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Effect of Longitudinal Ridges on the Aerodynamic
Characteristics of an Airfoil
Tufan Kumar Guha1, Erik Fernandez
2and Rajan Kumar
3
Florida Center for Advanced Aero-Propulsion
FAMU-FSU College of Engineering, Florida State University, Tallahassee, FL - 32310
A pair of longitudinal ridges is employed to enhance the aerodynamic efficiency of an
USA-35B airfoil used in a remotely controlled aircraft wing. Measurements include surface
static pressure distributions, 2-dimensional and stereo particle image velocimetry. In the
present study, the height and inter-ridge spacing were varied to find optimal control
configuration for maximum effectiveness. The results show that the baseline airfoil exhibits a
massive flow separation on the suction side at an angle of incidence of 12 and beyond. With
the addition of optimal longitudinal ridges, the stall angle was significantly delayed and the
flow was completely attached up to an angle of incidence of 16 . These results clearly show
the benefit of using longitudinal ridges as an effective flow control technique for improving
aerodynamic characteristics of an aircraft wing.
NomenclatureRec = Reynolds number based on airfoil chord length
Cp = Pressure coefficient (= (ps - p)/q)
PS = Local static pressure on the airfoil surface
P = Freestream static pressure
P0 = Freestream total pressure
q = Freestream dynamic pressure
D = Diameter of the longitudinal ridges
S = Inter-ridge spacing
U = Ensemble-averaged (mean) steamwise velocity
Uin = Freestream velocity (=20m/s)
I. IntroductionVER the years, many different methods for enhancing the aerodynamic performance of a wing have been
proposed, verified by wind-tunnel and flight tests and eventually incorporated in the flight vehicle. An airfoil
section, an essential part of a wing, has its primary task as a lift generator, therefore, an optimized design of the
airfoil is the prerequisite to the satisfactory performance of the lifting surface. It has always been a challenge of
optimizing and enhancing its lift without seriously increasing the drag. Also, it is now well known that the flow
separation over an airfoil at high angle of attacks severely degrades its performance, particularly reduction in lift and
increase in drag. To counteract this problem a wide variety of active1-5
and passive6-8
flow control devices have been
developed and researched upon and undoubtedly they help to mitigate the problem to a certain extent. But a lot of
these techniques can be termed as empirical, since there is a lack of complete understanding of the underlying
physics of interaction between the device and the flow. One of the passive control techniques that has been studied
in depth is the use of ribblets for skin friction drag reduction. Riblets9-10
, which are micro-grooves on the surface and
aligned to the freestream direction have shown a viscous drag reduction in the range of 48% on a variety of two-dimensional flows with zero or mild pressure gradients at subsonic speeds. However, the performance of these
devices in adverse pressure gradients and at supersonic speeds is questionable. An experimental study conducted by
Zverkov et al.14
have shown that a wavy wing surface affect the location of the transition point in the boundary layer
and reduces the size of separation bubble. Motivated by the techniques of drag reduction observed in aquatic
animals, Fish et al.11
studied the use of leading edge tubercles (typically found in humpback whale flipper) for lift
1Graduate Research Assistant, Department of Mechanical Engineering, Student Member AIAA.
2Graduate Research Assistant, Department of Mechanical Engineering, Student Member, AIAA
3Assistant Professor, Department of Mechanical Engineering, Senior Member AIAA.
O
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enhancement and drag reduction. The leading edge tubercles may function to generate streamwise vortices by
excitation of flow to maintain lift and prevent stall at high angles of attack. Another approach to increase the lift at
moderate angles of attack has been to somehow trap the vortex on the suction side of the wing. Rossow12
at NASA
Ames through his 2D inviscid, incompressible calculations proposed the use of two spanwise wing fences to trap the
vortex.
The aim of our research is to focus on the understanding of the mechanisms associated with the stall
characteristics, particularly flow separation at high angles of incidence and develop a flow control technique to
enhance the aerodynamic performance of an airfoil over a wide range of Reynolds number. In the present study we
have used USA-35B airfoil (Fig. 2) and characterized its performance over a range of test conditions. We have
introduced a flow control method by attaching longitudinal (chordwise) thin ridges, near the mid-section of the
airfoil. We anticipate that the placement of chordwise ridges along the airfoil section will have multiple flow control
effects in terms of restricting the spanwise vortex movement, generation of streamwise vorticity and energizing the
flowfield due to entrainment near the surface. Measurements include surface static pressure distributions and particle
image velocimetry. The initial results of this control technique are very encouraging in terms of increase in
maximum lift coefficient, reduction in size or even elimination of separation bubble and a delay in stall angle.
