AIAA 2011-3911
41st AIAA Fluid Dynamics Conference and Exhibit
27 - 30 June 2011, Honolulu, Hawaii
Passive Device Flow Control for Normal Shock/Boundary
Layer Interactions in External Compression Inlets
Trajaen J. Troia1, Amish A. Patel2, Dick Crouse3, and Gordon R. Hall4
Northrop Grumman Aerospace Systems, Redondo Beach, CA, 90278
A passive method of controlling normal shock/boundary layer interactions (SBLIs) for
external compression inlet applications is the use of micro-ramps for boundary layer control
(BLC). These devices were placed upstream of normal SBLIs of various shock strengths and
analyzed with Reynolds Averaged Navier-Stokes (RANS) CFD. A simplified CFD model
produced normal shock phenomena identical to that encountered near the throat of an
external compression inlet. Micro-ramp configurations were compared to baseline no flow
control cases, and both qualitative and quantitative performance metrics were assessed.
Simulations indicated that micro-ramps have the potential to reduce BL separation and
improve inlet performance as well as the potential to improve stability for normal SBLIs.
Converged RANS solutions were demonstrated up to normal shock strength of Mach 1.8. A
preliminary hybrid RANS / Large Eddy Simulation (LES) investigation was also performed,
indicating that further research should move in the direction of time accurate CFD and
experiment to assess the dynamic flow environment and stability characteristics introduced
by the passive devices. Based on the results of this study further investigation of passive
devices is warranted to achieve the goal of zero bleed inlets for supersonic aircraft.
Nomenclature
AOA
BL
BLC
c
CFD
δ
DOE
h
H
L
LES
M
Plcoal
P0
Ps_Ps0
Pt
Pt_Pt0
RANS
SBLI
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
Angle of Attack
Boundary Layer
Boundary Layer Control
Chord Length
Computational Fluid Dynamics
Boundary Layer Height
Design of Experiments
Actuator Height
Throat Height, or Shape Factor
Diffuser Length
Large Eddy Simulation
Mach
Local Static Pressure
Freestream Static Pressure
Static Pressure Rise
Total Pressure
Total Pressure Recovery
Reynolds Averaged Navier-Stokes
Shock Boundary Layer Interaction
1
Engineer, Propulsion System Design & Integration, One Space Park Drive, Redondo Beach, CA 90278, AIAA
Senior Member.
2
Engineer, Propulsion System Design & Integration, One Space Park Drive, Redondo Beach, CA 90278, AIAA
Senior Member.
3
Engineer, Propulsion System Design & Integration, One Space Park Drive, Redondo Beach, CA 90278, AIAA
Senior Member.
4
Subject Matter Expert/Consultant, Propulsion System Design & Integration, One Space Park Drive, Redondo
Beach, CA 90278
1
American Institute of Aeronautics and Astronautics
Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
U
U0
VG
WL
X
y+
z
=
=
=
=
=
=
=
Streamwise Velocity
Freestream Streamwise Velocity at Given X Location
Vortex Generator
Water Line
Streamwise location
Non-dimensional wall distance
Distance from Surface
I. Introduction
Shock Boundary Layer Interaction and Shock Induced Separation
There are two basic forms of shock boundary layer interaction (SBLI): oblique SBLI and normal SBLI. While
the underlying physics of both remains similar, normal SBLI tends to result in increased performance degradation
due to the higher losses inherent in a normal shock, as well as the larger pressure gradients a boundary layer must
negotiate when passing through a normal shock. When a momentum deficit boundary layer encounters the pressure
rise through a shock, the potential for boundary layer separation and accompanying total pressure loss exists. The
Hall Criterion1 has shown that separation will occur when the Pt at any height of the upstream boundary layer is less
than the downstream static pressure after the shock. Essentially any flow that does not have enough momentum to
overcome the pressure rise will stagnate and reverse. Numerous computational and experimental studies2,3 have
shown that SBLI’s above M = 1.3 involving typical turbulent flat plate boundary layers will result in separation
regions of sufficient magnitude to have a negative impact on aircraft performance.
In an external compression inlet, the terminating normal shock is the most aggressive SBLI a boundary layer will
encounter. Most supersonic external compression inlet designs have terminal normal shock strengths between M =
1.4-1.8. While total pressure loss is the main performance detriment of boundary layer separation, increased
distortion occurs due to the separation pockets travelling down the diffuser duct, resulting in concerns for engine fan
aeromechanics and operability. Moreover, separation can set the inlet into buzz and cause engine stall1.
