical relation is appropriate for static airfoils, however for pitching airfoils it was
found that under certain conditions a lower forcing frequency (lower St) could be more beneficial in terms of
9 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
4 mm
3 mm
2 mm
NACA0012
y
x
= 0.1 mm
Covered Electrode
Exposed Electrode
(a)
Covered Electrode
Exposed Electrode
(V)
∗= 1
∗
= 0
∗= 0
∗
∗
= 0
∗=,
(b)
(m/s)
DBD
Single-Actuation
(c)
(m/s)
DBD-1
DBD-2 DBD-3
Multi-Actuation
(d)
Figure 6. a) Detail of the grid used for the computation of flow with DBD plasma actuation. b) Boundary
conditions imposed for the DBD numerical model and the computed distribution of the non-dimensional
electric potential field. c) Quasi-steady ionic wind obtained with a single DBD. b) Induced flow computed
with the multi-DBD arrangement.
flow control. It was shown by Post and Corke28 that a St= 0.25 was responsible for an increase of the CL
in almost the entire pitching cycle. The work performed by Mitsuo et al.30 have shown that the best results
in terms of efficiency were obtained for a St= 0.5. For the pitching airfoil setup of Lee and Gerontakos21
that we are using here (c= 0.150 m and U∞= 14 m/s) the optimal forcing frequency for a St= 1 is equal
to f=U∞/c = 93.3 Hz. In this paper we compute the unsteady actuation for St= 0.5, meaning that the
applied forcing frequency is equal to f= 46.67 Hz.
Figure 7-a) shows the lift coefficient as a function of the pitching angle, where the values for steady and
unsteady actuation could be compared with the case with no-actuation. We can see that the CLremains
unchanged for the unsteady actuation in almost the entire half-pitching cycle. Here the flow is completely
attached to the suction surface of the airfoil and the introduction of small disturbances will not affect the
generated force. In Fig. 7-b) we can see that the stall angle is similar to the one where there isn’t any
actuation but the maximum lift is slightly reduced. This could be related to a small perturbation of the
LEV that at this point could not be beneficial. Such perturbation is visible in Fig. 4-b) for the upstroke
angle of 23.4 deg (near stall). The biggest visible effect is perhaps in the downstroke cycle, when the flow
is completely separated. We can see in Fig. 7-a) that the unsteady actuation can improve the aerodynamic
efficiency of the airfoil for certain pitching angles (20-15 deg). For angles within this range, namelly 20.6
deg and 18.8 deg a visible reduction of the separtaion bubble is observed in Fig. 4-b). The intersection point
between the half-cycles is similar to that obtained with no-actuation.
10 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
-1
-0.5
0
0.5
1
1.5
2
2.5
3
-5 0 5 10 15 20 25
CL
α (º)
No Actuation
Steady SDBD
Steady Multi-DBD
Unsteady SDBD
(a)
1.7
1.8
1.9
2
2.1
2.2
2.3
2.4
20 21 22 23 24 25
CL
(º)
No Actuation
Steady SDBD
Steady Multi-DBD
Unsteady SDBD
(b)
-0.1
-0.05
0
0.05
0.1
0.15
0.5 1 1.5 2 2.5
CL
(º)
No Actuation
Steady SDBD
Steady Multi-DBD
Unsteady SDBD
(c)
Figure 7. a) Lift coefficient as a function of the pitching angle with and without plasma actuation. b) Detail
of CLin the stall region. c) Detail of the intersection between the upstroke and downstroke CLlines.
IV. Conclusion
In the present paper we have analysed with numerical tools the influence of plasma actuation in pitching
airfoils under deep-stall conditions. Two modes of actuation were analyzed, namely steady actuation and
unsteady actuation. For the steady actuation case different arrangements of DBDs were studied, namely a
single-DBD, positioned at the leading edge, and a multi-DBD configuration.
We have demonstrated that the inclusion of multi-DBDs could be more beneficial than using a single
DBD for delaying stall and to obtain a faster reattachment of the flow. However an optimal arrangement
of actuators in terms of position need to be selected in order to obtain a maximum multiplication effect
of the induced flow velocity. A control algorithm similar to the one of Lombardi et al.29 should also be
implemented in the steady actuation mode in order to shutdown the actuation in most of the downstroke
cycle where the DBDs were reducing the aerodynamic efficiency of the airfoil.
The unsteady actuation have demonstrated a superior performance on the downstroke phase of the cycle.
