AIAA-98-0964
A COMPARISON OF TWO HYPERMIXING FUEL INJECTORS IN
A SUPERSONIC COMBUSTOR
Matthew J. Gaston*, Neil R. Mudford* and Frank Houwing†
* Dept. of Aerospace & Mechanical Engineering,
University College, University of New South Wales
Canberra, ACT, Australia
† Dept. of Physics, The Faculties,
Australian National University
Canberra, ACT 0200 Australia
Abstract
An experimental study has been undertaken to
evaluate the performance of two hypermixing
injectors designed for supersonic combustion ramjet
(scramjet) applications. Supersonic mixing and
combustion studied in a free-piston driven shock
tunnel is examined using surface pressure
measurements and shadowgraphy. Tests were
conducted at two inlet Mach numbers: M=2.5 and
M=3.7.
Introduction
With the present interest in aerospace planes,
considerable effort is being devoted to the
development of propulsion systems that would power
these vehicles. One such propulsion system is the
supersonic combustion ramjet (scramjet) which
involves combustion at supersonic speeds; that is
both primary flow (such as air) and the injected fuel
(such as hydrogen) have to mix and burn at supersonic
velocities. However, it is known that mixing at high
Mach numbers is not very efficient and significant
efforts are therefore being made to study supersonic
turbulent mixing and to seek ways to enhance the
mixing.
Several mixing enhancement devices have been
proposed. One such class of devices is termed
hypermixers. These have the basic objective of
generating streamwise vorticity. It is hoped that the
streamwise vorticity will sweep through and entrain a
parallel-injected fuel jet, thereby increasing the fuelair interface area such that small-scale mixing is
significantly increased1.
Copyright © 1998 The American Institute of
Aeronautics and Astronautics Inc. All rights reserved.
Experiment
Injectors
The two hypermixing injectors tested in this
investigation are shown in Fig. 1 with the plane base
injector which forms the datum for the study. The
first hypermixing injector has a segmented blunt
trailing edge and is referred to here as the castellated
injector. The second hypermixing injector is a midplane version of the swept compression-expansion
ramp injector similar to that used by Davis and
Hingst2. Note that there are two and five nozzle
versions of the plane base and castellated injectors.
2 Nozzle Plane Base
2 Nozzle Castellated Swept CompressionExpansion Ramp
5 Nozzle Plane Base
5 Nozzle Castellated
Figure 1: Injector types.
Earlier studies have shown that certain castellated
trailing edge aerofoils have lower base drag than their
blunt trailing edge counterparts in supersonic flow3.
The drag reduction is primarily due to the entrainment
of fluid from the upper surfaces of the projections into
the recess regions. This geometry is similar to that of
a straight expansion ramp. It was hoped that the
benefit of reduced drag could be enjoyed at the same
time as the induced flow promoted the mixing.
Swept ramps are known to produce pairs of large
counter-rotating streamwise vortices to enhance
mixing1. The SCER (swept compression-expansion
ramp) should also produce a pair of large and very
strong counter-rotating streamwise vortices energised
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American Institute of Aeronautics and Astronautics
AIAA-98-0964
by the pressure difference between the faces of the
compression and expansion ramps2. In addition to
this source of streamwise vorticity, this injector is
subject to shock-induced vorticity due to the
positioning of the expansion surfaces and the
associated recompression shock waves4. Mixing may
be further enhance by a shock-vortex interaction
between the two5. It is expected that this injector will
have high drag. To be beneficial, therefore, the
increased heat release arising from its promotion of
mixing must be sufficient to offset the drag penalty in
any practical application of this injector.
from the injector base. An extra pressure transducer
was positioned 67 mm downstream from the injector
base in the constant area section. The Mach 2.5
condition was obtained by using a two-dimensional
supersonic diffuser in conjunction with a 304.8 mm
exit 34.9 mm throat conical nozzle, as shown in
Fig. 3 (b). The two angled plates that make up the
diffuser are positioned such that the reflected shock
waves pass outside the scramjet inlet and a Mach
number of 2.5 is produced at the diffuser exit.
