Effect of boundary layer thickness on secondary structures in a short inlet curved duct

Effect of boundary layer thickness on secondary structures in a short
inlet curved duct
Jeremy Gartner 1
, Michael Amitay 2,⇑
Rensselaer Polytechnic Institute, Troy, NY 12180, USA
a r t i c l e i n f o
Article history:
Received 4 February 2014
Received in revised form 15 October 2014
Accepted 17 October 2014
Available online 13 November 2014
Keywords:
S-shape inlet
Secondary flows
Three dimensional separation
Secondary flow structures
a b s t r a c t
The flow pattern in short inlet ducts with aggressive curvature has been shown to lead, in some cases, to
an asymmetric flow field at the aerodynamic interface plane. In the present work, a two-dimensional
honeycomb mesh was added upstream of the curved duct to create a pressure drop across it, and there-
fore to an increased velocity deficit in the boundary layer. This velocity deficit led to a stronger stream-
wise separation, overcoming the instability that can result in an asymmetric flow field at the
aerodynamic interface plane. Experiments were conducted at Mach numbers of M = 0.2, 0.44 and 0.58
in an expanding aggressive duct with rectangle to a square cross section with area ratio of 1.27. Steady
and unsteady pressure measurements, together with Particle Image Velocimetry (PIV), were used to
explore the effect of the honeycomb on the symmetry of the flow field. The effect of inserting a honey-
comb was tested by increasing its height from 0 to 2.2 times the boundary layer thickness of the baseline
flow upstream of the curve. Using the honeycomb, flow symmetry was achieved for the specific geomet-
rical configuration tested with a negligible decrease of the pressure recovery.
Ó 2014 Elsevier Inc. All rights reserved.
1. Introduction
Heightened interest in short and curved inlet ducts for aircraft
has led us to further explore the flow field existing in such devices
as well as to a better knowledge of the issues associated with their
design and methods of mitigating such issues. Several factors
related to engine and aircraft performance and operation drive
the use of short inlet duct designs, such as the overall airframe
length reduction enabled by a shorter duct and reduction of frontal
planform by burying the engine into the airframe. These factors led
to a reduction in weight and fuel consumption, and allowed for
innovative external and integrated aerodynamics, such as Blended
Wing Bodies (Dagget et al., 2003). Another factor that must be
taken into consideration is the stability margin for operation of a
jet engine following the duct, where uneven pressure distribution
and secondary flow structures can lead to engine stall at the fan/
compressor stages (surge stall) (Scribben et al., 2006; Mattingly
et al., 2002).
A considerable body of work is available in the literature con-
cerning the analysis of the flow field in short inlet ducts (Bansod
and Bradshaw, 1972; Launder and Ying, 1972; Enayet et al.,
1982; Wellborn et al., 1992, 1993; Whitelaw and Yu, 1993a,b; Ng
et al., 2008, 2006). These previous research efforts have shed light
on the main features of the flow field existing in aggressively
curved ducts, where the rapid curvature in the duct results in pres-
sure gradients in the direction normal to the turn, leading to the
onset of secondary flow structures in the form of two counter
rotating vortices. Other structures were also noticed to co-exist
with these counter-rotating vortices, such as cross stream flow at
the internal surfaces, which invade the local boundary layer lead-
ing to further flow detachment (Wellborn et al., 1992, 1993; Ng
et al., 2008, 2006; Chen, 2012) disrupting the flow and creating
recirculation zones in the duct. The symmetric counter-rotating
vortices can be described by inviscid flow equations, caused solely
by the turning of the flow. These pressure driven counter-rotating
vortices convect the low momentum fluid of the boundary layer
towards the center of the duct impacting flow uniformity and pres-
sure recovery at the face of the engine located downstream, at the
aerodynamics interface plane.
Implementation of passive and active flow control techniques in
short inlet ducts has been an active field of research (Scribben
et al., 2006; Ng et al., 2008, 2006; Chen, 2012; Vaccaro, 2011;
Debronsky, 2012; Amitay et al., 2002; Gissen et al., 2011; Jirasek,
http://dx.doi.org/10.1016/j.ijheatfluidflow.2014.10.016
0142-727X/Ó 2014 Elsevier Inc. All rights reserved.