Although not being able to delay stall it was responsible for an increase of CLwithin a relevant portion of
the downstroke phase.
11 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
Acknowledgments
The present work was performed as part of Project CROP, supported by European Union within the 7th
Framework Programme under grant number 323047, and also supported by C-MAST, Centre for Mechanical
and Aerospace Science and Technology Research Unit No. 151.
References
1Sachse, H., Kirsten-Boeing Propeller, Tecnhical Memorandum Nr 351, National Advisory Committee for Aeronautics,
1926.
2Strandgren, C., The Theory of the Strandgren Cyclogiro, Tecnhical Memorandum Nr 727, National Advisory Committee
for Aeronautics, 1933.
3Wheatley, J. B., Simplified Aerodynamic Analysis of the Cyclogiro Rotating-Wing System, Tecnhical Note Nr 467,
National Advisory Committee for Aeronautics, 1933.
4Wheatley, J. B., Wind-Tunnel Tests of a Cyclogiro Rotor, Tecnhical Note Nr 528, National Advisory Committee for
Aeronautics, 1935.
5Nozaki, H., Sekiguchi, Y., Matsuuchi, K., Onda, M., Murakami, Y., Sano, M., Akinaga, W., and Fujita, K., “Research and
Development on Cycloidal Propellers for Airships,” 18th AIAA Lighter-Than-Air Systems Technology Conference, No. AIAA
2009-2850, American Institute of Aeronautics and Astronautics, Reston, Virigina, May 2009, pp. 1–5, doi:10.2514/6.2009–2850.
6Ilieva, G., P´ascoa, J. C., Dumas, A., and Trancossi, M., “MAAT - Promising innovative design and green propulsive con-
cept for future airship’s transport,” Aerospace Science and Technology, Vol. 35, 2014, pp. 1 – 14, doi:10.1016/j.ast.2014.01.014.
7Trancossi, M., Dumas, A., Xisto, C., P´ascoa, J., and Andrisani, A., “Roto-Cycloid Propelled Airship Dimensioning and
Energetic Equilibrium,” SAE 2014 Aerospace Systems and Technology Conference, No. 2014-01-2107, 2014.
8Iosilevskii, G. and Levy, Y., “Experimental and Numerical Study of Cyclogiro Aerodynamics,” AIAA Journal , Vol. 44,
No. 12, Dec. 2006, pp. 2866–2870, doi:10.2514/1.8227.
9Sirohi, J., Parsons, E., and Chopra, I., “Hover performance of a cycloidal rotor for a micro air vehicle,” Journal of the
American Helicopter Society, Vol. 52, No. 3, 2007, pp. 263–279, doi:10.4050/JAHS.55.022002.
10Hwang, S., Min, Y., Lee, H., and Kim, J., “Experimental Investigation of VTOL UAV Cyclocopter with Four Rotors,” 48th
AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, No. April, American Institute
of Aeronautics and Astronautics, Reston, Virigina, April 2007, pp. 1–10.
11Hwang, S., Min, Y., Lee, H., and Kim, J., “Development of a Four-Rotor Cyclocopter,” Journal of Aircraft, Vol. 45,
No. 6, Nov. 2008, pp. 2151–2157, doi:10.2514/1.35575.
12Benedict, M., Ramasamy, M., and Chopra, I., “Improving the Aerodynamic Performance of Micro-Air-Vehicle-Scale Cy-
cloidal Rotor: An Experimental Approach,” Journal of Aircraft, Vol. 47, No. 4, July 2010, pp. 1117–1125, doi:10.2514/1.45791.
13Benedict, M., Ramasamy, M., Chopra, I., and Leishman, J., “Performance of a Cycloidal Rotor Concept for Micro Air Ve-
hicle Applications,” Journal of the American Helicopter Society, Vol. 55, No. 2, 2010, pp. 022002, doi:10.4050/JAHS.55.022002.
14Benedict, M., Mattaboni, M., Chopra, I., and Masarati, P., “Aeroelastic Analysis of a Micro-Air-Vehicle-Scale Cycloidal
Rotor in Hover,” AIAA Journal, Vol. 49, No. 11, Nov. 2011, pp. 2430–2443, doi:10.2514/1.J050756.
15Benedict, M., Jarugumilli, T., and Chopra, I., “Effect of Rotor Geometry and Blade Kinematics on Cycloidal Rotor
Hover Performance,” Journal of Aircraft , Vol. 50, No. 5, Sept. 2013, pp. 1340–1352, doi:10.2514/1.C031461.