Injector
(a) Splitter Plate
Flow
Scramjet Model
25mm
4.8mm
The experiments were performed on a laboratory
model of a scramjet, shown in Figs. 2 and 3, in the
T3 free-piston shock tunnel facility at the Australian
National University6. Tests were conducted at two
Mach numbers: M=2.5 and M=3.7. The model
consists of a rectangular duct having two side
windows to provide optical access with an injector
situated mid-way between the floor and roof of the
duct. The duct has a 25 x 52 mm cross section and is
425 mm in length. The injector is 4.8 mm thick, the
upstream section forming a splitter plate that extends
far enough into the flow so that shock waves
generated at the leading edge and the expansion waves
at the shoulders pass outside the duct. Pressure
transducers were mounted on the centre line in the
duct floor beginning 55 mm downstream of the
injector base and every 20 mm thereafter as shown in
Fig. 2 (a).
(a) Splitter Plate Injector
Flow
9
188mm
10 11
12
Side Window
180mm
Scramjet
Inlet
(b)
M=3.7
1 2
Side Window
180mm
3.5 deg
3 4 5 6 7 8 9 10
11
(b)
Diffuser
M=5.8
M=2.5
Shock
waves
Scramjet
Inlet
304.8 mm exit conical nozzle
Figure 3: Scramjet models (a) Mach 2.5 model and (b)
nozzle and diffuser set-up to produce Mach 2.5.
Flow Conditions
Each injector was subjected to three tests: noinjection, non-combustion, and combustion. The noinjection test allows the study of the wake of the
injector. In the non-combustion test, the fuel jet
exhausts into a co-flowing nitrogen stream so that
mixing proceeds without combustion. In the
combustion test, the fuel jet exhausts into a co-
25mm
1234567 8
4.8mm
150mm
87mm
F1
Scramjet
Inlet
88.9 mm exit contoured nozzle
Figure 2: Scramjet models (a) Mach 3.7 model and (b)
Mach 3.7 nozzle set-up.
A transducer was also mounted in the roof directly
above the injector and 35 mm upstream of its base, to
measure the inlet pressure of the duct. This
configuration was used for the Mach 3.7 condition.
The Mach 3.7 condition was obtain by using a 88.9
mm exit 25.4 mm throat contoured nozzle shown in
Fig. 2 (b).
M=3.7
Condition
H o=4 (MJ/kg)
Mach No.
Pressure(kPa)
Temperature(K)
Density (kg/m3 )
Velocity (m/s)
M=2.5
Condition
H o=2.9 (MJ/kg)
Mach No.
Pressure(kPa)
Temperature(K)
Density (kg/m3 )
Velocity (m/s)
Inlet
2 - Nozzle
Conditions φ = 0.5
5 - Nozzle
φ = 0.5
3.72
2.0
120
113
1020
160
0.44
0.17
2330
1960
Inlet
2 - Nozzle
Conditions φ = 0.8
1.9
71
170
0.1
1880
5 - Nozzle
φ = 0.8
2.5
90
1230
0.25
1710
2.0
46
160
0.07
1960
1.9
46
170
0.07
1880
Table 1. Flow conditions.
The configuration shown in Fig. 3 (a) was used for
Mach 2.5 experiments. In order to prevent thermal
choking7, the bottom plate was inclined at an angle of
-3.5 degrees, starting at a point 87 mm downstream
flowing air stream allowing both mixing and
combustion to take place. In all tests involving fuel
injection, hydrogen was used as the fuel. The
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AIAA-98-0964
freestream flow conditions including the total specific
enthalpy, Ho, are summarised in Table 1.
Mach 3.7 condition
Shadowgraph system
Combusting flow behaviour
The flows were visualised using a single-pass
shadowgraph system8. The light source was a flash
lamp-pumped dye laser with a 0.5 µs pulse width. A
schematic of the shadowgraph optical system is
shown in Fig. 4.
For each of the no-injection, non-combustion and
combustion tests, floor static pressures were measured
throughout the test time. The significant pressure
rises observed in the combusting flows are attributed
to heat release accompanying the combustion, with
higher pressures indicating
more
complete
combustion. Figure 6 shows the floor pressures
recorded for the five nozzle plane base injector at φ =
0.5.