Boundary Layer Control Methods
Legacy methods for normal SBLI control include porous wall bleed and slot bleed. Porous wall bleed is based
upon the principal of removing the momentum deficit in the boundary layer, thus leaving only high momentum flow
to pass through the SBLI and avoiding separation - thus it is a preventative means of eliminating BL separation
which requires flow upstream of the normal shock to be removed. Porous wall bleed relies on a pressure differential
across the ramp surface to drive flow through flush holes along the inlet surface whose diameter is typically about
equal to the boundary layer displacement thickness. Slot bleed removes boundary layer separation after it has
occurred, and requires much larger amounts of flow; and is therefore less efficient. Porous wall bleed is a tried and
true method of boundary layer control and provides shock stability benefits, as well as, a degree of off design
performance, first demonstrated by Wong circa 19744 and with application to many supersonic aircraft designs
beginning with the YF-17 of the early 70’s.
Bleed systems in general require ducting and integration volume for the low pressure plenums. Additionally the
aircraft incurs a drag penalty due to the ram drag of the air that is bled and dumped overboard. Typically the bleed
exhaust can be located in a favorable pressure gradient on the aircraft or used for engine bay cooling.
Passive device flow control encompasses airflow management systems that do not require any energetic input to
manipulate the flow field and provide boundary layer control. The main category of passive device generally
discussed in literature and analyzed for inlet integration are vortex generators (VGs). The traditional VG
configuration is a rectangular or triangular blade attached to the wall, protrudes into the boundary layer, and is
oriented at an angle of attack (AOA) to the streamwise flow, flow to generate vorticity. The vorticity serves to mix
high momentum flow from the upper portions of the BL with low momentum boundary layer flow near the wall;
increasing the likelihood the boundary layer can negotiate an adverse pressure gradient.
VGs are an attractive option for integration for many reasons. First, they do not require supporting hardware and
plenums below the compression surface, thus saving on weight and space allocation. Moreover, due to the fact that
passive devices do not remove or add mass flow to the inlet, they do not incur bleed drag and associated engine
performance penalties. Passive devices also provide the added benefit of being naturally failsafe. Unless the device
is physically removed from the inlet, it will continue to provide vortex generating action, albeit at a reduced
effectiveness at off design conditions. Some VG designs however, such as blade type VGs, can present challenges
for maintainability in practical applications. A crewmember working inside an inlet performing routine maintenance
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activities must take care to not bend or distort any of the VGs, while also ensuring not to harm oneself on the sharp
surfaces of the VG.
One passive device that is more robust and maintainable than the traditional blade VG is the micro-ramp.
Although the blade type VG was shown to be 4X more effective than the micro-ramp in producing cross-stream
vorticity5, it has inherent structural/fatigue risks for dynamic, high temperature environements. The micro-ramp
device has been investigated in experiment and CFD for its potential as a VG for BLC. Initial studies were focused
on micro-ramps that had heights greater than that of the boundary layer, in an attempt to physically force flow down
into the boundary and create vortices that would mix the high momentum free stream with the low momentum
boundary layer. Previous studies have shown6 that micro-ramps larger than the BL exhibited unsatisfactory
performance due to excessive drag and the inability to provide efficient mixing within the BL. Recent experiments7,8
have shown promise for a new class of passive devices termed sub boundary layer VGs to provide not only lower
drag, but the same if not better SBLI performance as traditional scale VGs.
The type of sub boundary layer VGs chosen for this study is the micro-ramp. To date, many micro-ramp studies
have been conducted on oblique SBLI control for mixed-compression type inlets9. Other more recent work has been
done on normal SBLIs. In most cases, experiments have been dominated by corner flows that do not allow for the
effect of the micro-ramps to be assessed at a unit problem level. The CFD model investigated herein was created to
remove any 3-D effects induced by sidewalls, and study the physics of micro-ramp normal SBLI control at the most
basic level.
Mirco-ramp Configuration Screening
Design of Experiments (DOE) optimization methods were used to generate micro-ramp devices for Mach
numbers of interest. This optimization process makes use of some preliminary CFD screening and analysis using
desirability function. Figure 1 shows an example optimal desirability surface for the free stream Mach number (M)
actuator height (h) factor interaction (M*h). This surface allows an actuator height (h) to be chosen in order to best
manage SWBL interactions over a range of design normal shock Mach number (M). This is particularly important
for supersonic inlet design, as the inlet must operate efficiently at off-design conditions. This desirability function
allows for design considerations to be accounted for in selecting a BLC solution that will provide optimal
performance over the vehicle’s flight envelope. Other geometric parameters like micro-ramp Angle of Attack
(AOA) and device spacing were investigated using a similar approach in achieving the optimal micro-ramp
configuration for this study. Information on the DOE optimization process can be found in literature published by
Ventresca7.
Freestream Mach
Micro-ramp Height
Figure 1. Micro-ramp Height Desirability DOE Response Surface (Image Courtesy of NASA/GRC)
II. Normal Shock Model Description
The CFD model utilized for this study was created to analyze BL separation and BLC to expand the external
compression inlet design space. Of particular interest for the present study is the use of passive micro-ramp, subboundary layer flow control devices to potentially prevent normal shock BL separation and enable normal shock
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stability upstream of an external compression inlet throat. The CFD model consists of a pressure far field boundary
condition along the upper domain, sufficient flat plate length to allow for BL growth, a straight cowl and constant
area duct. At the exit of the constant area duct a back pressure condition is set to simulate the effect of an inlet, and
to form a standing normal shock upstream of the simulated inlet. The BL growth section length and height of the
cowl were chosen to simulate a BL height to throat height ratio similar to that of a Mach 2+ class supersonic vehicle.
