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J O U R N A L
O F P R O P U L S I O N A N D P O W E R
Vol. 10, No. 2, M a rc h -A p ri l 1994
Advanced
Injectionand
Mixing
Techniques
for
ScramjetCombustors
D a v id W . B o g d a n o f f*
Eloret Institute, Palo
Alto,
California 94303
Scramjet combustor fuel injection and mixing enhancement techniques are reviewed. The injection techniques
include
hole injection from
th e
combustor wal l, slot injection parallel
to the
f low,
an d
injection
from
struts
and
the rear of ramps. Three new advanced mixing techniques are presented. Th e first is a combustor, curved so
that
buoyancy forceswill
aid in the
penetration
of the
fuelacross
the
combustor.
Th e
second
is
pulsation
of the
fuel injectors
to
increase penetration
and
mixing.
A
fluidic technique,
a
modified Hartmann-Sprenger tube,
is
identified as a
strong candidate
to
generate
the
pulsations.
Th e
third
is the
injection behind pylons
to
allow
deep penetration into the air stream. This technique is likely to produce high base pressures on the injector
structure, particularly if base burning is encouraged. Curved or slanted pylons can be used to increase the
recovery of
fuel
je t
momentum.
Th e
potential
of the new
mixing techniques
to
increase scramjet engine per-
formance
isassessed.
I.
Introduction
S
CRAMJET operat ion a t flight Mach numbers of 10-20
is generallybelieved
1
-
2
to
require mixing
an d
combustion
at
combustor inlet Mach numbers
of
roughly one-third
th e
flight M a ch n u m b er .T he fuel,g enerally takento beh y d r o g en ,
is
injected an d
m u s t
mix and
burn
in the
very short combustor
stream residence time. Thrust is generated as relatively small
differences between th e large engine inlet an d outlet m o-
menta.There are inevitably
frictional,
shock, and other losses
in
th e main momentum stream. Additional losses due to in-
jection,
mixing, an d
combustion
of the
fuel must
be
kept
to
a minimum
a n d ,
at the
same t ime,
th e
most complete
fuel-
air
mixing
an d fuel chemical energy release must be achieved
to
maximize
thrust. The most imp ortant add itional losses due
to injection an d mixing comprise shock wave losses on the
fuel jets , shock wave losses
an d
pressure
an d
friction drag
on
injector
mechanical structures
(i f
any), shear layer mixing
losses
betweenfuel an d air, an d loss of the m o m e n t u mof the
fuel
jets
(i n
some configurations).
Many
be nchmark studies
3
11
have been done with circular
hole
injectors at anglesof 90 deg (normal to the stream
flow).
O t h e rs tudies
1 2
1 6
have been done with circular hole injectors
at
angles rangingfrom 0 deg (parallel to the stream
flow)
to
150 deg (angled
upstream against
the
stream flow). Slot
injectors
17
21
ar e
usually oriented
so
that
th e
fuel
an d
stream
velocity
vectors ar e aligned. A n u m b e r of techniques have
been
used
to
enhance
th e
mixing
of the
injected
flow
with
th e stream flow. Perhaps the most basic is to use some kind
of
strut
22
-
23
or extended tube
15
so that th e actual injector
orifice is
lifted
away
from
th e
main stream wall into
th e
body
of the stream
flow.
Other mixing enhance me nt tech-
Injector types and m ixing enhancem ent techniques are re-
viewed
in Sec. II. Three new advanced mixing techniques are
described in Sec. Illand their potentialto increase scramjet
engine performance is assessed.
II.
Injectors and
Mixing Enhancement
Techniques
A.
Injectors
W e
first
discuss injectors which are on the wall and do not
have mechanical structure protrud ing into the flow. W edefine
6 as the angle between the injector jet and
th e
freestream.
For example, 6 = 0 deg an d 6 = 90 deg denote injection
parallel and norm al to the stream, respectively. W e also in-
troduce
th e following
definitions:
R
ci
is the
ratio
of the mo-
mentum of the injected jet to that of the adjacent freestream,
an dR
m
is the
ratio
of the
mass flux
of the injected jet to
that
of th e adjacent freestream. R eferences 3- 7 present data on
norm al circular hole injectors. R eferences 3-5 all study a
singlesonic injector injecting into a Mach 4 stream. For these
cases R
q
ranged from 0.5-3.0
an d
concentrat ion measure-
ments were taken betweenX ID = 1 an d
X ID
= 200, where
D
is the
injector port diameter
an d
X
is
distance downstream
from
th e injector. The penetrat ion w asf o u n d to vary asI ? -
5
,
/C;
54
,
or
R
3
in
R efs. 3-5, respectively. R eferences
4 and 5
found th e penetrat ion (a t maximum injectant concentrat ion)
to decrease between X ID = 1 an d XID = 15-40, an d then
to increase as X ID increases towards 200.