SPHERICAL MIRROR f = 3.0 m
DYE LASER
FPDL
PRISMS
CUBE BEAM
SPLITTER
in place for
alignment only
LENS f = 40 mm
3.0 m
300
2 mm RADIAL APERTURE
FLAT
MIRROR
d = 0.24 m
FLOW
TEST SECTION
LENS f = 2.86 m
CAMERA
2.86 m
No-Injection
Non-Combustion
Combustion
250
Pressure (kPa)
ALIGNMENT LASER
200
150
100
NEUTRAL DENSITY FILTER
50
FLAT MIRROR d = 0.5 m
Figure 4: Single-pass shadowgraph system.
0
0
Results and Discussion
Duct f l o w
Figure 5 shows the results of a method of
characteristics calculation of the complex wave
patterns associated with the plane base injector wake9
together with the corresponding shadowgraph of the
flow. Flow is from left to right in this and in all
subsequent shadowgraph images.
Recirculation Zone
Recompression Shock
Base Expansion
Wake
Figure 5: Shadowgraph image of wake flow and
theoretical position of waves in the duct. M=3.7
The calculation accurately predicts the wave positions
and shock curvature arising from the interaction of the
reflected expansion fan with the recompression shock
wave. Other waves shown in the shadowgraph image
(top left of the image) are generated from a shock
wave/boundary layer interaction at the scramjet inlet
not included in the theoretical calculation.
100
200
300
Distance from Injector Base (mm)
400
Figure 6: Measured pressure profiles for no-injection,
non-combustion and combustion for the five nozzle
plane base injector at φ = 0.5. M=3.7
The maxima and minima on the two lower plots
indicate shock waves and expansion waves,
respectively, interacting with the wall. It is
interesting to observe that the non-combusting
pressure distribution is virtually identical to the noinjection distribution except far downstream of the
injector. It is also clear from the figure that the shock
waves in the combusting flow intersect the wall
further upstream than in the no-injection and nocombusting flows. This is due to the reduction in
Mach number caused by heat release. The combusting
plot shows a significant pressure rise due to heat
release just a short distance (100-130 mm)
downstream of the injector base.
Shadowgraph images taken from the three tests are
shown in Fig. 7 for the five-nozzle plane base injector
at φ = 0.5. There are clear differences between the
images of the no-injection and non-combusting flows.
The differences between the non-combustion image
Fig. 7 (b) and the combustion image Fig. 7 (c) are
more subtle. The difference is that the jet can be
easily seen across the entire image in Fig. 7 (c) but
not in Fig. 7(b).
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American Institute of Aeronautics and Astronautics
AIAA-98-0964
increase. This visual effect is accociated with
combustion, is probably due to a non-uniformities in
refractivity caused by the combustion heat release and
resulting turbulent flow.
(a)
(b)
(a)
(c)
(b)
0
40
(mm)
80
120
Figure 7: Shadowgraph images for (a) no-injection, (b)
non-combustion and (c) combustion for the five nozzle
plane base injector at φ = 0.5. M=3.7
(c)
(d)
In comparing Figs. 6 and 7, it seems that most of the
combustion takes place outside the optical viewing
window.
0
Shown in Fig. 8 is the effect of fuel mass flow rate.
The pressure plot shows no discernible difference in
pressure rise with φ until about 150 mm downstream
from the injector base. Beyond this point, the pressure
rise due to heat release increases with increasing
amounts of fuel flow as illustrated in the figure.
400
PHI = 0
PHI = 0.25
PHI = 0.5
PHI = 1.0
350
Pressure (kPa)
300
250
200
150
100
50
0
0
100
200
300
Distance from Injector Base (mm)
400
Figure 8: Measured pressure profiles for φ = 0, φ = 0.25
φ = 0.5, and φ = 1.0 for the 5 nozzle plane base injector.