⇑ Corresponding author.
E-mail address: amitam@rpi.edu (M. Amitay).
1
Department of Mechanical, Aerospace & Nuclear Engineering, 110 8th Street,
Troy, NY, USA.
2
James L. Decker’45 Endowed Chair of Aerospace Engineering, and Director of the
Center for Flow Physics and Control (CeFPaC), 110 8th Street, Troy, NY, USA.
International Journal of Heat and Fluid Flow 50 (2014) 467–478
Contents lists available at ScienceDirect
International Journal of Heat and Fluid Flow
journal homepage: www.elsevier.com/locate/ijhff

2006; Reichert and Wendt, 1994). The predominant forms of actu-
ation are vortex generators, steady and unsteady jet blowing tan-
gent to the surface, synthetic jet actuators, and many more.
Recent work has studied multiple actuation devices (Vaccaro,
2011; Gissen et al., 2011) including combination of flow control
techniques. Although most of the previous work was focused on
circular cross section ducts (Wellborn et al., 1992, 1993;
Whitelaw and Yu, 1993a,b; Gissen et al., 2011; Jirasek, 2006),
emphasis was also given to rectangular cross section ducts
(Launder and Ying, 1972; Ng et al., 2008, 2006; Chen, 2012;
Vaccaro, 2011; Debronsky, 2012; Amitay et al., 2002). All of the
work performed with flow control had the objective of improving
the pressure recovery and pressure distribution at the exit of the
duct.
Recent experiments (Vaccaro, 2011; Debronsky, 2012) and
numerical simulations (Chen, 2012) have shown that, under some
geometrical conditions, the flow can become asymmetric, where
one of the counter-rotating vortical structure supersedes the other.
In the case of the rectangular ducts, the secondary flow structures
were shown to move towards one of the corners of the duct (Chen,
2012; Vaccaro, 2011; Debronsky, 2012).
As noted by Chen (2012), the secondary flow phenomenon
(i.e., a turbulent flow with mean streamwise vorticity) is attrib-
uted to two mechanisms: (i) the skew induced, inviscid mecha-
nism, which is caused by any bend in the flow path of ducts
with any cross sectional shape as shown by Miller (1991) (and
Fig. 1 below), and (ii) a stress-induced mechanism occurring in
any non-circular ducts, straight or not, due to anisotropy of the
Reynolds stresses. A more in-depth description of secondary flow
can be found in Perkins (1970) and Bradshaw (1987). Also note
that further complexity in the flow structures is due to swirl
development in the second bend of the s-duct. This reverse in
the curvature is accredited with the crossover of the transverse
velocity component near the side walls, an essentially inviscid
process. Another feature of short inlet ducts is the adverse pres-
sure gradient caused by the opposite curvature of the second
bend. Therefore, the secondary flow generated by the first bend
is attenuated, being reversed depending on the aggressiveness
of the turn (i.e., the aspect ratio L/D, the offset and area ratio
between the inlet and the exit sections).
Fig. 2a (taken from Wellborn et al., 1993) shows a three-dimen-
sional perspective of the owl face separation topology with a coun-
ter-rotating pair of vortices orientated with upwelling along the
centerline. The skeleton drawing of Fig. 2b shows the schematic
of a symmetric but unstable owl face separation of the first kind.
Perry and Hornung (1984) suggested that this unstable symmetric
distribution could exist due to slight variations in the flow field.
However, they stated that it was a special condition and that any
asymmetry in the flow would cause the streakline pattern to shift
to one side as shown in Fig. 2c.