These
latter two
references also found the maximum injectant concentration
to vary as
(X/D)~
()
-
5
.
R eference 6 studied a single sonic in -
jector injecting into a Mach 2
stream.'R^
ranged from 4.4-
5.3. Th e height of the Mach disc w as f o u n d to vary as
/ ? J J -
5
,
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184
BOGDANOFF: TECHNIQUES FOR SCRAMJET COMBUSTORS
whe r e
7,
is the height of the Mach disc, D * is the diameter
of th e throat of the
injector ,
an d M j is the
injector
je t Mach
n u m b e r . The calculated integrated p lum e trajectories were
found
to
approach
horizontal as
XID* reaches
10, and to be
very
similarf or
equal
massflows at
varyinginjector
pressures.
A tXID* = 10, the penetrat ions of the calculated trajectories
were fo u n d to vary as (p////?,,,)
0
-
24
0
-
29
, wh e r e p
tj
is the
pitot
pressure of the injector j e t , and/?,,,is the s tream
pitot
pressure.
R e fe re nce s 12-15 present data on
angled circular hole
in -
jectors. Some key
parameters
for the
data
of
these
references
are
given
in
Table 1.
In
Table 1,S denotes
the
lateral spacing
between the holes an d 7 V
in j
, th e
n u m b e r
of
injectors.
A wide
range of R
t]
an d 0 values ar e
covered
by this
data .
W e
note
that th e injector of R e f . 15 is
above
th e wall,
outside
th e
boundary layer, while the injectors of R efs.
12-14
are on the
wall.Also,significant differences were
found to
exist between
th e flow patterns of
single
and multiple injectors.
8
W e will
retu rn to
this
pointlater.
R e f e re nc e s
12 and 15m ain lypresent
data on
initial
je t
penetrat ion, either
the height of the Mach
discor the centroid of the injector gas plum e.The
penetrat ion
was found
to
vary
as
R
1
-
5
fo r
b o t h references,
and as
1/(1
+
cos 0)or (cos 0)-
8
in
R e f s .
12 and 15, respectively. R ef eren c e
15 therefore finds th e m a x i m u m
penetrat ion
at
6 =
90
deg,
while
R e f . 1 2
finds
the maximum
penetrat ion
at the highest
value
of
8
tested (120 deg). Increasing penetration for the
data
of Re f. 12 for
0 increasing above
90 deg may be the
result of
boundary-layer
separation.
R eference
12
also showed
increased penetration with wall blowing upstream of the in-
jectors. Re ferences 13 and 14
measure
injected gas
concen-
trations
20-120
injector
hole
diameters
downstream of the
injectors an d fo u n d th e penetrat ion at these locations to be-
have quite differently than
th e
initial penetrations studied
in
R e f s .
12 and 15. R ef er en c e 13, using
five
injectors, found
penetration
to
increase withdecreasing
0,
being 1. 5 1. 6
times
greater at
6 =
30 deg than at
6 =
90
deg.
Reference 14,
with
a single
injector,
found
penetrat ion better
at
6 =
30 deg
than
at 0 - 15 deg. R e f e r e nc e 14 also fo u n d the
penetrat ion
to
vary
as
/ ? -
7
- -
9
at
XID =
20
an d
as
K J-
4
- -
8
at
X ID =
90.
From R e fs . 13 and 1 4, it thus
appearsthat
t he
best pene trat ion
well
downstream of the injector
occurs
at 0 30
deg.
R e f -
erence 13
compared data from single
and multiple
injectors.
The single jet apparently initially (at XID 30) penetrates
less
an d mixes
more
poorly, but further downstream (a t X lD
120), th e
single
jet is
superior
on both
accounts.
References 8-11 present
data on
tandem normal circular
hole
injectors. In the
work
of
R e f s.
8-10there are two
equally
sized injector holes downstream
of a
rearward facing step.
Time average and rms velocity
data
an d
injectant
concentra-
tion data were
obtained
using
laser-based techniques.
T he
path
of the
injectant p lum e
and the decay of the
injectant
concentration
were d et er m in ed . In the
work
of R e f . 11there
ar e
five
injector holes with diametersincreasing in the down-
stream direction. No concentration-based
penetrat ion
data
ar e available from R e f . 11. W i th ta nde m jets , one may
argue
that
an upstream jet can be viewed as creating streamwise
vorticity which
then
carries
th e
injected
gas of
downstream
jets
further into the main stream. It remains to be d em o n -
respect to the mainstream. The yaw will produce added
streamwise
vorticity in the f l ow. W e ll dow nstr e am of the in-
jectors ,
yaw was found to
increase
the lateral spreading of
th e
injected
gas by
10-30%.
