AIAA 2009-5415
45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit
2 - 5 August 2009, Denver, Colorado
AIAA 2009-5415
Scramjet Mixing Control Using Fin-Guided Fuel Injection
C. Aguilera*, B. Pang†, A. Ghosh†, A. Winkelmann‡, and K.H. Yu§
Department of Aerospace Engineering, University of Maryland, College Park, MD 20742
A specially designed fin was used to assist fuel jet penetration into a scramjet combustor
flow-field. The fin served two specific purposes – first, to enhance mixing between fuel and
air, and second, to minimize the stagnation pressure loss associated with the fuel injection
process. Non-reacting mixing experiments were conducted in a Mach 2.1 blowdown air
facility with transverse injection of choked helium flow simulating the fuel injection process.
The penetration of helium jet and the mixing with supersonic airflow were characterized
using spark schlieren and laser-sheet visualization techniques. Strength of injection-caused
shocks was determined using a series of wall pressure measurements. The present results
were consistent with the findings from our earlier study, which showed substantial increase
in the fuel-jet-penetration height while at the same time reducing the stagnation pressure
loss. In comparison with the baseline injector, the fin-guided injector caused at least a twofold increase in the fuel-jet-penetration height and up to 45% reduction in the shock-caused
static pressure rise. The present results suggest that a properly designed fin could be used to
increase both combustion efficiency from mixing enhancement and propulsion efficiency
from reduced pressure loss.
IXING enhancement is a problem in scramjet
mainly due to extremely short flow residence
time, which is on the order of msec. Also, supersonic
mixing layers spread at a much slower rate in
comparison to incompressible mixing layers subjected to
a similar level of shear stress. This reduced rate of
growth, which is widely known as compressibility
effect1-3, makes mixing in scramjet an even more
challenging problem.
For typical Mach numbers
encountered in scramjet internal flow field, one may
expect up to a five-fold decrease in shear layer spreading
rate due to the compressibility effect.4.
There have been many previous studies5-7 for mixing
enhancement in supersonic shear flows, especially for
propulsion and power applications. Although these
studies have shown that a variety of techniques can be
used to enhance mixing over a wide range of flow length
scales that may be particularly useful for reacting flows,
the associated cost of mixing on the system performance
has not been rigorously considered. Since mixing
enhancement must result in an increase of flow entropy,
there will be some reduction in the stagnation pressure.
Then the amount of decrease in stagnation pressure can
100
Stagnation Pressure Loss (%)
M
I. Introduction
10
1
M=2 (weak shock)
M=3 (weak shock)
M=2 (strong shock)
M=3 (strong shock)
0.1
0.01
0
5
10
15
Deflection Angle (deg)
Figure 1: Stagnation pressure losses
associated with planar deflection
*
Graduate Research Assistant, Aerospace Engineering Dept., University of Maryland, AIAA Student Member.
Postdoctoral Scientist, Aerospace Engineering Dept., University of Maryland, AIAA Member.
‡
Associate Professor, Aerospace Engineering Dept., University of Maryland.
§
Associate Professor, Aerospace Engineering Dept., University of Maryland, AIAA Associate Fellow.
1
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†
Copyright © 2009 by Aguilera et al. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.
20
Figure 2: Experimental setup for fuel injection experiment.
be used to quantitatively assess the cost of mixing. In this study, we consider a tradeoff between mixing
enhancement and stagnation pressure loss by making measurements on these quantities.
The mixing enhancement technique we have been developing utilizes a sharp fin not only for guiding fuel jet
penetration into the air flow but also for triggering a weak oblique shock just upstream of the fuel injection
location.8-10 The fin is designed to minimize the pressure losses associated with jet-induced shocks. For instance,
when a transverse fuel jet is introduced into a supersonic cross flow, a blunt shock is created upstream which
reduces the stagnation pressure downstream. This loss in stagnation pressure is mostly a wasted cost as the shock
helps very little in terms of enhancing turbulent mixing at scales important for combustion.
Typical transverse fuel injection causes deflection in supersonic airflow by setting up an oblique shock.