Based on previous experiments (Ng et al., 2008, 2006; Vaccaro,
2011) and numerical analysis (Chen, 2012), it was shown that a
critical length to diameter ratio exists in a rectangular cross section
compact inlet duct, controlling the asymmetry of the secondary
structures. For ducts longer than this critical length the flow pat-
terns are asymptotically stable. With the reduction of duct length
below this critical value, the streamwise pressure gradient
increases and interacts with the transverse invasion. Simulta-
neously, the forward moving main flow confronts the backflow
to the streamwise separation. The symmetric pattern becomes
unstable due to the saddle–saddle connections existing in the
topology of the flow, leading to a flow bifurcation (Tobak and
Peak, 1982).
This asymmetric configuration is the starting point for the cur-
rent work, which also derives from the observation (Enayet et al.,
1982; Chen, 2012) that the secondary flow structures forming on
the inside of duct bends have their strength dependent on inlet
flow conditions, specifically the momentum thickness of the
Nomenclature
AIP aerodynamic interface plane
D width of the duct (mm)
h honeycomb height (mm)
d boundary layer thickness (mm)
h/d relative height of the honeycomb
h momentum thickness (mm)
L length of the duct (mm)
Minlet Mach number at the inlet
PR pressure recovery
Pinlet static pressure at the inlet (atm)
c specific heat ratio of air
Cp pressure coefficient
P0 total pressure (atm)
P static pressure (atm)
Fig. 1. Development of secondary flows in a pipe bend showing the (a) presence of an adverse pressure gradient, and (b) direction and orientation of the secondary flow (from
Miller, 1991).
468 J. Gartner, M. Amitay / International Journal of Heat and Fluid Flow 50 (2014) 467–478

Fig. 2. Owl face separation of the ﬁrst kind topology showing the (a) three-dimensional perspective of the separation, (b) symmetric skeleton schematic of streaklines, and (c)
skeleton schematic of asymmetric streakline pattern (from Wellborn et al., 1993).
Fig. 3. The experimental facility.
Fig. 4. Exploded view of the inlet duct assembly.
∗
Fig. 5. Inlet duct with the Instrumentation Can for AIP measurements.
Fig. 6. Pressure sensors distribution at the AIP.
J. Gartner, M. Amitay / International Journal of Heat and Fluid Flow 50 (2014) 467–478 469

incoming boundary layer. Therefore, the objective of the current
work is to show that proper manipulation of the incoming bound-
ary layer can lead to modification of the flow and pressure fields
without incurring any significant pressure loss. The main gain
expected from the manipulation of the incoming boundary layer
is on the cross flow symmetry of the pressure distribution at the
AIP.
Honeycomb Mesh
Quarter of a dollar
Fig. 7. Compact inlet with the honeycomb inserted.
Fig. 8. Internal view with the honeycomb inserted.
Fig. 9. Inlet total pressure field measured at M = 0.44; (a) across the duct height,
and (b) across the duct span.
Fig. 10. Cross-stream distributions of the streamwise velocity profiles upstream of
the turn at M = 0.44.
Fig. 11. Effect of the honeycomb height on the shape of the velocity profile at x/
D = À0.58.
Fig. 12. Effect of the honeycomb height on the shape of the velocity profile at x/
D = À0.14.
Table 1
Momentum displacement at x/D = À0.58.
h/d0 h (mm)
0 0.55
0.74 0.95
1.47 1.04
2.2 1.16

2. Experimental setup
The high subsonic facility at the Center for Flow Physics and
Control (CeFPaC) at Rensselaer Polytechnic Institute was utilized
in the current experiments. The high subsonic facility is a blow
down, open return wind tunnel. The inlet duct geometry allowed
for Mach numbers up to 0.6 although most data presented was
conducted at a Mach number of 0.44 (mass flow rate up to
1.76 kg/s). Fig. 3 shows the flow path of the facility and labels each
component. The air begins in the blower and transitions through
the blower diffuser into the settling chamber. Next, it enters con-
traction followed by the inlet duct (i.e., the test section), and finally
exits the facility through a diffuser.