R eferences17-21studied slot
injection parallel
to the
main
stream.
For all
cases
th e
parallel injection
was at the
wall .
The fr e e str e am Mach numbe r s var ie d from 2 to 3 and the
freestream total
t e m p e r a t u r e
w asne a r 300 K for R efs. 17 , 18,
and 21, and the freestream static
t e m p e r a t u r e
w as 1 30 0 K
fo r R e f s .
19 and 20. The parallel
inj e ct ion
w as
su b so n ic
air
fo r
R e f .
17 ,
Mach
1.7 air for
R e f s .
18 and 21, and
Mach
1
hydrogen
fo r
R e fs .
19 and 20. The
apparatus
of
R e f .
21
h a d ,
in addit ionto a parallel injection slot,11
tubes
n o r m a lto the
flow, downstream of the slot and spanning the slot he ight.
Injection normal to the freestream flow could be made
through
these tubes.Testswere
ru n
with
no nor m a l injection an d
with
injection of air at Mach numbers of 1.0 and 2.2.
W it h
parallel
injection,
theoretically th e full
m o m e n t u m
of the injected gas
is available to add to the main
stream.
For the geometries of
R e f s .
17-21,
th e t h i n , high-gradient boundary layer at the
wall
downstream of the
slot
m ay
result
in significant je t mo-
m e n t u m
losses.
In general ,these experiments showed rather
poor
mixing. Mixing was improved by the
i m p i nge m e nt
of
oblique
shocks
18
-
20
on the
jets
and the addition of normal
injection.
21
The second
main
classo f
injectors
ha s
mechanical structures
of varioussortswhich lift
the injector
ports
out
into
th e
main-
stream.Muchimproved pen etrationan d
mixing
can beachieved
in this way at the
cost
of a ddi t i ona l m om e ntum losses due to
th e mechanical
structure.These losses comprise
pressure an d
friction
drag forces
on the structure and
addit ional
shock and
shear layer losses due to the s tructure. R eferences 22 and 23
study injection from jets
on
strutsentirely spanning
th e
stream.
Th e angled circular hole injector of Re f . 15 (a simple be nt
t u b e) is an example of an
inje ctor
"strut"
which
does not
extend completely across the s tream. E st imates made of the
m om e ntum los s due to
pressure
drag
forces
on these
struts
show
that
these
losses
can be very substantial compared to
representative estimated
engine thrusts. Inje ction can also be
at the
downstream
end of a
ramp.
24
"
28
W e
will
consider
this
type of
injector
in the following section on mixing
enhance-
m e n t .
B. Mixing Enhancement
A n u m b e r o f t h e t h e o r e t i c a l a n d e x p e r i m e n t a l s t u d -
j
es
i 8 , 2 o , 2 6 , 2 7 , 2 9 - 3 2 j^ye investigated th e effect of passing the in-
jected jets through shock
waves
or expansion
wave systems.
A large fraction of
these
studies show substantial increases in
mixing due to the
addition
of shock or expansion waves . Dif-
ferential
acceleration
of
different
density
gases
by wave sys-
tems
will
produce
baroclinic
torques
which will ,
i n t u r n ,pro-
duce
vo rticity
an d increase mixing. In two- dime nsional flows,
th e
additional vorticity
ca n
only
be in the
spanwise
direction;
in three-dimensionalflows
(i .e. ,
with
circular
j e ts ) , addit ional
str e amwise
vorticity
will
be
created
by the
wave s.
27
-
31
O n
two-
dimensional
interfaces
or withtw o-dimensional je ts , increased
velocity differences
due to wave
systems
will
increase
th e
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BOGDANOFF: TECHNIQUES FOR
SCRAMJET COMBUSTORS
185
H elm ho ltz or
R ichtmye r - Me shkov instabil it ies ,
30
'
34
-
35
respec-
tively.
Thus, there
ar e a numbe r of me chanisms by which
shock or expansion waves can
increase mixing
between the
jets and the
mainstream.
Injection
at the dow nstream end of ramps is studied in R efs.
24-28.
Figuresla-lc showthe ramp geom etriesof R e f s .24,
26 , and 28, R e f . 25 and R ef. 27, respectively. In
Fig.
1,
FS
denotes
freestream flow, /,
injector
je t f l ow, S,
shock wave
an dE, expansion
fan.