16Leger, J., P´ascoa, J., and Xisto, C., “Analytical Modeling of a Cyclorotor in Forward Flight,” SAE 2013 AeroTech
Congress & Exhibition, 2013, pp. doi:10.4271/2013–01–2271.
17Leger, J., P´ascoa, J., and Xisto, C., “Analytic Model Able to Assist in Parametric Design of Cycloidal Rotor
Thrusters,” Proc. ASME International Mechanical Engineering Congress and Exposition, No. IMECE2013-65238, 2013, pp.
doi:10.1115/IMECE2013–65238.
18Xisto, C. M., P´ascoa, J. C., Leger, J. A., Masarati, P., Quaranta, G., Morandini, M., Gagnon, L., Schwaiger, M., and
Wills, D., “Numerical Modelling of Geometrical Effects in the Performance of a Cycloidal Rotor,” 6th European Conference on
Computational Fluid Dynamcs, 2014.
19Gagnon, L., Quaranta, G., Morandini, M., Masarati, P., Xisto, C. M., and P´ascoa, J. C., “Aerodynamic and Aeroe-
lastic Analysis of a Cycloidal Rotor,” AIAA Aviation , American Institute of Aeronautics and Astronautics, June 2014, pp.
doi:10.2514/6.2014–2450.
20Benedict, M., Fundamental understanding of the Cycloidal-Rotor concept for micro air vehicle applications, Ph.D. thesis,
University of Maryland, 2010.
21Lee, T. and Gerontakos, P., “Investigation of flow over an oscillating airfoil,” Journal of Fluid Mechanics, Vol. 512, 2004,
pp. 313–341, doi:10.1017/S0022112004009851.
22Pruski, B. J. and Bowersox, R. D. W., “Leading-Edge Flow Structure of a Dynamically Pitching NACA 0012 Airfoil,”
AIAA Journal, Vol. 51, No. 5, March 2013, pp. 1042–1053, 10.2514/1.J051673.
23Spentzos, A., Barakos, G. N., Badcock, K. J., Richards, B. E., Wernert, P., Schreck, S., and Raffel, M., “Investigation of
Three-Dimensional Dynamic Stall Using Computational Fluid Dynamics,” AIAA Journal , Vol. 43, No. 5, May 2005, pp. 1023–
1033, doi:10.2514/1.J051673.
24Gleize, V., Costes, M., Pape, A. L., and Richez, F., “Numerical Simulation of a Pitching Airfoil Under Dynamic Stall
Conditions Including Laminar/Turbulent Transition,” Aerospace Sciences Meetings, American Institute of Aeronautics and
Astronautics, Jan. 2008, pp. doi:10.2514/6.2008–391.
12 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
25Wang, S., Ingham, D. B., Ma, L., Pourkashanian, M., and Tao, Z., “Turbulence modeling of deep dynamic
stall at relatively low Reynolds number,” Journal of Fluids and Structures, Vol. 33, No. 0, 2012, pp. 191 – 209,
doi:10.1016/j.jfluidstructs.2012.04.011.
26Gharali, K. and Johnson, D. A., “Dynamic stall simulation of a pitching airfoil under unsteady freestream velocity,”
Journal of Fluids and Structures, Vol. 42, No. 0, 2013, pp. 228 – 244, doi:10.1016/j.jfluidstructs.2013.05.005.
27Patel, M. P., Ng, T. T., Vasudevan, S., Corke, T. C., Post, M., McLaughlin, T. E., and Suchomel, C. F., “Scaling Effects
of an Aerodynamic Plasma Actuator,” Journal of Aircraft , Vol. 45, No. 1, Jan. 2008, pp. 223–236, doi:10.2514/1.31830.
28Post, M. L. and Corke, T. C., “Separation Control Using Plasma Actuators: Dynamic Stall Vortex Control on Oscillating
Airfoil,” AIAA Journal, Vol. 44, No. 12, Dec. 2006, pp. 3125–3135, doi:10.2514/1.22716.
29Lombardi, A. J., Bowles, P. O., and Corke, T. C., “Closed-Loop Dynamic Stall Control Using a Plasma Actuator,” AIAA
Journal, Vol. 51, No. 5, March 2013, pp. 1130–1141, doi:10.2514/1.J051988.