The blower used is a Cincinnati Fan model HP-12G29 run by a
100 HP motor that is controlled by a variable frequency drive.
The blower can produce a volumetric flow rate up to 170 m3
/
min. The air exits the blower and enters the diffuser section that
transitions the circular cross-section of the blower to the square
cross-section of the settling chamber. The air is slowed as it enters
the settling chamber where the fluid is conditioned through a set
of screens and honeycomb. In addition, a thermocouple and a static
pressure ring were monitored for all experiments and were used in
the calculation of the Mach number. The air then enters the con-
traction section with a contraction ratio of 142:1 and a conven-
tional 5th order polynomial curvature.
The air then enters the inlet duct, which has a constant cross-
section with a length of 0.3048 m for boundary layer growth as
well as to measure the inlet Mach number. Another static pressure
ring, as well as a thermocouple, was instrumented in this constant
cross-section section. Utilizing a one-dimensional isentropic flow
assumption, the inlet Mach number was found from the static
pressure ring in the inlet as well as the static pressure ring of the
settling chamber. This assumption was validated in the previous
work done by Vaccaro (2011) utilizing a total pressure Kiel probe
in the constant cross-section region of the inlet duct. The total
pressure of the inlet nearly matched the pressure measured in
the static pressure ring of the settling chamber, where the air
velocity is small enough to be assumed to be zero allowing the
pressure and temperature measured in the settling chamber to
be taken as the total quantities. From the total quantities measured
in the settling chamber and the static pressure measured in the
inlet, the Mach number from the isentropic flow assumption is:
Minlet ¼
ffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi
P0
Pinlet
cÀ1
c
À 1
#
Ã
2
c À 1

v
u
u
t ð1Þ
Following the constant area section of the inlet is the diffusing
S-shaped section. After the air goes through the inlet duct, it exits
the wind tunnel through a diffuser. This diffuser angle is 3° to
reduce flow speed as it exits into the open room.
The inlet duct has a length-to-diameter ratio of 1.6, where the
initial rectangular cross-section area of 90 mm tall by 114.3 mm
wide transitions to a square cross-section 114.3 mm by
114.3 mm resulting in an area ratio of Aexit/Ainlet of 1.27. The design
Fig. 13. Streamwise pressure coefficient distribution for the three different honeycomb actuator heights at M = 0.2.

was modular to allow for easy removal and exchange of parts as
can be seen in Fig. 4, which shows an exploded view of the inlet
duct. For example, two sets of windows were fabricated: one set
made of aluminum for most of the experiments and the other
out of optical grade acrylic solely for the use of PIV. This reduced
the chance of scratches occurring on the windows that would have
a negative impact on the PIV quality and potentially causing glares
and/or inaccuracies in the collected data.
The inlet duct was instrumented with 126 static pressure taps
along the lower surface of the duct. The spanwise density of static
taps increased towards the side-wall with an array on the opposite
side for checking symmetry (see Fig. 4). The static pressure taps
were sampled by the means of four pressure scanners (Scanivalve
DSA3217, 16 channels each, ±5 psid full scale and accuracy of
±0.05% of full scale, or ±0.0025% psi). Because only 64 ports could
be sampled at a time, two runs were required to sample all of
the static pressure taps.
The static pressures measured from these taps were used to cal-
culate the pressure coefficient CP at each location. The definition of
CP used for this calculation is:
CP ¼
2
cM2
inlet
P
Pinlet
À 1

ð2Þ
The coordinate systems is defined such the origin is at the mid-
dle of the cross-section plane at the beginning of the curvature. x is
the streamwise direction, y is the cross-stream direction (i.e., nor-
mal to the surface) and z is along the span. Another axis was
defined as y⁄
in order to quantify the boundary layers, where y⁄
is aligned with y, but is shifted such that it is defined to be zero
on the floor (see Fig. 5).