Potential amount of stagnation pressure loss due to oblique shock is calculated for a two-dimensional flow and the
results are shown in Fig. 1 as a function of shock strength and flow deflection angle. As shown in Fig. 1, the amount
of stagnation pressure loss due to oblique shock can vary widely between a small fraction of a percent and tens of
percent. Our ultimate objective is to optimize the tradeoff between combustion enhancement and stagnation
pressure loss. For now, we quantify the amounts of mixing and pressure loss from our controlled experiments.
II. Experimental Set-up
Fuel injection and mixing experiments were conducted using a blow-down supersonic tunnel driven with a
vacuum tank from the downstream end. This tunnel uses suction from the vacuum tank to establish Mach 2.1
airflow in the test section for approximately 15 seconds. The flow path dimensions and the fin shape are shown in
Fig. 2. Helium, which was used as surrogate fuel, was injected transversely into the supersonic airflow through a
12.7-mm diameter choked orifice, flush mounted on the lower wall. Wall pressure was measured using a series of
pressure ports that lined the combustor upper wall every 12.7 mm.
The fin employed in this study was a triangular blade with a rectangular base of 12.7 mm wide and 127 mm long
mm placed just upstream of the orifice injector. The height of the triangular cross-section varied from 0 at the
leading edge to 38.1 mm at the trailing edge. The fin was placed such that the injector orifice was just downstream
of the fin trailing edge allowing fuel injection into the fin’s wake. The resulting displacement in the tunnel crosssectional area was equivalent to a constant-angle ramp, corresponding to a simple 8.5° ramp for the same footprint
area as the fin base, or a 0.7° ramp if the entire span of the wind tunnel was considered.
A Scanivalve DSA-3217 Digital Sensor Array was used to make static pressure measurements along the top wall
of the combustor. This Array consists of 16 temperature compensated piezoresistive pressure sensors with a
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pneumatic calibration valve. The 16 sensors, or channels, all have a range of 0-1.5 MPa. Their associated error is
±0.2% of scale for pressures less than 0.1 MPa, ±0.12% of scale for pressures between 0.1 and 0.14 MPa, and
±0.05% of scale for pressures above 0.14 MPa. The measured pressures are sent via a TCP/IP connection to a
desktop computer and into a LabView virtual control panel. This virtual interface allows for monitoring of all 16
channels and the DSA’s settings as well as writing of the data to a text file to be read by post-processing software.
For flow visualization, a conventional schlieren technique was used to obtain side profiles while laser sheet
visualization was used to obtain cross-sectional view at selected downstream locations.
Fuel injection experiments were conducted with and without the guide fin in place. The configuration without
the fin was used as the baseline for performance comparison purpose. Table 1 shows the summary of various fuel
injection conditions that were tested.
Table 1: Flow conditions for fuel injection experiment.
m! air
m! fuel
(kg/s)
(g/s)
FuelAir
Ratio
1
1.56
2.7
2
1.56
3
4
case
Stagnation
Pressure (psi)
p0,air
p0,fuel
0.0017
14.7
75
3.5
0.0022
14.7
95
1.56
4.2
0.0027
14.7
115
1.56
5.0
0.0032
14.7
135
III. Results and Discussion
For each of the injection conditions shown in Table 1, the flow mixing behavior is compared for the two injector
configurations, one with the fin (Fin-Guided) and one without (Baseline). The results are presented in Figs. 3-6,
each containing the schlieren images comparison as well as the corresponding wall pressure distributions. The
surrogated fuel, helium, was injected at x/D=0. With increasing injection pressure or fuel flow rate, the fuel jet
penetrated deeper into the free-stream supersonic airflow. The images show clearly that fuel penetration depth was
substantially improved with the use of the fin. As the fuel penetration into the free stream was guided by the fin
wake, the fuel was pushed to the top of the fin relatively unobstructed. The fuel penetrating beyond the fin height
appeared to follow a similar trajectory as in the baseline case.