In the
rear
views, th e
dense
hatching
denotes
th e injector port. The second
shock
in Fig. la is
assumed to be reflected from a top wall
which
is not shown.
In
R e f s .24-26 and 28, both swept and unswept rampswere
studied. Swept an d u n s w e p t
ramps
look identical in the side
views.
F or
unswe pt
ramps, the
corners
A and B remain par-
allel
as one moves upstream
from
th e injector port
location.
For
swept
ramps,
these corners spread
outwards , as shown
in th e
rear
views. For the swept ramps, th e sweep angle is
-10
deg. R eferences24-26
each make comparisonsof
swept
vs unswe pt r a m p injectors and
find
that greater streamwise
vorticity
an d
bet ter
mixing ar e
obtained
with swept
r a m p
injectors. H ow e ve r , la rge r m om e ntum losses
ar e
found
25
with
the
swept
ramps. The ramp injectors of
R e f s .
25 and 27, in
contrast to those ofR e fs . 24,26 , and 28,
have
awallgeometry
whichproduces a strong oblique shock directly at the injector
port. This
shock
wave will produce
baroclinic
torques
an d
generate streamwise
vorticity
in the region of the main stream -
jet boundary (if a
density
difference exists), an d
hence,
im -
prove
mixing.
27
III.
New Advanced M ixingT echniques
In this section, three
new
advanced mixing techniques
ar e
presented and preliminary assessments ar e
m a d e
of the po-
tential which they
offer
fo r
increases
in scramje t
engineper-
formance .
A. Curved Combustor
Figure 2 shows sketches of two scramjet engines .
These
sketches are schematic only and do not
represent
actual can-
didate geometries. The fuel inje ctor locations are d en o t ed by
/. A
conventional
straight
c om bus tor
is
shown
in
Fig.
2a. The
proposed new concept is to curve the combustor, as shown
in Fig.
2b, so
that buoyancy forces
in the accelerating refer-
ence
system of the curved main flow
will
tend to carry th e
fuel
plume
from
the injector
across
th e
combustor, ensuring
good
penetration
an d
mixing wi th out having large
high-drag
injector
struts
in the flow. This concept is a
de ve lop m e nt
from
earlier work showing mixing en h a n c em en t due to shock or
expansion waves
18,20,26,27,29-32
(see Sec.
II.B.).
Fig. 2 Sketches of scramjet engines with a straight and b curved
combustors.Fuel injector locations
are
denoted
by / and
diffusersho ks
are shown
dashed. Geometries and shock locations are conceptual and
schematic only.
T he density
differences
necessary to d riv e th e fuel across
th e
combustor s tem from the low molecular weights of the
fuel
an d
combustion products
and the
he ating
due to com-
bustion.
W e
will
e x a m i n e briefly th e
buoyancy dr ive n accel-
eration
o f afuel p l u m e o f a different d en si ty t h a n the am b ien t
gas. L et p be the
density
of the p lum e ,
p
()
the density of the
sur r ounding ambie nt gas, and
de fine
= p/ p
()
. To ge tR
(l
, th e
acceleration
of the plume
(neglecting
viscous effects) divided
by
th e acceleration of the
local reference system,
w e m u s t
consider
th e acceleration of the plume gas plus the acceler-
ation of the ambie nt gas sur r ounding the plume . The acce l-
eration of the latter .gas
produces
an additional apparent
mass"effect, which isdiscussedfor an infinite circular cylinder
in
Hunsaker
a n d R i g h t m i re .
36
For the
circular
cy lin d er, th e
additional
apparent
mass is
e qua l
to
( v o l u m e
of cylinder) x
p
0
.
From
the solution for the
flow
over a
sphere
given in
A n d er s o n ,
37
a few
pages
of
algebra suffices
to
show th a t ,
fo r
a
sphere,
th e additional apparent mass is e qua l to
0.375
x
(volume of
sphere)
x
p
()
.
The
total
mass (buoyant gas +
additionalapparent mass
from
th e
ambient gas)
to be
accel-
erated
and the
b u o y a n t
force
available
can now
readily
be
calculated as functions off for cylindrical and spherical buoy-
an t bodies, yielding
R
(l
as funct ions of
f.
For a cyl inde r , R
tl
= (I
-
)/( + >
andfora
s p h e r e , R
a
=
(1 - )/(().375
+
) .
T he
value s
o fR
a
differ
fo r
cylinders
an d sphe r e s since
the additional apparent masses are proportional to different
fractions
of the
buoyant
volum e for the two
geom etries. Spheres
always have th e greater R
{l
values;
e . g . ,
fo r f
=
0.5, fo r
spheres R
a
=
0.571,
while
for cylinders,
R
(l
=
0.333.