Figure 2 shows a notional schematic of the CFD model used in the baseline cases.
Figure 2. CFD Model Description
The CFD domain is considered a strip model, with one micro-ramp device and half the device spacing on
either side of the ramp modeled. A symmetry boundary condition is used on the sidewalls to simulate an array of
micro-ramps across the width of the inlet. The upper boundary of the freestream domain is modeled as a far field
boundary condition to allow for MFR effects to be properly simulated. The grids were structured with the baseline
grids without micro-ramps consisting of 2.30M points while the cases with the micro-ramps consist of 2.82M points.
Completely converged RANS solutions were achieved with a Northrop Grumman Aerospace Systems (NGAS) inhouse CFD solver. Menter's SST turbulence model was used. Viscous wall spacing fully resolved the boundary layer
down to y+ of 1 or less. Figure 3 displays the CFD grid domains for the baseline and passive micro-ramp models.
Baseline
Micro-ramp
(M = 1.4)
Figure 3. Baseline and Micro-Ramp Meshes
III. Discussion of Results
Selection of Freestream Mach Number
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An appropriate reference was chosen for evaluation of device effectiveness. The baseline cases consist of the
Normal Shock Model run with no BLC applied, resulting in reference cases that exhibit BL separation. In flow
control research and experimentation it is important to report on flow control device effectiveness in solving a flow
field issue that is imperative to correct, and cannot be resolved by conventional means of geometric shaping. Some
publications demonstrate successful flow control on flow fields that are of no consequence to aircraft performance
and thus hold little merit as a tool to the aircraft designer. The baseline reference flow field for this study was
chosen for consideration of applicability to an M2+ supersonic aircraft system.
Cases of particular interest for investigation into micro-ramp SBLI control were run at freestream Mach numbers
of 1.4, 1.6 and 1.8, with Mach 1.4 representing the lower bound shock strength in which SBLI beings to result in
appreciable levels of BL separation. To illustrate this effect, Figure 4 provides a comparison of the Normal Shock
Model, at freestream Mach numbers from Mach of 1.2-1.8, run at Reynolds numbers from 2.85E5 – 4.27E5 [1/in].
M = 1.20
M = 1.40
M = 1.20
M = 1.40
M = 1.25
M = 1.45
M = 1.25
M = 1.45
M = 1.30
M = 1.60
M = 1.30
M = 1.60
M = 1.35
M = 1.80
M = 1.35
M = 1.80
Figure 4. Minimum Mach for Appreciable BL Separation
Figure 4 clearly shows a negligible effect due to the SBLI between Mach 1.2-1.3 (slight BL growth but no
separation); whereas the Mach 1.4 case exhibits a BL separation that would result in decreased inlet performance.
Since Mach 1.4 marks the lower end separation normal shock strength, it was chosen as the first case. For external
compression inlet design, the most aggressive design normal shock strength is Mach 1.8, due to the increasing total
pressure loss with increasing shock strength. Mach 1.6 represents nominal cruise normal shock strength for a Mach
2+ supersonic vehicle.
Qualitative Micro-ramp Comparisons
In the discussion of results regarding the flow field observed with micro-ramp BLC devices in place, “upwash”
and “downwash” regions may be referenced. These streamwise planes correspond to spanwise locations in which
vortex flows emanating from the micro-ramps have in plane trajectories in either the upward or downward
directions. Figure 5 illustrates these flow regions with a basic spanwise CFD image depicting the counter-rotating
vorticity shed by the micro-ramps.
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Upwash Region
Upwash Region
Downwash Region
Figure 5. Upwash and Downwash Reference Planes (Image Courtesy of NASA/GRC)
Baseline cases with no flow control applied were run to create a proper reference for determining the
effectiveness of the different flow control configurations. An optimized micro-ramp configuration was selected
based on the DOE desirability optimization method for each Mach number. Mach contours from the RANS CFD
solutions are provided in Figure 6 for normal shock Mach numbers of 1.4, 1.6 and 1.8, run at Reynolds numbers
equivalent to Reynolds numbers of 3.32E5, 3.80E5 and 4.27E5 [1/in]. The left side of Figure 6 shows the baseline no
flow control cases, while the images on the right show the micro-ramp configurations at streamwise planes in the
vortex upwash and downwash regions.