M=3.7
The shadowgraph images in Fig. 9 are those
corresponding to the pressure traces in Fig. 8. The
visible effects of the increased fuel flow include the
increasing visibility of the jet in the region x <
60 mm and can be attributed to increasing quantities
of the cool, unburnt hydrogen fuel, with its high
refractive index, on the refractivity field in the duct.
The appearance of turbulence-like flow patterns,
visible at the right hand end of the field of view,
becomes more pronounced and seems to occupy a
greater proportion of the height of the duct as φ
40
(mm)
80
120
Figure 9: Shadowgraph images for (a) φ = 0, (b) φ =
0.25 (c) φ = 0.5, and (d) φ = 1.0 for the five nozzle plane
base injector. M=3.7
Influence of Injector Geometry
The pressures recorded for the φ = 0.5 combusting
flow tests with the five-nozzle and the two-nozzle
plane base injectors are presented in Figure 10. The
figure shows that the combustion pressure increase in
the flow with the five-nozzle injector greatly exceeds
that in the flow with the two-nozzle injector. The
only differences between the two flows are the initial
surface area of the fuel/air interface, which will be
greater for the five-nozzle injector, and the fuel jet
exhaust pressure, which will be greater for the twonozzle injector. These differences cannot influence the
flow chemistry; the flow temperature cannot be altered
significantly anywhere in the flow, nor can there be
any increase in precursor radical species. The pressure
differences must therefore arise from differences in
mixing efficiency between the two injector
geometries. This leads immediately to the conclusion
that the flow is mixing limited, as expected at the
outset of this work. A corollary is that the five-nozzle
injector produces faster mixing than the two-nozzle
injector.
Hereafter, combustion induced pressure rise will be
interpreted as a measure of mixing efficiency. It is
possible that the oblique shocks on the compression
ramps of the SCER may alter the chemistry via the
associated entropy rise but this is considered unlikely.
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American Institute of Aeronautics and Astronautics
Pressure (kPa)
AIAA-98-0964
No-Injection Two Nozzle
280
260
240
220
200
180
160
140
120
100
80
60
40
20
0
(a)
Combustion Two Nozzle
Combustion Five Nozzle
(b)
(c)
0
100
200
300
400
Distance from Injector Base (mm)
Figure 10: Measured pressure profiles for injection
from plane base injectors with two and five nozzles,
respectively, for φ = 0.5. M=3.7
Figure 11 shows the pressure distributions for the
two-nozzle plane base, castellated base and SCER
injectors for φ = 0.5 and M = 3.7 combusting flow
compared against that of the no-injection plane base
flow. The plane and castellated base injectors can be
seen to have very similar mixing efficiencies while
the SCER can be seen to significantly enhance
mixing efficiency beyond that of the plane base
injector datum case.
No-Injection Plane Base
280
5
45
(mm)
85
125
Figure 12: Shadowgraph images for injection from (a)
plane base, (b) castellated and (c) SCER injectors with
two nozzles, for φ = 0.5. M=3.7
Close inspection of Fig. 12 reveals that the fine
structured, turbulence-like region, associated earlier
with combusting flow, is similar in extent for the
plane and castellated injector flows but extends further
upstream for the SCER injector flow. This accords
with the pressure distributions which show the SCER
reaching a given pressure level upstream of the other
injectors.
Combustion Plane Base
260
Combustion Castellated
240
Combustion Swept CompressionExpansion Ramp
220
Non-Injection Plane Base
260
Combustion Plane Base
240
Combustion Castellated
180
220
160
200
140
180
Pressure (kPa)
Pressure (kPa)
200
280
120
160
100
140
80
120
60
100
40
80
20
60
0
0
50
100
150
200
250
300
Distance from Injector Base (mm)
350
400
Figure 11: Measured pressure profiles for injection
from plane base, castellated and swept compressionexpansion ramp injectors with two nozzles, for φ = 0.5.
M=3.7
Figure 12 shows the corresponding combusting flow
shadowgraph images, which appear to show an
increase in the wake-jet thickness as the injector is
changed from the plane base to the SCER injector.
This impression may be misleading as shadowgraph
is sensitive to the second spatial derivative of
refractivity and the castellated and SCER injector base
flows are highly three-dimensional. Interpretation is
therefore difficult.