30Mitsuo, K., Watanabe, S., Atobe, T., Kato, H., Tanaka, M., and Uchida, T., “Lift Enhancement of a Pitching Airfoil
in Dynamic Stall by DBD Plasma Actuators,” 51st Aerospace Sciences Meetings , No. AIAA 2013-1119, American Institute of
Aeronautics and Astronautics, 2013, pp. 10.2514/6.2013–1119.
31Abdollahzadeh, M., P´ascoa, J. C., and Oliveira, P., “Numerical investigation on efficiency increase in high altitude propul-
sion systems using plasma actuators,” European Congress on Computational Methods in Applied Sciences and Engineering,
2012.
32Abdollahzadeh, M., Pascoa, J., and Oliveira, P., “Two-dimensional numerical modeling of interaction of micro-shock
wave generated by nanosecond plasma actuators and transonic flow,” Journal of Computational and Applied Mathematics,
Vol. 270, 2013, pp. 401–416, doi:10.1016/j.cam.2013.12.030.
33Abdollahzadeh, M., Pascoa, J., and Oliveira, P., “Modified Split-Potential Model for Modeling the Effect of
DBD Plasma Actuators in High Altitude Flow Control,” Current Applied Physics, Vol. In press, No. 0, 2014,
pp. doi:10.1016/j.cap.2014.05.016.
34P´ascoa, J. C., Xisto, C. M., and Gottlich, E., “Performance assessment limits in transonic 3D turbine stage blade
rows using a mixing-plane approach,” Journal of Mechanical Science and Technology, Vol. 24, No. 10, 2010, pp. 2035–2042,
doi:10.1007/s12206–010–0713–9.
35Xisto, C. M., P´ascoa, J. C., Oliveira, P. J., and Nicolini, D. A., “A hybrid pressure density-based algorithm for the Euler
equations at all Mach number regimes,” International Journal for Numerical Methods in Fluids, Vol. 70:8, 2012, pp. 961–976,
doi:10.1002/fld.2722.
36Xisto, C. M., P´ascoa, J. C., and Oliveira, P. J., “A pressure-based method with AUSM-type fluxes for MHD flows
at arbitrary Mach numbers,” International Journal for Numerical Methods in Fluids, Vol. 72:11, 2013, pp. 1165–1182,
doi:10.1002/fld.3781.
37Xisto, C. M., P´ascoa, J. C., and Oliveira, P. J., “A pressure-based high resolution numerical method for resistive MHD,”
Accepted for publication in Journal of Computational Physics , 2014.
38Suzen, Y., Huang, G., Jacob, J., and Ashpis, D., “Numerical Simulations of Plasma Based Flow Control Applications,”
35th AIAA Fluid Dynamics Conference and Exhibit, 2005.
39Bouchmal, A., Modeling of Dielectric-Barrier Discharge Actuator , Master’s thesis, Delft University of Technology, 2011.
40Menter, F. R., Langtry, R. B., Likki, S. R., Suzen, Y. B., Huang, P. G., and Volker, S., “A Correlation-Based Transition
Model Using Local Variables Part I: Model Formulation,” Journal of Turbomachinery, Vol. 128, No. 3, March 2004, pp. 413–422,
doi:10.1115/1.2184352.
41Langtry, R. B., Menter, F. R., Likki, S. R., Suzen, Y. B., Huang, P. G., and Volker, S., “A Correlation-Based Transition
Model Using Local Variables Part II: Test Cases and Industrial Applications,” Journal of Turbomachinery, Vol. 128, No. 3,
March 2004, pp. 423–434, doi:10.1115/1.2184353.
42Langtry, R. B. and Menter, F. R., “Correlation-Based Transition Modeling for Unstructured Parallelized Computational
Fluid Dynamics Codes,” AIAA Journal , Vol. 47, No. 12, Dec. 2009, pp. 2894–2906, doi:10.2514/1.42362.
43P´ascoa, J. C., Dumas, A., Trancossi, M., Stewart, P., and Vucinic, D., “A review of thrust-vectoring in support of a
V/STOL non-moving mechanical propulsion system,” Central European Journal of Engineering, Vol. 3, No. 3, 2013, pp. 374–
388, doi:10.2478/s13531–013–0114–9.
44Whalley, R., Turbulent Boundary-layer control with DBD Plasma actuators using spanwise travelling-wace technique,
Phd thesis, University of Nottingham, 2011.
13 of 13
American Institute of Aeronautics and Astronautics
Downloaded by Carlos Xisto on July 28, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-3854