No BLC Reference
Mach=1.4
Passive Device BLC
Mach=1.4
Mach=1.4
Mach=1.6
Mach=1.6
Mach=1.8
Mach=1.8
Appreciable
Separation
Mach=1.6
Massive
Separation
Mach=1.8
Massive
Separation
Upwash Region
Downwash Region
Figure 6. RANS CFD Study Comparison Summary
The no flow control cases for Mach 1.6 and 1.8 show massive BL separation upstream of the inlet throat (Mach
1.4 shows a smaller separation due to the weak pressure gradient through the normal shock). Qualitative assessment
of micro-ramp configurations show that massive BL separation has been suppressed for all simulated Mach
numbers; although there are still large regions of depressed momentum flow downstream from the normal shock.
Massive surface separation has been translated to a smaller surface separation bubble and off-surface, free shear
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layer viscous losses. Aspects of the flow field that were determined necessary for further investigation are: 1) the
effect of vortex mixing and redistribution of BL momentum; 2) the pre-compression shock structure generated by
the micro-ramps; and 3) three-dimensional (3-D) relief due to the spacing between the micro-ramps. The first
potential mechanism is the most common for passive device / vortex generator flow control in both subsonic and
supersonic flow fields. The initial hypothesis was that the oblique shock generated by the micro-ramps is a major
player in the prevention of separation due to its softening of the pressure gradient through the SBLI. It is now
believed that this may not be true, but that all 3 mechanisms work in unison to help suppress massive BL separation.
The ensuing discussion will provide rationale for this claim.
Throat Total Pressure Recovery Analysis
The most valuable measurement of total pressure recovery for this model is taken at the simulated inlet throat.
Generated to investigate the SBLI and inlet throat physics, the model possesses a simplistic cowl and diffuser to set
up a normal shock and control MFR. Accomplished with a constant area, straight section, the benign “diffuser” flow
does not encounter adverse pressure gradients due to area divergence or curvature. In an aircraft’s external
compression inlet integration, the throat flow field would enter a 3-D diffuser with area divergence, and surface
curvature. Therefore, the most relevant information from the CFD results presented herein is that from the throat
plane, as this can be thought of as a boundary condition for flow incoming to an aircraft’s unique diffuser. Figure 8
compares the inlet’s throat recovery for the baseline and micro-ramp cases as a function of normal shock Mach
number.
Throat Plane
Figure 7. Throat Recovery Comparison for Baseline and Micro-ramp Solutions
First to notice from Figure 8Figure 7. Throat Recovery Comparison for Baseline and Micro-ramp
SolutionsFigure 7 is the loss in total pressure exhibited by the baseline cases (2-3% lower than ideal normal shock
total pressure loss). For an aircraft inlet with a 3-D diffuser, these levels would translate to a higher loss in total
pressure upon reaching the engine face, due to internal viscous mixing losses associated with the separated layer, as
well as friction through the diffuser’s length, and normal shock separation aggravating the flow through the diffuser.
The first point to be made is that the baseline separation is higher recovery than expected in a real aircraft inlet
integration, due to the lack of a real diffuser. The Mach 1.8 baseline separation would result in a much lower engine
face total pressure recovery relative to the Mach 1.6 (and Mach 1.6 relative to Mach 1.4) case if they were coupled
with a 3-D duct. This is due to viscous losses from the irreversible thermodynamic process that occurs as the
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separation induced free wake mixes out, wall friction, and induced diffuser performance loss as unhealthy BL flow
negotiates 3-D geometry. This means that the 2-3% total pressure recovery loss measured at the throat plane is only
the start of mixing losses. For a real inlet with a relatively straight length to diameter ratio (L/D) = 4, and the
observed separated flow throat profile, would result in a total pressure recovery loss of approximately 5-7%.
The micro-ramp case at Mach 1.6 is shown to be the most effective in increasing the pressure recovery over its
corresponding baseline case. All micro-ramp cases are still short of achieving the ideal pressure rise, but it should be
noted that any improvement in recovery is offset due to the drag on the micro-ramp itself. For most flow fields
requiring BLC, the effectiveness of passive devices diminishes as the adverse pressure gradient increases in severity,
and many times a threshold is met at which redistribution of the BL momentum is not enough to eliminate
separation and other means of BLC are required. The effectiveness in improving recovery is somewhat diminished
compared to the Mach 1.6 case.
Static Pressure Rise Effectiveness of Micro-ramp BLC
The next effectiveness evaluation method is to investigate the static pressure rise through the SBLI region. Figs.
8-10 provide static pressure rise comparisons for the baseline and micro-amp cases. In each of the static pressure rise
plots, there are separate data series for the micro-ramp upwash and downwash regions.
Figure 8. Static Pressure Rise Comparison for Mach 1.4 Case with M = 1.4 Micro-ramp
At Mach 1.4 the micro-ramps are effective at keeping the shock near its ideal position. The separation of the
baseline case is mild in comparison, but a steeper static pressure rise is observed with the micro-ramp, whereas the
baseline no BLC case exhibits a more gradual rise and does. More noticeable are the effective static pressure rises
for Mach 1.6 and 1.8. For Mach 1.6, the micro-ramp allows the static pressure to reach the ideal level over a short
distance near the SBLI region. Comparing the Mach 1.8 case shows that the micro-ramp again improves the static
pressure rise effect (Figure 10), as the baseline no BLC case pressure rise begins very far forward, and only recovers
to the same separation level observed in the Mach 1.6 case. The micro-ramps are able to achieve a relatively steep
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rise in static pressure, but are unable to achieve the full ideal rise. These results indicate that the effectiveness of
micro-ramp SBLI has somewhat diminished at the maximum normal shock strength.