T h us ,
it
would seem to be beneficial, on this account, to
break
up
th e
fuel p l u m e .
W e
will r e tur n
to
this point
in the
f o llo w in g
section.
R e f e r e nc e
1
gives conditions
in a
representative hydr oge n
fueled scramjet c om bus tor at the
inlet ,
exit, and fuel injector
throat .
W e consider th e
case
at Mach 15
with
a fuel equiva-
lence ratio (ER) of 1.0. T he c om bus tor
inlet
conditions are,
pressure =
0.568 x
10
6
dyne /cm
2
, te m p e r a tur e
= 1908 K ,
Mach number
= 5.189,a ndvelocity = 4.34 x
10
5
cm/s.
Based
on possible material limitations, we take th e
fuel
stagnation
temperature
to be
1000
K ,
rather
than
1667
K asgiven in
R e f.
1. For
these
conditions, the density of pure u n b u r n e d fue l ,
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186
BOGDANOFF:
TECHNIQUES
FOR
SCRAMJET COMBUSTORS
cept with me thane (or highe r hydr ocar bon) fuels appears less
desirable
than
with
hydrogen
fuel
due to the much
weaker
buoyancy forces for u n b u r n e d fuel p l u m e s .
R e t u r n i n g
to
hydr oge n
f u el ,
with
0.47,
as
es timated
above, R
(l
can be estimated from th e equations given above
to rangefrom0.36-0.63,depending upon thedegreeo f
breakup
of
the
fuel/products p l u m e s .
If an average
R
a
of 0.40 can be
achieved
and thecurvedc om bus tordeflects a
total
of 2.5 times
its height,
th e
fuel/products p l u m e s will cross
th e c om bus tor ,
based on the
present
simple
inviscid analysis.
T w o
factors
ca n
reduce the combustor
curvature
and/or the average R
a
value
required below
those
values
given above using
th e present
mode l.
First, the injectors (even if on the wall)
will
presum-
ably
achieve
some penetration
on the ir own, so a
full
crossing
of
th e combustor by the fuel/products p l u m e s ,
driven only
by
buoyancy
effects, is notrequ ired. Second,fuel injection could
begin
some what
upstr e am
of the c om bus tor
proper.
If only
relatively small a m o u n t s
of hydr oge n combustion take place
"early,"such
injection m ay
produce
little
reduction
in
engine
perfo rm an ce.
W i th
a curved c om bus tor ,
there
are concerns about
n o n -
uniform wall
heating, momentum loss
and the generation of
shock
waves,
and the
attendent loss
of
stagnation pressure.
To assess the severity of
these
effects,
a two- dime nsional in -
viscid planar
C F D
solution
w as
computed
for a curved
com-
bustor
(without
fuel injection).
The
inlet conditions
were those
giveni n
R e f.
1 for Mach 15.The gas was
ta ke n
to be
calorically
and volumetrically
perfect with
a
molecular weight
of 28.84
g/g-mole
and a
specific
heat ratio of
1.4.
The com bustor chan-
ne l
is comprised of two circular arcs
with
20-deg
included
angle, inne r
radius 425 cm, and height 10 cm. The length/
height
ratio of the channel is 15, and the channel
deflects,
over its
len g th,
2 . 7 5
time s its
height.
The gridding
used
w as
50 x
100, with
100 cells in the streamwise direction.
The CFD
solution shows
a
series
of
continuous expansion
an d compression
wav e
systems,
but no
shock waves,
very
little
stagnation
pressure loss, and a
m o m e n t u m
loss of 0.28% of
th e inlet stream
m o m e n t u m . A s
expected,
due to the flow
curvature,
there
are substantial density
gradients
in the
height
direction. The
density
at the
outer
wall averages about 1 . 3
1.4 times the mean value, with
m a x i m u m
densities of
1.5-
1.6 times the mean value. At the inner wall, the density
averages
0.7-0.8
time s
the mean
value ,
with
m i n i m u m
den-
sities of
0 . 5 time s
th e
m e a n value . Heating would
be ex-
pected to be increased at theouter
wall
an d
decreased
at the
inne r
wall
roughly
in proportion to these densities. Clearly,
attention
must
be
paid
to the
wave systems
an d
density var-
iations
which
would occur in a curved combustor , but our
preliminary CFD calculations do not show any very bad flow
patterns which
wo u ld
weigh heavily against the curved com-
bustor
concept.