Static pressure distributions along the upper wall for the corresponding cases were also compared. In these plots,
the x-axis is normalized with the injector orifice diameter, while the y-axis is normalized with the initial stagnation
pressure, p0,air. Also shown here are the corresponding wall pressure distributions prior to the fuel injection, that are
shown as dotted lines. There was a slight increase in the starting pressure for the configuration with the fin as the fin
acting as a shallow ramp induced a weak oblique shock. With fuel injection, the wall pressure was observed to
increase from the pre-injection values due to jet-induced shocks. The amount of pressure increase was thus related
to the strength of shocks caused by fuel injection. Comparison of these pressure plots suggests that the injection
shock strength was substantially lower in the fin-guided injection case than that in the baseline case. Also, as the
fuel flow rate was increased from 2.7 g/s in Fig. 3 to 5.0 g/s in Fig. 6, the shock strength in each case increased as
expected.
For another type of comparison, the amount of static pressure rise associated with fuel injection was normalized
using the pressure distribution before the injection. The injection shocks caused static pressure rise depending on the
shock strength. The results are shown in Fig. 7, which suggest that the fin effectively reduced the strength of the
shock that the flow encountered. The pressure rise associated with fin-guided fuel injection was significantly
smaller. In Fig. 8, the amount of reduction in wall pressure rise associated with shocks is plotted for various
injection pressures. This reduction reached nearly 33% to 47% depending on the injection pressure.
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0.14
Baseline
w/o injec
0.14
Fin-Assisted
w/o injec
0.12
p/p
0
0
p/p
Fin-Assisted
w/o injec
0.12
0.10
0.08
0.10
0.08
0.06
0.06
0
5
10
15
x/D
Figure 3: Schlieren images of fuel injection at 75 psi and
the corresponding pressure distribution on the upper wall
0
0.14
Baseline
w/o injec
0.12
5
10
15
x/D
Figure 5: Schlieren images of fuel injection at 115 psi and
the corresponding pressure distribution on the upper wall
0.14
Fin-Assisted
w/o injec
Baseline
Fin-Assisted
w/o injec
w/o injec
0.12
p/p
0
0
p/p
Baseline
w/o injec
0.10
0.10
0.08
0.08
0.06
0.06
0
5
10
15
x/D
Figure 4: Schlieren images of fuel injection at 95 psi and
the corresponding pressure distribution on the upper wall
0
5
10
15
x/D
Figure 6: Schlieren images of fuel injection at 135 psi and
the corresponding pressure distribution on the upper wall
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Baseline, 75psi
Baseline, 95psi
Baseline, 115psi
Baseline, 135psi
1.5
Normalized Pressure
50
Reduction in Pressure Rise (%)
Fin, 75psi
Fin, 95psi
Fin, 115psi
Fin, 135psi
1.4
1.3
1.2
1.1
1.0
45
40
35
30
60
0.9
0
5
x/D
10
Figure 7. Static pressure rise associated with fuel
injection. The pressure results were normalized
with static pressure before the injection.
15
80
100
120
Injection Pressure (psi)
140
Figure 8. Amount of reduction in shock-induced static
pressure rise associated with fin-guided fuel injection
Cross-sectional views of fuel-air mixing were obtained at several axial locations using planar Mie-scattering
technique. These images are presented in Figs. 9-11. Figures 9 and 10 represent instantaneous images of the flow
fields at various flow conditions and distances downstream from the injector, corresponding to the baseline case and
the fin-guided case respectively. The air flow was illuminated by fog condensation from moist air, while the
surrogated fuel, helium, remained dry and thus appeared as dark spot. As the density of the fog particles in the
airflow were affected by pressure and velocity, the post-shock region appeared brighter than the pre-shock region
and so did the boundary layer where the velocity was slower. The size of the viewing area being presented is 4
injector diameters wide over a full height of the combustor.
It can be inferred from Fig. 9 that the fuel jet penetration in the baseline case was relatively shallow and the fuel
jet dispersed rather slowly. Comparing the corresponding images in Fig. 10, it can be vividly shown that the fuel
penetration depth substantially increased with the fin-guided fuel injection. Averaged images are presented in Fig.