Figure 9. Static Pressure Rise Comparison for Mach 1.6 Case with M = 1.6 Microramp
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Figure 10. Static Pressure Rise Comparison for Mach 1.8 Case with M = 1.8 Microramp
Boundary Layer Comparison
A comparison of boundary layers from the M = 1.6 cases at two constast fuselage stations and buttline
combinates is shown in Figure 11, with the boundary layers from the baseline, no flow control case shown in red,
and the case with the M = 1.6 microramp shown in blue. Upstream of the microramp location the boundary layers
are shown to be exactly equal and agree well with the 1/7 power law profile shape. The downstream boundary
layers, BL2 and BL4, for the no flow control case clearly depict separation and have a shape factor of H ~ 6, well
above the H = 2.2-2.4 shape factor that is traditionally accepted as indicative, while not required, for separation.
The boundary layer profiles at the same locations for the the M = 1.6 microramp case show that the micro-ramp is
able to reduce the shape factor to H<1.8, and eliminates the reverse flow at the boundary layer locations shown,
bringing the profile back closer to a 1/7 power law profile shape. This result shows that the micro-ramps were
successful in improving the health of the BL downstream of the shock and near the inlet throat. Entering the 3-D
contour of a real aircraft diffuser, this flow field should perform substantially better than the flow exhibited in the
baseline cases with massive separation.
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BL1
1.00
0.90
0.80
0.70
0.60
0.50
0.40
0.30
0.20
0.10
0.00
FC
BL2
1/7 Power Law Profile
No FC
H = 1.32
H = 1.32
z/δ
z/δ
No FC
-0.40 -0.20 0.00 0.20 0.40 0.60 0.80 1.00
1.00
0.90
0.80
0.70
0.60
0.50
0.40
0.30
0.20
0.10
0.00
FC
H = 6.10
H = 1.79
-0.40 -0.20 0.00 0.20 0.40 0.60 0.80 1.00
U/U0
U/U0
BL3
FC
BL4
1/7 Power Law Profile
No FC
H = 1.32
H = 1.32
z/δ
z/δ
No FC
1.00
0.90
0.80
0.70
0.60
0.50
0.40
0.30
0.20
0.10
0.00
1/7 Power Law Profile
1.00
0.90
0.80
0.70
0.60
0.50
0.40
0.30
0.20
0.10
0.00
FC
1/7 Power Law Profile
H = 6.10
H = 1.55
Approximate
Normal Shock
Location
BL2
-0.40 -0.20 0.00 0.20 0.40 0.60 0.80 1.00
-0.40 -0.20 0.00 0.20 0.40 0.60 0.80 1.00
U/U0
U/U0
BL4
BL1
BL3
Figure 11. Boundary Layer Profile Comparison for M = 1.6 Baseline (Blue) and M = 1.6 with M = 1.6
Microramp (Red)
Micro-ramp BLC Separation Reduction Mechanisms
Previous studies of ramp type vortex generators in supersonic flows have suggested the vortices shed from the
VG are the primary mechanism by which the micro-ramp enables BLC. While it is true that the vortices enhance
momentum transfer between the deficit boundary layer and momentum rich freestream, this paper proposes that in
the case of micro-ramp normal SBLI control, the vortex induced momentum mixing is one of three mechanisms by
which BLC is achieved.
Figs. 12-15 show four groups of streamlines seeded in the flow upstream of the VG and propagated downstream
through the normal SBLI for the Mach 1.6 case. Figure 12 and Figure 13 are streamlines that were initiated at a
height about equal to the incoming boundary layer height and half that height. Both sets of streamlines show how
the high momentum flow of the upper regions of the boundary layer pass over the VG and begin to integrate into the
vortices after about one micro-ramp length downstream. This agrees with what most technical literature claims is
the primary mechanism of boundary layer control, momentum mixing.
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Figure 12. Stream trace Group of Max Height BL Flow for Mach 1.6 Case
Figure 13. Stream trace Group of Mid Height BL Flow for Mach 1.6 Case
Figure 14 is a set of streamline traces taken very near to the wall and propagated downstream. These images reveal
that a good portion of the lowest momentum flow from the bottom of the boundary layer is ejected up and away
from the lower surface. This depressed flow passes through the normal shock away from the wall, a crucial aspect
contributing to the elimination of massive BL separation. This low momentum flow mixes with higher momentum
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flow and is better able to negotiate the pressure rise through the normal shock. The remaining boundary layer flow
that is not ejected off the surface stagnates at the normal shock and separates, but the recirculation zone is confined
to a small bubble as opposed to a wedge, as seen in Figure 15, creating a localized recirculation zone, as opposed to
a massive, recirculating separation wedge along the surface.