It is also likely that c om bus tor wall profiles
more sophisticated than simple circular arcs could produce
some reduction in the s trengths of the wave systems and the
magnitudes of the m om e ntum loss and de nsi ty variations. F or
ex am ple,
supersonic flow can be turned in symmetric elbows
K), and produce rapid, intense pulsations. The
fuel
je t flow
should be reduced tonear zero at the pressure m i n i m a .
I t seems unlikely that mechanical pulsation devices could
produce
frequencies
fast
e nough
to pulse the jet
each
time it
travels
a small
n u m b e r
(e.g. , 2-3) of diameters . (I t would
also
be
difficult
to m a k e such devices
survive
the high-tem-
perature e nvi r onm e nt . )
B y
analogy with
th e
subsonic
jet ex-
periments,
39
-
40
such rapid pulsing would likely be necessary
to attempt to break up the jet and increase them ix in g .
Slower
pulsing would merely produce
a quasisteady jet
with
a pe-
riodically
varying
flow r a t e t h is would
be
expected
to
pro-
duce little increase in
mixing.
Hence, it seems likely t h a t
fluidic
techniques
would
be
required
to
pulse
th e
j e t . M a n y
fluidic
oscillator techniques,e . g . ,
th e
cylindrical
resonant
cav-
it y
technique
41
and the beam
deflection amplifier with feed-
back technique,
42
have
th e
disadvantages
of
producing
lo w
ampli tude
(and
insome cases, low frequency) oscillations and
require
large,
b u lk y flow passages.
T h e H a r tm a nn- S p r e nge r
(H.-S.)tube,
43
"
48
inconstrast, iscompact an dproduces very
large
amplitude, rapid pulsations . Peak-to-peak oscillation
amplitudes
observed
43
"
46
in
H.-S. tube s
ar e
ro u g hly e q u a l
to
th estagnationpressurein the excitingj e t .Hence,w econsider
th e
H.-S.
tube to be a s trong ( though not
necessarily
th e
o n l y )
candidate
to
pulse
th e
fuel
jets .
A possible
injector
geometry using an a n n u l a r
H.-S.
tube
isshown
schem atically
in
Fig.
3 . ( W e note
th a t
th e
H a r t m a n n -
Sprenger tube
was men tioned in R ef . 49, but as
applied
to
th e production of
oscillatory shock waves
and not to fuel
injectors.)
B e low,
w e
m a k e
an es timate of how
finely
th e
H.-S.
geometry
of Fig. 3 can break up the jet . W e
assume
that
th e
length
of the
H.-S. tube
is 1.4 cm. To the same scale
the nozzle
throat
and exit diameters are 2 and 4 cm,
respec-
tively.
W e take th e
fuel
to be
ideal
hydrogen at a stagnation
temperature
of 1000 K. Using the
open-closed
organ
pipe
fo rm u la
for the
H.-S.
tube and
assuming
it to
operate
at 1000
K ,
its frequency is calculated to be 43kHz. Using standard
isentropic
flow
tables ,
th e
nozzle exit
velocity can be calcu-
lated to be 4.3
km/s. Hence,
th e injected fuel plume moves
about 2.5 nozzle
exit
diameters pe r
cycle.
This is
sufficiently
fast
to
break
th e p lum e up if, of
course,
th e
pulsation
of the
flow rate
ha s
sufficient
amplitude.
Two important points
m us t
b e m a d e here. First,
Fig.
3 is
conceptual
only; obviously a n u m b e r o f
geometries
and op-
erating conditions would have
to be
tes ted
and the most suit-
able of
both f o u n d .
Second, it is well k n o w n that very
high
temperatures can be produced in
H.-S. t u b es .
4 3
4 8
This could
be very detrimental to the survivability of
such
a
device
in
the scramjet
c om bus tor
e n v i r o n m e n t .H o w e v e r,w e
notefrom
th e same
references
that the temperatures reached are very
m u c h
dependent on the geometry and operating
conditions
of
th e
H.-S.
tubes .
Also,
m a n y of theH.-S. tube geometries
and operating conditions described have
been
developed spe-
cifically to
produce these
high
temperatures.
In the course of
developing aH.-S. tube/nozzle geometry
suitable
for scramjet
c om bus tor injectors, emphasis would be placed on
pr oducing
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BOGDANOFF: TECHNIQUES FOR SCRAMJET COMBUSTORS
187
strong
pulsations while
limiting
th e temperatures
reached
to
valuesacceptablebasedo n m aterials considerations . The data
given in R e f s . 43-48 suggest
that
it may well be possible to
achieve this goal.
C. Pylon
Injection
Injection Behind
Pylons)
The basic
concept
is shown, for a
90-deg inje ctor ,
in
Fig.