11. Between 80 and 100 instantaneous images were averaged for each case, spanning over 11-14 seconds of run
time. Also, remarkably, much of the fuel in the fin-guided injection case was delivered to the core of the supersonic
airflow rather than staying in the boundary layer as in the baseline case. This difference in mixing characteristics
could have an additional implication if combustion efficiency and wall heat transfer characteristics were considered
in the reacting flow.
IV. Summary and Conclusions
An experimental study on fin-guided fuel injection was conducted to characterize the enhanced mixing and the
reduced pressure loss. Non-reacting fuel injection experiments were conducted in a simulated scramjet flow path.
The results clearly showed that the fin-guided injection approach could not only enhance the mixing between the
supersonic air stream and the injected fuel flow, but it could also reduce the stagnation pressure loss by lowering the
strength of the jet-induced shocks.
From the experimental results, it is clear that the location of fuel delivery into the supersonic airflow can be
controlled by the fin height. In the current configuration, the width of the fin base was matched to the size of the
injector orifice, which resulted in most of the fuel being delivered to the top of the fin. This resulted in more than
doubling the fuel penetration depth in comparison with the baseline transverse injection case. It should also be
possible to control the vertical distribution of fuel by adjusting the fin base width.
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The pressure distribution results showed that the amount of stagnation pressure loss, as inferred by the shockinduced pressure rise, was effectively controlled by the fin initiating the weak oblique shock rather than the fuel jet.
In the current configuration, the maximum saving was achieved with the lowest fuel flowrate case which yielded
more than 45% reduction in shock-caused static pressure rise. Extrapolating from such results, it would be possible
to further reduce the shock-induced pressure loss by optimizing the fin design in the future. The ultimate
optimization test will require combustion experiments with actual thrust measurements.
x/D = 0
x/D = 2
x/D = 4
x/D = 6
x/D = 8
x/D = 10
(a)
p0,fuel
/p0,air
=5.1
(b)
p0,fuel
/p0,air
=6.4
(c)
p0,fuel
/p0,air
=7.8
(d)
p0,fuel
/p0,air
=9.2
Figure 9. Instantaneous images of baseline case fuel injection visualized
with a laser sheet at various axial distances downstream.
(a) fuel flow rate = 2.7 g/sec
(b) fuel flow rate = 3.5 g/sec
(c) fuel flow rate = 4.2 g/sec
(d) fuel flow rate = 5.0 g/sec
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x/D = 12
x/D = 0
x/D = 2
x/D = 4
x/D = 6
x/D = 8
x/D = 10
(a)
p0,fuel
/p0,air
=5.1
(b)
p0,fuel
/p0,air
=6.4
(c)
p0,fuel
/p0,air
=7.8
(d)
p0,fuel
/p0,air
=9.2
Figure 10. Instantaneous images of fin-guided fuel injection visualized
with a laser sheet at various axial distances downstream.
(a) fuel flow rate = 2.7 g/sec
(b) fuel flow rate = 3.5 g/sec
(c) fuel flow rate = 4.2 g/sec
(d) fuel flow rate = 5.0 g/sec
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x/D = 12
x/D = 0
x/D = 2
x/D = 4
x/D = 6
x/D = 8
x/D = 10
x/D = 12
(a)
p0,fuel
/p0,air
=5.1
(b)
p0,fuel
/p0,air
=5.1
(c)
p0,fuel
/p0,air
=9.2
(d)
p0,fuel
/p0,air
=9.2
Figure 11. Time-averaged images comparing the fin-guided fuel injection and the baseline fuel injection
under various conditions.
(a) baseline case with fuel flow rate = 2.7 g/sec
(b) fin-guided case with fuel flow rate = 3.5 g/sec
(c) baseline case with fuel flow rate = 4.2 g/sec
(d) fin-guided case with fuel flow rate = 5.0 g/sec
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Acknowledgment
This study was sponsored by the Office of Naval Research with Dr. Gabriel D. Roy as the scientific officer.
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2
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