Figure 14. Stream trace Group of Near Surface BL Flow for Mach 1.6 Case
Figure 15. Isosurface of U = 0 for M = 1.6 No Flow Control (Left) and Micro Ramp (Right) Cases
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Figure 16. Stream trace of Flow through Different Spanwise Extents of Microramp
Figure 16 shows three streamline rakes seeded at different spanwise extents in the boundary layer. The orange
and red streamline groups encompass the outermost 10% of the span while the blue streamline group is the 80%
remaining in the middle. The starboard grouping, in red, feeds into the separation and recirculation zone, and is
subsequently pulled into the vortex region, which originates from the blue streamlines which are ejected off the
surface of the microramp.
The first key mechanism is the oblique shock generated from the micro-ramp, which accomplishes two flow
field benefits. First, the oblique shock acts as natural BL diversion of flow away from the surface; low momentum
fluid that cannot negotiate the oblique shock pressure rise will travel up the shock. Combined with the ramp AOA,
this effect induces an initial trajectory of the low momentum fluid moving away from the wall. Secondly, the shock
increases the static pressure on the micro-ramp surface and enhances the pressure gradient from the top of the ramp
to the sidewalls, which in turn increases the strength of the counter-rotating vortex generation. This is further
amplified by the expansion from the ramp sidewalls that drops the sidewall static pressure (Figure 17).
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WL = 0.003
z/δ
WL = 1.667
z/δ
WL = 3.333
z/δ
WL = 0.003
z/δ
WL = 1.667
z/δ
WL = 3.333
z/δ
WL = 0.003
z/δ
WL = 1.667
z/δ
WL = 3.333
z/δ
Figure 17. Mach, Total Pressure Recovery and Static Pressure Recovery Contours at Various Waterlines for
M = 1.6 Case
Three-dimensional relief and the micro-ramp sidewall expansion combine with the vortex generation to draw
high energy fluid into the location near the wall. This effect allows the trajectory of the BL to persist in its trajectory
away from the wall. Counter-rotating vortices in the sense created by this micro-ramp orientation has been shown to
combine and create its own “lift off” trajectory9. Under normal circumstances in which pure vortex mixing is wanted
near the wall, this “lift off” flow feature is undesirable; however, for this normal SBLI environment, it creates a
favorable outcome by further enhancing the effect of removing the lowest energy BL flow from the surface.
Moreover, by pushing the boundary layer off the surface, the micro-ramp creates a region of low pressure which in
turn pulls boundary layer flow from the space in between the micro-ramp and its neighbor in towards the centerline
of the micro-ramp, as shown in Figure 18, where the vortices are most effective at pulling boundary layer up and
away from the surface. These effects combine to result in a more healthy (although more three-dimensional)
boundary layer at the shock interaction along the surface, thus reducing the separation. A successful ejection of the
BL to eliminate massive BL separation has been achieved, at the expense of a smaller recalculating bubble and offsurface free shear layer viscous losses. What is unknown at this time is the impact of the phenomena observed on
normal shock stability. Figure 19 clearly shows that the flow after the shock re-expands to close to M = 1.7, yet does
not result in separation in the downwash region, providing evidence of the fact that the boundary layer in that region
is relatively nonexistent. Also shown in Fig. 18 is the smaller recirculation zone occupying the upwash region.
Figure 18. Oil Flow Visualization of Lower Surface for M = 1.6 Case
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Figure 19. Close Up of Shock and Vortex Interaction M = 1.6 Case
Multiple mechanisms were shown to contribute to the reduction of massive BL separation. In order to determine
the full effect of the traditional vortex generation vs. the oblique shock and 3-D relief contributions to the BLC
physics, a body force model for vortices could be used to run an equivalent case where the micro-ramp is removed
and a pair or vortices of equivalent strength are produced. These vortices would be inserted at the same steamwise
location that the micro-ramp sheds its counter-rotating vortex structure. Since the physical geometry will not exist,
the oblique shock will not be produced and the effect of the vortex mixing will be fully captured, independent of any
oblique shock and geometry dependent 3-D effects, and can be quantitatively subtracted from results with the microramp in place.
The presented results of the RANS CFD analysis provided insight into the flow physics involved with passive
device normal SBLI control, and took a first look at evaluating its effectiveness in reducing normal shock induced
separation. By no means are these results considered 100% accurate or final in any way – more work is required to
verify this CFD model and validate that the passive devices can indeed be used in real aircraft inlet integration. One
major area of concern is normal shock stability when implementing micro-ramp BLC. An experiment will be
required to fully investigate this consideration in the future, but time accurate CFD can be used as both a screening
and optimization tool to reveal dynamic features of the flow field. Future work will utilize a hybrid RANS/LES time
accurate analysis to provide insight into the shock stability.