II. Test Facility and MeasurementsA. Low Speed Wind Tunnel
Measurements were made in the low speed wind tunnel at the Florida Center for Advanced Aero propulsions
(FCAAP) at the Florida State University. The facility is a closed loop wind tunnel with continuous operation over
extended period. The test section is manufactured with acrylic walls and has dimensions of24 x 24 x 60, which
allows for flow visualization from all the four sides. It is driven by a 240 HP fan which is controlled with a variable
frequency drive (VFD; Toshiba H9425KAA) and a precision regulator. The precision regulator is used to
manipulate the fan pitch angle while maintaining a particular fan RPM with the help of the VFD. The flow travels
through a heat exchanger and a flow straightener section which allows for constant flow temperature and minimum
flow fluctuations. Accurate velocities can be obtained between a range of 2 90 m/s with a test section freestream
turbulence intensity of 0.5% at 20m/sec. The present tests were carried out at a freestream velocity of 20m/sec and
corresponding Reynolds number of 5.1 x 105
based on a chord length of 14.45 inches. Wind tunnel velocity is
measured by a pitot-static probe mounted 0.75 m upstream of the test section and is monitored using a 0 - 0.1 in H2O
Omega differential pressure transducer with 0.5% full scale accuracy.
B. Test ModelUSA-35b AirfoilThe airfoil studied is a section of a USA-35b wing used in a RC plane. The span of the airfoil section is 24, so
that it is flushed with the walls of the test section and its chord length is 14.45 (Fig. 2). The airfoil is mounted at a
distance of 8 downstream of the beginning of the test section and 12 from the floor of the wind tunnel , as shown in
in Fig. 1.
a) Subsonic windtunnel facility b) Mounting of Airfoil
Fig. 1. Experimental Setup
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To measure the surface static pressure distributions, 31 pressure ports of 0.02 diameter are drilled at the mid -
section of the airfoil. The airfoil can be rotated with respect to the freestream and fixed at a chosen angle of attack in
the range of -4 to 20. The initial tests were carried out on the baseline airfoil and later two ridges on the two sides
of pressure generator were introduced to study the effect of control (Fig. 2a). The separation distance (S) between
these two ridges was varied between 1 8 (Fig. 2b). The ridges covered the airfoil surface up to 90 % of chord
length on both the suction and the pressure side. This was done so that the airfoil maintains the characteristic of a
sharp trailing edge. The longitudinal ridges were made of two flexible circular tubes and were mounted on the airfoil
surface along its contour (Fig. 1). The diameter of the circular tubes was also varied from 0.05 0.2 in theseexperiments (Fig. 2b). These tubes were covered with a smooth tape to achieve a nearly Gaussian surface profile and
to avoid non-uniformity near the surface. The non dimensionalized height (D/C) of the ridges with respect to the
chord length is 0.0035 to 0.014. The incoming boundary layer thickness () at quarter-chord was estimated to be
approximately 0.15-inch and the corresponding non-dimensionalized /C is equal to 0.01. Table 1 shows the set of
experimental test conditions.
Parameter Symbol Parametric Values
Angle of Attack 0, 12, 14, 16
Diameter of RidgeD 0, 0.05, 0.1, 0.2
D/C 0, 0.0035, 0.007, 0.014
Inter-ridge Spacing
S 0, 1, 2, 4, 6, 8
S/C0.07, 0.14, 0.28, 0.42, 0.56
D
x/C = 0.9C = 14.45
S
b = 24
Fig. 2. Airfoil with longitudinal ridges
a) Experimental wing section b) Schematics of test model
Table 1. Experimental test conditions
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a)2D PIV measurement plane
b) Stereo PIV measurement plane
Fig. 3. PIV measurement locations on the airfoil section
C. MeasurementsMeasurements include surface static pressure distributions and particle image velocimetry at the mid chord of the
airfoil. The Reynolds number based on chord length at a freestream velocity of 20 m/s is 5.1 x 105. The angle of
incidence was varied from 0 - 20. The static pressures were measured with a Omega 0-0.3 psid pressure
transducer (Model PX138) with a full scale error of 1%. The pressure data was sampled (1000 samples) at each port
at 100 Hz for 10 secs to ensure flow stabilization and minimize statistical errors.
Particle Image Velocimetry (PIV) was used to obtain quantitative measurements of the flow field of interest.