4. The pylon allows th e fuel jet to
penetrate
deeply into the
combustor. This
concept
is
related
to the r a m pinje ctor s,
24
~
28
b u t , in
part ,
ha s
some advantages
since the
fuel does
not
pass
thro u g h the pr otr uding inje ctor s tr uctur e . For simplicity, w e
have shown
a circular injectorport in
Fig.
8, but by
elongating
th e port
in the
direction
of the
stream flow,
th e
same size
pylon
could
be made to serve for an
injector with
greater mass
flow. Particularly with
an
elongated
injection
port,
the pylon
can
be
likely
be made smaller
than
th e
correspondingramp
24
"
28
injectors
since
th e fuel passage is inside the r a m p in the latter
cases.
Also,
th e
rear (base) pressure
on the
pylon
is
likely
to
be
larger than
that for
ramp
injectors due to the
presence
of
an adjacent parallel jet and
possibly
base c om b us t ion. ( R e f -
erences 50 and 51 show that large reductions in the
base
drag
of
projectiles can be
achieved
using
base
bur ni ng. )
Hence,
th e drag of the pylon is likely to be considerably less than
thatfor the
r a m p
o f
r a m p
injectors. The
90-deg
inje ctor
shown
in
Fig.
4 has the
disadvantage
that th e fuel je t m o m e n t u m
will be totally
lost. More
viable pylon injector
concepts
ar e
showni n
Figs.
5a and 5b.Here, th e
pylons
are tilted or
curved
to
allow
partial
recovery
of the
fuel
je t
m o m e n t u m .
The
pres-
sure on the base of the curved
pylon
will be increased due to
th e
tu rn in g
of the
jet.
T w o
additional possible
refinements of the
pylon
injector
will now be presented (as applied to a 90-deg inje ctor ) . Fig-
ures
5c and 5d show the concept of the
partial trapping
of the
fuel
jet in the
recessed
rear
of the pylon.
This concept would
help to resist the tendency of the jet to be
deflected
away
from th e
rear
of the
pylon
by the
mainstream flow
an d con-
seq u en t reduction of jet
penetration. Figure
6 shows the ta-
pered
pylon
concept. The tapering, whichneed not be l inear ,
would be
adjusted
to
optimize
th e
vertical distribution
of fuel
in the combustor . For simplicity th e above concepts were
shown
applied singly to a 90-deg
injector ,
but
they
could
very
well
be applied to tilted or curved
injectors and/or
combined
with
each
other.
The three new advanced
mixing
techniques presented in
this
section could be used
together
in the
same
combustor to
maximize combustor pe r for mance .
a)
b
Fig.
5
Pylon injector with
curved
rear
to
trap
the
injector
jet: a)
tilted pylon
injector,
b)
curved
pylon
injector,
c)
side view
section),
and d) top view.
c)
Fig. 6
Three
viewsof
tapered
pylon
injector
concept:a) top view, b)
side
view
section), and c)
rear
view
section).
D. PerformanceAssessment ofProposed Advanced
MixingConcepts
A t
this point,
an assessment ismade of the potential for
scram
jet
engine performance increases offered
by the ad-
vanced
mixing
concepts presented in thethree pr e ce ding sec-
tions. W econsider asrepresentative th ecombustor inl e t
con-
ditions fo r flight at Mach 15 given in R e f . 1.
These
are 1)
pressure
=
0.568
x 10
s
d y n e/cm
2
,2)
t e m p e r a tu r e
= 1908K,
3)
Mach n u m b e r
=
5.19,
and 4)
velocity
= 4.34 x 10
s
cm /
s. For
these conditions, using
a
simple
five
species
ai r
eq u i-
librium solver, we calculatethat th e
m o l e c u l a r
weig ht = 28.8
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188
BOGDANOFF: TECHNIQUES FOR SCRAMJET COMBUSTORS
instead, a fuel total te m p e r a ture of 1000 K .
Using
a scramjet
engine cycle solver described in R e f . 53, we have estimated
that at C'
T
will d r o p by -0.02 for a fuel total temperature
drop
from 1667to 1000K .Hence,for the followingdiscussion,
we will take C'
T
to be0.147 for a fuel injection total temper-
ature of 1000 K.
T he first comparison will be
between
a curved combustor
withinjectors on theouterwall, not projecting into th e stream,
an d a
benchmark
straight combustor with representative,
well-
documented injectors
with
structures protruding intothe
stream.