40
20
0
0
50
100
150
200
250
300
350
400
Distance from Injector Base (mm)
Figure 13: Measured pressure profiles for injection
from plane base and castellated injectors with five
nozzles, for φ = 0.5. M=3.7
The comparison between the pressure distributions for
the combusting flows with the five nozzle plane base
and castellated injectors is shown in Fig. 13. The
injectors produce virtually identical pressure profiles
and higher pressures than their two nozzle
counterparts as noted earlier for the plane base
injector. The corresponding shadowgraph images are
also similar, as may be seen from Fig. 14.
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American Institute of Aeronautics and Astronautics
AIAA-98-0964
(a)
(a)
(b)
(b)
0
40
(mm)
80
120
(c)
Figure 14: Shadowgraph images for injection from (a)
plane base and (b) castellated injectors with five nozzles,
for φ = 0.5. M=3.7
5
Mach 2.5 Condition
The pressure profiles for the Mach 2.5 condition are
shown in Figs. 15 and 17. Fig. 15 shows the pressure
distributions for the two-nozzle plane base, castellated
and SCER injectors. The first point on the graph is
the pressure measured by the pressure transducer
mounted on the floor of the constant area section. The
lower curve is the pressures measured for no-injection
and shows the floor static pressure dropping due to the
increasing area of the duct. The pressure rises due to
shock waves seem to be suppressed by this
expansion.
No-Injection Plane Base
Combustion Plane Base
Combustion Castellated
Combustion Swept Compression-Expansion Ramp
160
140
45
85
(mm)
125
Figure 16: Shadowgraph images for injection from (a)
plane base, (b) castellated and (c) SCER injectors with
two nozzles, for φ = 0.8. M=2.5
plane base and castellated injectors. However, the
castellated injector seems to show some performance
improvement over the plane base injector beyond 200
mm of the injector base.
Figure 17 shows the pressure profiles for the two five
nozzle injectors at M=2.5. There appear to be some
significant differences between the injectors between
150 to 200 mm downstream of the injector base and a
similar trend to that shown by their two nozzle
counterparts 220 mm downstream of the injector base.
No-Injection Plane Base
Combustion Plane Base
Combustion Castellated
160
120
Pressure (kPa)
140
100
120
Pressure (kPa)
80
60
40
20
100
80
60
40
0
0
50
100
150
200
250
300
350
400
Distance from Injector Base (mm)
Figure 15: Measured pressure profiles for injection
from plane base, castellated and SCER injectors with two
nozzles, for φ = 0.8. M=2.5
The pressure rises measured for combustion show a
similar result to that obtained for the two nozzle
injectors at M=3.7, as the SCER injector produces a
substantially higher pressure due to combustion than
either of the
20
0
0
50
100
150
200
250
300
350
400
Distance from Injector Base (mm)
Figure 17: Measured pressure profiles for injection
from plane base and castellated injectors with five
nozzles, for φ = 0.8. M=2.5
The shadowgraph images in Figs. 16 and 18 also have
a similar appearance to those taken at M=3.7 with the
steeper shocks here indicating the lower freestream
Mach number.
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American Institute of Aeronautics and Astronautics
AIAA-98-0964
allowance for the uncertainties introduced by this
simplification, the calculation can be used to obtain
an approximate value for the percentage of injected
fuel consumed as a function of distance for each
injector/flow combination. As argued earlier, the
mixing limited nature of the flow allows this
percentage to be interpreted as a measure of the
mixing efficiency of injector/flow combination.
(a)
(b)
5
45
(mm)
85
125
140
120
Figure 18: Shadowgraph images for injection from (a)
plane base and (b) castellated injectors with five nozzles,
for φ = 0.8. M=2.5
Pressure (kPa)
One-Dimensional flow with heat
addition
Quantitative data obtained in this project are limited
to the duct floor pressures. In order to deduce other
flow properties such as temperature, Mach number
and percentage fuel burnt, from these pressure
distributions and the flow boundary conditions, a onedimensional finite difference calculation was
performed similar to that used by Stouffer et al.12 .