Sensitivity Studies
A few basic analysis excursions were performed to evaluate sensitivities of micro-ramp SBLI control. Fig. 19
provides Mach profiles of results for a Reynolds number comparison. Results presented in subsequent sections were
run at a Reynolds numbers of 3.32E5, 3.80E5 and 4.27E5 [1/in]; these results are compared here to a Reynolds 1/3
lower. The micro-ramp sized for Mach 1.6 based on the DOE optimization method was implemented at Mach 1.4,
1.6 and 1.8 for this investigation.
Two important results were achieved from this comparison. The first is that the Mach 1.8 case exhibits massive
separation for the lower Reynolds number run, whereas it demonstrated SBLI control at the higher Reynolds
number. The BL is larger for the higher Reynolds number case, and at this point the Mach 1.6 micro-ramp, operating
at an off-design Mach number (1.8), is no longer of sufficient shape to control the SBLI. This indicates that an
aircraft designer must consider a vehicles flight envelope when performing passive device optimization studies. A
positive result is that the Mach 1.6 micro-ramp works over the full Mach range at the higher Reynolds number,
showing that the devices can potentially be optimized to persist over a Mach range without having to change shape
mechanically.
Fig. 20 presents results obtained for various micro-ramp streamwise positions at Mach 1.4 and Reynolds number
of 3.32E5 [1/in]. Inlet throat recovery was plotted for the three cases. The first case is the original reference microramp position, and the other cases are for one and two chord lengths upstream of the reference. Results indicate that
at two chord lengths upstream from the reference position BLC effectiveness has begun to diminish, although the
change in total pressure recovery is small. Terminal shocks in front of inlets will move in streamwise position as a
function of atmospheric conditions and engine airflow, and as such the success of BLC devices must persist over the
full range of the normal shock’s travel.
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1/3*ReReference
M = 1.80
M = 1.60
M = 1.40
ReReference
Figure 20. Effect of Reynolds Number on Performance of M = 1.6 MR at M = 1.4, 1.6 & 1.8
1.000
0.990
0.980
∆ = 1c
Pressure Recovery
0.970
0.960
0.950
0.940
0.930
0.920
∆ = 2c
0.910
0.900
0.000
0.500
1.000
1.500
2.000
x/c, Microramp Shift from Reference Location
Positive is Upstream
Figure 21. Effect of Microramp Position on Performance of M = 1.4MR at M = 1.4
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2.500
These sensitivities were performed to provide initial insight into off-design effects relevant for aircraft design,
but more work is required to fully understand the impact of Reynolds number, micro-ramp position, and fixed
geometry devices operating over a broad range of flight conditions.
Unsteady
RANS\LES
Steady RANS
Inlet Instability/Buzz Considerations
Figure 22 and Figure 23 provide a comparison between RANS and hybrid RANS/LES solutions for both no flow
control, and micro-ramp configurations at Mach 1.6. The two solution sets show the same overall features: massive
separation with no flow control, and attached flow for the micro-ramp cases. Apparent qualitative differences in the
flow field for the RANS/LES method are the eddy structures and variations in shock structure. Figure 22 and Figure
23 only depict one snapshot in time for the RANS/LES solution.
Baseline Flow (Separated)
Micro-ramp Flow Control
Figure 22. RANS and Hybrid RANS/LES CFD Streamwise Comparisons for Mach 1.6 (Image Courtesy of
NASA/GRC)
RANS Analysis
Steady State
RANS/LES Analysis
Instantaneous
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Figure 23. RANS and Hybrid RANS/LES CFD BL Spanwise Comparisons for Mach 1.6 (Image Courtesy of
NASA/GRC)
For external compression supersonic inlets, instability leading to buzz is typically the result of externallygenerated disturbances within the incoming flow field. The categories of disturbances which can lead to inlet buzz
include shock induced separation of boundary layer on the compression surfaces, slipstream ingestion of an oblique
shock/normal shock “triple point” near the inlet cowl, vortex ingestion from external aerodynamic surfaces, and
ingestion of other viscous wakes/ flow fields resulting from highly integrated aircraft configurations operating at
AOA/AOS and/or at low MFR.
Reference 1 addresses all of the above externally –generated disturbances leading to inlet buzz and provides a
common numerical criterion shown to be in good agreement with test data for each of the buzz inducing flow field
phenomena noted above. While each of these types of disturbances are necessary conditions to induce inlet buzz,
they are not necessarily sufficient conditions, but rather inlet stability is determined by the numerical criteria
presented in the paper by Hall1.
For the present study of the use of micro-ramps using the existing CFD model, the instability problem is
bounded such that only large boundary layer separation going thru the normal shock (or normal shock plus subsonic
pressure rise downstream of the shock that results from reduced MFR) are considered to be a potential source of
inlet buzz. It is noted that compression ramp flow separation ingested by an inlet can be a stable situation. For the
stable case, a RANS CFD solution would be expected to be a valid representation of the flow field/inlet
performance.