In addition to the static surface pressure sensors, an Instrumen-
tation Can was mounted at the exit of the inlet duct to acquire the
total pressure field at the aerodynamic interface plane (AIP). Fig. 5
shows the Instrumentation Can mounted to the inlet duct with
eight rakes each containing five sensors (a total of 40 sensors). Each
sensor location on the rake has a high frequency pressure trans-
ducer (Kulite model XCQ-062, ±5 psid full scale and repeatability
of ±0.5% of full scale, or ±0.025% psi) as well as a steady total pres-
sure tap that can be sampled with the four Scanivalves. The num-
bering scheme of the Kulite sensors at the AIP is shown in Fig. 6.
The perspective of Fig. 6 is the same orientation as Fig. 5 where
the observer is looking downstream toward the AIP.
The 40 Kulites at the AIP were utilized to measure the steady
and unsteady pressure. The steady pressure was used to plot con-
tours maps of the pressure recovery, which is the total pressure at
the AIP normalized by the total pressure at the inlet. The average
pressure recovery, PRavg, and the lower half pressure recovery,
PRlowerhalf, are defined respectively as
PRavg ¼
ð
P40
1 P0;iAIPÞ=40
P0inlet
ð3Þ
PRlowerhalf ¼
ð
P35
21P0;iAIPÞ=15
P0inlet
ð4Þ

where the Kulites are numbered as represented in Fig. 6.
Fig. 7 shows the compact inlet with the honeycomb inserted
inside it. In order to provide some proportional perspective of
the experiment’s scaling, a quarter of US dollar is situated next
to the honeycomb mesh. An additional internal view of the tunnel
is presented in Fig. 8, where the flow direction is out to the page
with the same quarter displayed in Fig. 7.
In this work, Particle Image Velocimetry (PIV) measurements
were conducted in order to collect quantitative flow measurement
of the velocity field at different streamwise and spanwise planes
within the duct. All PIV experiments conducted in this work uti-
lized a commercial LaVision System of software, and TSI hardware.
The hardware included two 120 mJ Nd:YAG lasers and a single
1000 Â 1016 pixel resolution TSI CCD camera. Additional hardware
was a Martin Magnum 850 fog machine utilized for seeding the
flow with O(1 lm) smoke particles. An array of different optics
included a cylindrical lens utilized to create the laser light sheet,
focal lens to focus the sheet within the measurement field, and
camera optics for focusing the camera at different focal lengths
and desired fields of interest. Also, the camera and the laser were
mounted on computer-controlled traverses to provide a precise
location of the measurement planes.
The velocity components were computed by the DaVis 7.1 soft-
ware using the cross-correlation technique of pairs of successive
images with 50% overlap between the interrogation windows.
The successive images were processed using a multi-pass method
in which the initial and final passes were 32 Â 32 pixels and
16 Â 16 pixels, respectively. Lastly, the averaged velocity field in
each plane was averaged over 500 instantaneous velocity fields.
The maximum velocity (approximately 150 m/s) corresponds to
an average displacement of about 4 pixels with an approximate
maximum error of ±0.1 pixel, which corresponds to an error of
±2.5% of the inflow velocity.
3. Results
The results section is divided into three sections: (i) upstream of
the curved duct, (ii) along the curved duct, and (iii) at the AIP. Both
PIV and pressure measurements were conducted and are presented
in this section.
3.1. Upstream of the curved duct
Prior to the analysis of the effect of the boundary layer thickness
on the flow field downstream, we must ascertain that the incoming
flow is uniform and symmetric. Thus, the pressure distributions
and the boundary layer velocity profiles were measured upstream
of the curved section, i.e. prior to the inlet plane of the duct (x/
D = 0).
Two techniques were utilized to quantify the incoming flow
field. The first method utilized a total pressure Kiel probe, which
was traversed across the height and span upstream of the inlet
plane at x/D = À1.56. The traversing of the probe was computer
controlled and stepped in 0.5 mm increments close to the wall,
and 2.5 mm intervals towards the center of the duct. The height
and span were normalized by the inlet width (D = 114.3 mm),

and the pressure measured by the Kiel probe was normalized by
the total pressure in the settling chamber of the flow facility.
Fig. 9 shows the total pressure at Mach 0.44, which is nearly uni-
form across the height as well as the span of the duct, except close
to the duct walls, where a boundary layer exists.