The governing flow equations for this model are,
ρuA = const.
Continuity :
Momentum : ρu
State :
ρu
80
60
40
20
0
0
50
100
150
200
250
300
350
400
Distance from Injector Base (mm)
Figure 19: Measured pressure profiles for no-injection
from plane base, castellated and SCER injectors with two
nozzles, for φ = 0.8. M=2.5
Figure 19 shows the no-injection pressure profiles for
the three injector configurations tested. These
pressures are virtually identical. Thus, the measured
combustion pressure profiles have a common
reference point. This allows the results from the onedimensional calculation for each injector configuration
to be compare directly with each other.
0.5
dh
dp dq
=u
+
dx
dx dx
0.45
0.4
0.35
p = ρRT
% Fuel Burnt
Energy :
du
dp
=−
dx
dx
Non-Injection Plane Base
Non-Injection Castellated
Non-Injection Swept Compression-Expansion Ramp
100
where q is the heat released.
It is assumed that the only reaction which occurs in
the flow is the water formation reaction,
0.3
0.25
0.2
Plane Base
0.15
Castellated
Swept CompressionExpanison Ramp
0.1
0.05
0
2 H2 + O2 → 2 H2 O
0
50
100
150
200
250
300
350
400
Distance from Injector Base (mm)
This assumption is justified on the grounds that since
the maximum calculated temperature is below 2500K
(see Fig. 21), which is too low for the operation of
endothermic reactions such as the NO and OH
formation reactions.
This model does not take into account the pressure
rise due to shock waves. It assumes that the change
of pressure with distance measured for the combustion
case is due to heat release only. With suitable
Figure 20: Percentage fuel burnt along duct for plane
base, castellated and SCER with two nozzles, for φ =
0.8. M=2.5
Figure 20 shows the percentage fuel burnt or
efficiency along the duct for the two-nozzle injector
configurations at M=2.5. The SCER injector shows
significantly higher efficiency throughout the duct
approaching nearly double that of either the plane base
or castellated injector from 120 to 200 mm
downstream of the injector base. The castellated
injector shows 10 -15% improvement in efficiency
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American Institute of Aeronautics and Astronautics
AIAA-98-0964
over the plane base injector
from 220 mm
downstream of the injector base. The way in which
the percentage fuel varies along the duct for the
castellated injector is not consistent with the notion
that the effects of hypermixing injectors ought to be
confined to the near-field. At the stage this cannot be
explained.
3000
Temperature (K)
2500
2000
1500
1000
Plane Base
Mixing Enhancement
Mixing enhancement primarily arises from the
stretching of the fuel-air interface. The plane base
injector achieves this via spanwise vorticity. An
explanation for the unchanged level of mixing shown
by the castellated injector might be found in some of
the mechanisms responsible for its lower drag. The
drag reduction is primarily due to the entrainment of
fluid from the top of the projections into the recess
region3 and this is a possible source of streamwise
vorticity. The entrainment of fluid also results in a
thicker wake neck region than that of the plane base
injector which reduces the amount of spanwise
vorticity3. This source and sink of vorticity may
counteract each other.
Castellated
500
Swept Compression-Expansion
Ramp
0
0
50
100
150
200
250
300
350
400
Distance from Injector Base (mm)
Figure 21: Temperature variation along duct for plane
base, castellated and SCER injectors with two nozzles,
for φ = 0.8. M=2.5
The temperature variation shows a similar profile to
that of the measure static pressure profile (Fig. 21)
with temperatures reaching a maximum 2100K,
2300K, and 2600K, for the plane base, castellated and
SCER respectively.
The Mach number distributions are shown in Fig. 22.
The freestream Mach number drops to approximately
M =1.7 therefore remaining supersonic throughout the
duct.
Increasing Mach number may also have an effect on
the mixing properties of the castellated injector. The
region of influence on the upper surface of the
projection and the flow expansion angle around a
blunt trailing edge decrease with increasing Mach
number. Therefore the effectiveness of presence of the
recess region in entraining high momentum fluid
down beside the fuel-jet to enhance mixing11 will be
decreased as the Mach number rises.