However, at a critical level of throat blockage by the separated layer, the flow will go unstable resulting in inlet
buzz. For this case, a RANS CFD solution which results in a stable solution would not at all be representative of
buzz dynamics that may occur in a real situation. Empirical data exists, which when combined with the Hall
Criterion, can serve as a guideline relative to avoiding instability leading to buzz; however, the onset of buzz,
amplitude, and frequency for a specific inlet configuration can only be determined with relative certainty by either a
time accurate CFD investigation or wind tunnel test with high frequency instrumentation.
As follow-on to the current study, it is planned to fully evaluate the micro-ramps in terms of stability/buzz using
hybrid RANS/LES time accurate CFD analysis.
This study has shown that potential for performance enhancement exists with the implementation of passive
devices in external compression inlets as a means to eliminate traditional porous wall bleed for normal shock
stability and BL separation mitigation. Future fuel burn and efficiency requirements provide drive for further
investigation of this technology. Using passive devices in place of wall bleed results in drag reduction and weight
reduction. These factors translate into improved engine efficiency and vehicle range and endurance capability. An
additional benefit of using passive devices, such as micro-ramps, is lower system complexity, higher reliability, and
ease of integration/maintenance.
IV. Conclusion
RANS CFD solutions were presented for micro-ramp SBLI control for normal shock Mach numbers of 1.4, 1.6
and 1.8 at a Reynolds numbers from 3.32E5 to 4.27E5 [1/in]. For these cases, the micro-ramps were successful in
reducing massive BL separation, but at the expense of off-surface viscous free shear layer loss. It was shown that the
throat total pressure recovery was improved over the baseline no flow control cases, and the static pressure rise
through the SBLI region was closer to the ideal normal shock pressure rise.
This model provided a simple way to begin understanding of an extremely complex flow field, and was required
to identify future work in achieving high Mach, bleedless external compression inlets with passive device BLC. As
investigation into passive devices capability to control normal SBLIs continues and these configurations are
optimized, higher-level vehicle performance benefits can be assessed. Without knowledge of how the throat flow
field migrates through a 3-D aircraft diffuser, or how sensitive the shock stability is to MFR variations and offdesign conditions, a performance evaluation using results from the present model would be misleading.
The micro-ramps improved BL shape factor downstream of the SBLI, which indicates a healthy BL would be
provided to an aircraft diffuser duct, unlike baseline cases with massive separation. However it is not improved over
what is typically expected of a traditional bleed system (but bleed systems incur other penalties). A healthy BL
entering the inlet’s diffuser is crucial in modern inlet design – with today’s highly integrated aircraft configurations;
the luxury of long, straight diffuser designs are less prevalent.
A detailed investigation into the unit flow physics was provided, which revelaed that multiple phenomena work
together to eliminate the BL separation. The typical useage of passive devices in redistributing the BL momentum
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through streamwise vorticity generation was shown to be a prominent effect. The pre-compression oblque shock and
shear trajectory of the ramp itself help provide an initial upward component of the BL’s motion as it passes through
the location inhabited by the micro-ramp. Three-dimensional relief between the devices allows for high energy flow
to fill to the location in which the BL would have resided without the device present. Including the effect of spacing,
as well as local expansions due to the micro-ramp shape, these 3-D relief mechanisms provide the means to keep the
BL away from the lower surface. Finally, counter-rotating vortices having the effect of vortex liftoff allow for
persistence of the BL’s ejection as it progresses downstream toward the normal shock.
Overall the result of these combined effects is that the low energy BL is ejected from the lower surface, and then
interacts with the normal shock in a localized wake shock interaction. Future work will include investigation into
inlet buzz/shock stability by first utilizing hybrid RANS/LES time accurate CFD solutions. Preliminary results using
this technique were presented. Using this method, shock stability and passive device optimization can be
accomplished in parallel to prepare for experiments. Unit problem experiments, similar to the unit problem CFD
model from this study, must first be executed to verify the physics observed in the RANS and hybrid RANS/LES
evaluations.
Finally, corner flows have been known to cause problems with micro-ramp SBLI control – to date there have
been no solutions, and more work is required in this area. The final step in the process is a wind tunnel test of an
external compression inlet that has the capability for testing both passive device BLC and traditional bleed. Testing
of this nature would include a full shock stability assessment and total pressure recovery measurements over a broad
range of operating conditions to reveal any potential limitations of passive device normal shock BLC.
Acknowledgments
The authors would like to thank NASA Glenn Research Center Inlet and Nozzle Branch for providing support
necessary to make this work possible. The authors give special thanks to Bernard Anderson of NASA Glenn
Research Center for his technical support and contributions in terms of DOE optimization methods for micro-ramp
shape selection, preliminary RANS CFD screening analysis, and an initial look into time accurate RANS/LES
solutions. We look forward to continued research in this area and truly believe that bleedless supersonic inlet
designs are achieveable and will be beneficial to future, highly efficient supersonic aircraft.
References
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