Two dimensional planar PIV and three dimensional stereoscopic PIV were employed in this study. For all planar 2D
PIV cases presented, measurements were carried out at the airfoil mid-span (x-y plane) and the flow was examined
over the aft 60% of the airfoil to capture the
separation effects. Stereo PIV cross planes (y-
z plane) at selected chordwise locations
covered a center span of approximately 0.52
z/C. Figure 3 shows the PIV measurement
locations and extent of laser sheet on the
airfoil section. The flow was illuminated by a
pulsed Quantel
Nd:YAG laser triggered at a
specified time interval. The beam is focused
using a single spherical lens of appropriate
focal length and the laser sheet is created when
the beam passes through two cylindrical lenses(two lenses were used in order to obtain a
sheet wide enough to cover the area of
interest).
The air flow is seeded using a ROSCO
fog machine and introduced into the wind
tunnel upstream of the flow straighteners. The
seed particles are approximately 2-5m in
size. The time interval between laser pulses
was approximately 30s for the presented
cases. Image pairs were acquired at 15Hz with
a resolution of 2560 x 2160 pixels yielding a
spatial resolution of approximately 12.4pixels/mm. Each PIV test case shown here consists of 500 ensemble averaged instantaneous image pairs in order to
estimate the mean statistics of interest. Images were acquired and processed using LaVision Davis 8 software with a
5.5 megapixel sCMOS camera equipped with a 55mm focal length lens. Velocity correlations were made using a
final adaptive interrogation window size of 48 x 48 pixels with a decreasing size multipass algorithm. All passes
used a 50% window overlap. The measurement uncertainty is estimated to be about 1% in ensemble-averaged
velocity measurements with a 95% confidence level.
III. Results and DiscussionAs briefly mentioned in the introduction, the main objective of the present study is to characterize the airfoil near
stall conditions particularly measure the effect of control on stall angle, the location and the reduction in the extent
of flow separation. Here we present results in the form of surface static pressure distributions, 2D velocity field and
a few stereo PIV measurements at selected conditions.
FLOW
x/C z/b
A 0.5 0.4
B 0.5 0.7
A
B
x
zy
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A. Surface Static Pressure DistributionsSurface static pressure distributions are presented in the form of coefficient of pressure Cp= (p - p)/q, (Where
p
is the free-stream static pressure and qis the free-stream dynamic pressure) as a function of non-dimensionalized
chordwise distance, x/C. Figures 4-5 show the effect of longitudinal ridges at 0, 12 and 14 angles of attack,
respectively. As may be seen from these pressure distributions that for the baseline airfoil, the flow is attached at 0
angle of incidence (Fig. 4) and remain attached up
to angle of 11
(not shown here). The flow isseparated at 12 (Fig. 5) with a large separation
bubble as indicated by a plateau region on the
suction side of the airfoil. At higher angles the
size of separation bubble is further increased and
larger area of airfoil is subjected to reverse flow.
With the addition of two longitudinal ridges (D/C
= 0.008, S/C = 0.14) on the surface of airfoil, it is
clearly observed that there is an increase in the
value of suction peak at all angles of incidence
and the effects are more pronounced at higher
angles of incidence, particularly near the stall
angles. The plateau region corresponding to
separation bubble is modified to attached flow
profile. From these pressure distributions it
appears that with the addition of longitudinal
ridges the flow is attached up to 14 angle of
attack. As expected the flow on the pressure side
of the airfoil remain unaffected with ridges at these angles of incidence.
B. Velocity FieldParticle image velocimetry (PIV) was carried out at a few selected conditions to obtain 2D velocity field along
the streamwise direction perpendicular to the airfoil surface. The PIV window was selected from x/C of 0.45 to 1.03
to capture the reverse flow in the separation bubble.