The considerable mixing enhancement shown by the
SCER is likely to be due to the generation of pairs of
streamwise vortices, in the wake of each ramp,
stretching the fuel-air interface3 possibly boosted by
shock-vortex interaction4. It is unlikely that there is
any jet-shock interaction mixing augmentation as the
recompression shocks and the fuel jets are well
separated in the current implementation of this
injector type2.
3.5
Conclusions and further work
3
Mach No.
2.5
An experimental comparison
between
two
hypermixing injectors, at two Mach numbers has
been presented using floor static pressure
measurements and shadowgraph images. A onedimensional model was used to determine combustion
efficiency, temperature and Mach number distributions
along the combustor. The following observations
were made:
2
1.5
1
Plane Base
Castellated
0.5
Swept Compression-Expansion Ramp
0
0
50
100
150
200
250
300
350
400
Distance from Injector Base (mm)
Figure 22: Mach No. variation along duct for plane
base, castellated and SCER injectors with two nozzles,
for φ = 0.8. M=2.5
(i) Combustion and attendant heat release and pressure
rise occurred in flows in which hydrogen fuel was
injected into an incoming air flow.
The one-dimensional calculations for the M=3.7
condition are not presented due to dominating effect of
the shock waves on the pressure profiles in the
constant area duct.
(ii) Combustion induced pressure rise increased with
increasing equivalence ratio, φ, up to φ = 1 which
was the highest value tested. The fact that the pressure
8
American Institute of Aeronautics and Astronautics
AIAA-98-0964
rise due to combustion for φ = 1 was not twice that of
φ = 0.5 suggests that a plateau would be reached.
(iii) A consistent interpretation of the flow was
obtained from shadowgraph, pressure distributions and
method of characteristics calculation of the flow field.
can be made, such as the inclusion of more chemical
reactions, the effects of the multiply reflected inlet and
recompression shocks and possibly some detailed
modelling of the mixing .
Acknowledgments
(iv) The flows are mixing limited. Relative mixing
efficiencies could therefore be deduced from the
pressure distributions produced by combustion heat
release.
(v) For the same injector base geometry, greater
mixing was obtained by increasing the number of fuel
exhaust nozzles. This is one manifestation of the
well known phenomenon that increasing the surface
area of the fuel/air interface, increases the mixing rate.
(vi) Of the injectors examined here, the SCER
injector proved to have the highest mixing efficiency
for all flow conditions. In the M = 3.7 flows, the
mixing efficiencies of the castellated and plane base
injectors were found to be very similar. In the M =
2.5 flows, the castellated injector was found to have a
slightly greater mixing efficiency than the plane base
injector.
(vii)
Application of the one-dimensional model
showed that approximately 30% of injected fuel was
consumed using a two nozzle plane base injector
while the proportion consumed rose to 45% with the
SCER injector.
(viii) The one-dimensional model shows, amongst
other things, that the combusting flow remained
supersonic throughout the combustor.
(ix) It seems that any flow effect generated by an
injector geometry needs to be either very strong or
very close to the fuel-air interface to have an effect on
the mixing due to the short residence times of these
flow effects.
The shadowgraph images provide some qualitative
information about the flow field, but due to the threedimensional nature of the flow field a quantitative
conclusion cannot be drawn from them. Experiments
are currently in progress to investigate the injector
configurations further in an unconfined, noncombusting flow environment using planar laserinduced fluorescence (PLIF). PLIF can be used to
image a two-dimensional slice of the flow field; from
these images a three-dimensional picture of the flow
field can be constructed. It is hoped that these
experiments will lead to a better understanding of the
mixing enhancement properties of the injectors.
Further improvement to the one-dimensional model
The authors would like to thank Paul Walsh for his
technical expertise and Paul Tant for his technical
assistance with the tunnel and scramjet. Inputs from
Jodie Fox, Sean O'Byrne, Paul Danehy, Phil Palma
and Joe Kurtz are gratefully acknowledged. Special
thanks goes to Sean for his help with the onedimensional model flow model.
This research has been financially supported by the
Australian Research Council.
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American Institute of Aeronautics and Astronautics