The second method utilized was the PIV, which captured the
velocity flow field before the turn (x/D = À0.58) at three spanwise
locations (z/D = 0, ±0.25) and enabled the extraction of the bound-
ary layer profiles. As presented in Fig. 10, the flow field upstream of
the turn is two dimensional.
As was mentioned above, the main goal of the present work is
to change the weighted contribution of the flow separation to
affect the secondary flow structures. Thus, a honeycomb was
inserted upstream of the turn at various heights (into the flow)
to create a pressure drop across it which generated a velocity def-
icit. As the honeycomb was inserted deeper into the flow (i.e.,
increase h/d0) the velocity deficit of the boundary layer increased,
which made it more susceptive to separation. Fig. 11 shows the
velocity profiles at a streamwise location of x/D = À0.58. The
velocity was normalized by the freestream velocity, U1, and the
height was normalized by the boundary layer thickness
(d0 = 0.99U1) of the baseline. The value of d0 is defined at x/
D = À0.58 for M = 0.44 and is kept the same throughout the paper
for all the different Mach numbers. Similar trends were seen for
M = 0.2 and 0.58 but not shown here for brevity.
The velocity profiles were also measured at a streamwise loca-
tion of x/D = À0.14 and are shown in Fig. 12. The same trend seen
upstream is also present at this streamwise location, except that
the flow accelerates due to the proximity to the turn.
In order to quantify the results presented in Fig. 11 and to show
the momentum deficit, which increases due to the insertion of the
honeycomb, the momentum thickness, h, is presented in Table 1
for each velocity profile.
3.2. Along the curved surface
Next, the distributions of the pressure coefficient, Cp, along the
curved surface are presented at the three Mach numbers of 0.2,
0.44 and 0.58 (Figs. 13–15, respectively), at three spanwise loca-
tions of z/D = 0, ±0.25 for four different honeycomb heights. These
plots are divided into three regions:
Region 1, where the flow accelerates around the turn and three
dimensionalities develop for the baseline case. However, when
the honeycomb is introduced into the flow it results in quasi
two-dimensional flow acceleration, where the asymmetry
decreases as h/d0 increases.
Region 2 represents the region where the baseline flow is sep-
arated as indicated by the constant pressure. The presence of
the honeycomb affects the separation region by making it
more two-dimensional and extends it farther downstream, as
is expected since the velocity deficit of the incoming boundary
layer was increased (as discussed above).
Region 3, where without the honeycomb, the flow on left side of
the duct (z/D = À0.25) reattaches first. By increasing the honey-
comb height, the difference in the pressure coefficient between
z/D = À0.25 to the other two spanwise locations (z/D = 0 and
0.25) decreases. This means that the reattachment point is far-
Fig. 16. Pressure recovery at the AIP for the baseline and the three different honeycomb heights at M = 0.2.

ther downstream with the honeycomb than without it. There-
fore, elevating the honeycomb increases the extent of the sepa-
ration region.
3.3. At the AIP
Next, the pressure recovery at the AIP was measured and is
discussed in this section. Figs. 16–18 present the time-averaged
pressure recovery at three Mach numbers and four honeycomb
heights. The red region represents a pressure recovery of 1,
meaning that there are no pressure losses, and the blue region
represents pressure losses. As can be seen, without the honey-
comb, the asymmetry, which is a result of the superposition of
the separation and the secondary flow structures, is clearly pres-
ent for all three Mach numbers. Here, the lowest pressure recov-
ery is on the bottom right side. However, increasing the
honeycomb height decreases the asymmetry, with minimal effect
of the overall pressure recovery.
Table 2 presents the pressure recovery values averaged over the
entire AIP plane, PRave, and over only the lower half of the AIP
plane, PRlowerhalf, for the different honeycomb heights. An important
result from this study is the low impact of the honeycomb mesh on
the pressure recovery while improving the symmetry of the pres-
sure distribution. By looking at the values of Table 2, the largest
impact on the pressure recovery is at M = 0.58 between the base-
line case (without honeycomb) and the case with the honeycomb
inserted at h/d0 = 2.2, where the difference is 0.68%.