Fig. 5. Effect of ridges (D/C = 0.008 & S/C = 0.14) on surface static pressure distributions
Fig. 4. Surface static pressure distributions at = 0
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a)Baseline airfoil, = 0 b) Baseline airfoil, = 12
c)Baseline airfoil, = 14 d) Baseline airfoil, = 16 Fig. 6. Effect of angle of incidence on streamwise velocity distribution
a)Ridge height D/C = 0.007 b) Ridge height D/C = 0.014Fig. 7. Effect of ridge height on streamwise velocity distribution, S/C = 0.14, = 14
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a)Ridge Spacing S/C = 0.07 b) Ridge Spacing S/C = 0.014
c)Ridge Spacing S/C = 0.028 d) Ridge Spacing S/C = 0.042
e) Ridge Spacing S/C = 0.056
Fig. 8. Effect of inter-ridge spacing on streamwise velocity distribution, S/C = 0.14, = 14
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a) Baseline airfoil b) With longitudinal ridgesFig. 10. Effect of longitudinal ridges (D/C = 0.014 & S/C = 0.42) on streamwise vorticity
i. Baseline FlowfieldFigure 6 shows the effect of angle of attack on the streamwise velocity contours for the baseline airfoil. The
arrowheads in the streamlines indicate the direction of flow.As expected, the flow is completely attached at = 0(Fig. 6a) and the surface streamlines follow the contour of the airfoil. An increase in angle of incidence to = 12
(Fig. 6b), the flow is separated on the airfoil suction side approximately at x/C = 0.65 and the separation bubbleshows an open separation. With further increase in angle of incidence to 14 and then 16 (Figs 6c and 6d), the size
of separation bubble is increased and the separation location is moved upstream leading to massive stall (indicated
by reverse flow) characteristics. Due to the presence of separation bubble, the streamlines in the freestream flow
have been pushed away from the surface.
ii. Airfoil with Longitudinal RidgesTwo chordwise longitudinal ridges were placed on the airfoil near the mid-section on either side of the airfoil
centerline. As mentioned in the experimental section the height of ridges and their inter-spacing was varied in this
study. Figure 7 shows the effect of the ridge height in controlling the separation on the airfoil for a fixed inter-ridge
a) Baseline airfoil b) With longitudinal ridges
Fig. 9. Effect of longitudinal-ridgs on streamwise velocity distribution, D/C = 0.014,
S/C = 0.14, = 16
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spacing of S/C = 0.14 at an angle of incidence of 14. The dotted line indicates the airfoil surface and the solid line
represent the new surface due to chordwise ridges. The results clearly show that in comparison to baseline flow at
= 14(Fig. 6c), even a small thickness ridges (D/C = 0.007) reduces the size of separation bubble and pushes the
separation location downstream. With an increase in ridge height to D/C = 0.014, the flow on the entire surface is
completely attached. In respect to the boundary layer thickness estimated at quarter-chord ( = 0.15), the small
thickness ridges are well within the boundary layer whereas a ridge height of D/C =0.014 is slightly outside.
The effect of inter-ridge spacing on the streamwise velocity is shown in Fig. 8. As observed, when the ridges aretoo close to each other (Fig. 8a), the effectiveness in terms of separation reduction is minimal, if any. As the spacing
between the two ridges was increased to S/C = 0.14 (S = 2) and 0.28 (Figs. 8b and 8c), its effectiveness improved
and the separation was eliminated. With further increase in spacing to S/C = 0.42 (Fig. 8d), the results were even
better in terms of completely attached flow, higher velocities near the airfoil surface and reduction in boundary layer
thickness. However, when the spacing was further increased to S/C = 0.56 (Fig. 8e), a small separation bubble was
observed near the trailing edge and the flow velocity near the surface was reduced. These results clearly indicate that
there are optimal values of ridge height and inter-ridge spacing for which the effectiveness is the best. The optimal
case of D/C = 0.014 and S/C = 0.42 was further tested at higher angles of incidence of 16 (Fig. 9) and the results
show attached flow features. These results are consistent with those observed in pressure distributions, confirming
the effectiveness of longitudinal ridges in enhancing the performance of airfoil near stall angles.
iii. Stereo Particle Image VelocimetryIn order to better understand the mechanisms responsible for the effectiveness of longitudinal ridges in
eliminating separation near stall angles, the flowfield was further investigated using stereo particle image
velocimetry at a few selected spanwise planes (Fig. 4). Figure 10 shows the contour plots of streamwise vorticity for
the baseline flowfield and with longitudinal ridges (D/C = 0.014 and S/C = 0.42) at = 14. It may be observed that
the longitudinal ridges significantly alter the flowfield close to the surface through the production of streamwise
vorticity and energizing the boundary layer. These results will be further analyzed to understand the flow physics
associated with this control scheme.
IV. Concluding RemarksAn experimental study measuring the effect of longitudinal ridges on the aerodynamic characteristics of an
airfoil is presented. Surface static pressure distributions and particle image velocimetry are used to measure the flow
control effectiveness. The results show that two chordwise ridges of optimal height and inter-ridge spacing can
eliminate the separation on the suction side of this airfoil. This will result in a significant lift enhancement and drag
reduction at these moderate angles of incidence and a delay in its stall angle by about 4 - 5. These results alsoindicate that there may be multiple flow control mechanisms responsible for the effectiveness of this flow control
technique.
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