In addition to the time-averaged pressure recovery, the effect of
the honeycomb on the fluctuating total pressure was also explored.
Fig. 19 presents the standard deviation of the pressure recovery for
the same cases showed in Fig. 18. For the baseline cases (i.e., with-
out the honeycomb present) at all three Mach numbers, there is a
large concentration of pressure fluctuations in the lower portion of
the AIP, which is the region where the PR is the lowest. When the
honeycomb is inserted into the duct (where h/d0 = 2.2), the fluctu-
ating pressure fields show symmetric distribution of the pressure
fluctuations field.
In addition to a global evaluation of the fluctuating pressure
field, more detailed examination was conducted where the power
spectra at three spanwise locations (sensors 23, 28, 33) at the AIP
were calculated and are presented in Fig. 20. For all three Mach
numbers, without the honeycomb in the flow, the power spectra
at the three spanwise locations are different, whereas the spectral
content on the left side and the middle of the AIP (sensors 23 and
28, the red and green lines, respectively) contains higher energy
than that on the right side (sensor 33, blue line). When the hon-
eycomb is inserted into the flow there is a negligible change on
the energy content in the left side of the AIP, meanwhile the
energy on the right side increases considerably. This modification
in the amplitude of the power spectra further suggests that the
introduction of the honeycomb induce a more symmetrical flow
field (both the time average and the fluctuating). Also, the mag-
nitude of the peak, which represents the dominant shedding fre-
quency of the separated flow, increases with the honeycomb
inserted into the flow, corroborating the results shown previ-
ously. The increase in amplitude, especially at the shedding
energy, is another indication that the separation becomes more
dominant and by doing so helps push the saddle–saddle point
farther downstream in the duct. This saddle–saddle point, as
mentioned previously, causes the onset of the instability respon-
sible for the asymmetry encountered in the flow without the
honeycomb actuator.
Table 2
Pressure recovery at the AIP.
h/d0 M = 0.2 M = 0.4 M = 0.58
PRave PRlower half PRave PRlower half PRave PRlower half
0 0.98904 0.98185 0.94934 0.91631 0.92183 0.87502
0.74 0.98916 0.98217 0.94905 0.91563 0.92115 0.87326
1.47 0.98879 0.98126 0.94690 0.91068 0.91731 0.86505
2.2 0.98857 0.98070 0.94517 0.90656 0.91560 0.86134
Fig. 19. Standard deviation of the pressure recovery at the AIP at M = 0.2, 0.44 and 0.58. Upper row: baseline flow (h/d0 = 0), and lower row: the most symmetrical
configuration achieved in the current experiments (h/d0 = 2.2).

4. Conclusion
The present work showed that, under some conditions, an
asymmetric flow field could develop inside a short inlet duct. In
order to correct this asymmetry, a two-dimensional honeycomb
mesh was inserted upstream of the curved duct to increase the
velocity deficit of the incoming boundary layer. It was shown
experimentally, using steady and unsteady pressure measure-
ments and PIV, that the manipulations of the velocity deficit led
to a more promptly flow separation, which, as a result, modified
the secondary flow structures and corrected the flow asymmetry
at the AIP. The experiments were conducted at Mach numbers of
0.2, 0.44 and 0.58, and the effect of inserting a honeycomb was
tested by increasing its penetration into the flow from 0 to 2.2
times the boundary layer thickness of the baseline flow upstream
of the curve. Using the honeycomb, flow symmetry was achieved
for the specific geometrical configuration tested with a negligible
decrease of the pressure recovery at the AIP. Moreover, the span-
wise uniformity of the flow increased as the height of the honey-
comb was increased.
Acknowledgements
This work was funded by the Northrop Grumman Corporation
(monitored by Ms. Florine Cannelle). The authors would also like
to thank the help of Dr. Israel Salvador and Mr. Brian Debronsky.
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