W e will make the assumption that e quivale nt combustion ef -
ficiencies
can be achieved using these tw o combustors , an d
will e stimate the mome ntum loss penalty due to the intrusive
injector structures
in the
straight combustor . (Detailed
C F D
analyses
of
these
flows
fo r
these
cases
would
undoubte dly
be
of great interest, but could easily comprise several separate
papers and are far beyond the
scope
of thepresent
article.)
W e consider the swept ramp geometry investigated in R efs.
24 ,
26, and 28. In
R e f .
24, tests were made
with actual
hy -
drogen fuel
combustion in a
h ot ,
vitiated ai r facility with
oxygen
r e ple nishme nt. For equivalence ratios of
1.0,
com-
bustion
efficiencies
were
estimated
to be50-60%. In R e f .
24,
the ramp deflection
angle
is
10.3
deg, an d
from
th e r a m p
dimensions given
therein, th e
fraction
of the channel blocked
by
th e ramp exits is
11.5%.
H owe ve r ,
th e
10.3-deg
faces of
th e
r amps occupy
22.3% of the
channe l
flowarea.
R e f e r e nc e
54 gives e quations pe r mitt ing
one to solve the
oblique
shock
problem
fo r
given
values of upstream Mach number, deflec-
tion
angle,and gasspecificheat ratio. For thecombustor
inlet
flow of
R e f .
1 for a flight
Mach
numbe r of 15, and a
r a m p
deflection angle of
10.3 deg,
th e
pressure
and Mach
n u m b e r
behind
t he
oblique
shock may
readily
be calculated to be
0.174
x
(stream
dynamic
pressure)
and 4.29, respectively. Assum-
ing th epressure to be
uniform
on the10.3-deg
surface
of the
ramps,
an d zero on the
rear
and side
surfaces
of the ramps,
th e
force
on the ramps can
easily
be
shown
to
correspond
to
a decrease in C'
T
of
0.0353,
a
very significant
fraction of the
total
available
C '
T
. Based on the
Mach angle
after th e
r a m p
oblique
shock,
91% of the
10.3-deg
surfaces of the
rampswill
be at the
full
pressure downstream of the shock.
Ho w ev er ,
there
will be some pressure recovery on the side an d rear
surfaces of the
ramps.
Hence, w e
m a k e
th e
rough
estimate
that th e effective
loss
of C'
T
due to pressure forces on the
ramp
will
be two-thirds of the valueestimated above or0.0236.
This is 16% of the
total
available
C '
T
,
still
a very
significant
loss.
The geometry of the c om bus tor channel analyzed a bove ,
lookingupstreamfrom aposition downstream of the injectors,
isshown in Fig. 7a. The penetrationdata fo r
swept
r a m p an d
similar injectors given in R efs.
25-28
and the relatively low
combustor
efficiencies estimated in Ref . 24suggest that th e
injector
configuration of Fig. 7a may not provide
sufficient
penetrat ion an d mixing of the fuel j e ts . If additional
r a m p
injectors
were provided, as shown in Fig. 7b, penetration,
mixing,
an d
combustor efficiency
ar e
likely
to be
significantly
enhanced. However, as a first es timate, th e effective
loss
of
C^duetopressureforceson the rampswouldnow bed o u b l ed ,
to 32% of the
total available C '
T
,
a
very large loss.
If, as
postulated,
th e
curved combustor
could
achieve e quivale nt
mixing an d combustion efficiencies ( due to buoyancy effects)
without drag forces
on injector
structures
protruding
into
th e
stream, substantial increases in
C'
T
could be achieved. (W e
note that the loss in
m o m e n t u m
calculated inSec.
III.A.
fo r
the curved combustorcorresponds to only
3.5%
of the
total
available C '
T
,
and this
could perhaps
be
fu rther reduced
by
optimization of the curved combustorshape.)
For an
assessment
of the
performance
of injection behind
pylons, we use the same ramp injector combustor geometries
discussedabove as the benchmark. The combustor inlet con-
ditions are again those from R ef. 1,
given
at the b eg in n in g of
this section. T he fuel total temperature is again taken to be
1000 K, and the fuel total pressure is taken
from
R e f . 1 a s
4.41 x 10
7
dyne/cm
2
(for a fuel equivalenceratio of1.0). The
pylongeometry
considered
is shown in
Fig.
8. The height of
th e pylon is taken to be equal to that of the ramp injector,
and it is considered to stand in the same combustor channe l
used
with
the ramp injector. The
pylon
and the
injection
channel are raked back at 30 deg to the
flow
direction, and
th e
fuel
is fully expanded to the combustor static pressure
upstream of the pylons before injection.
This corresponds
to
injection at Mach3.51.For the
fuel
equivalence
ratio
of1.0,
th earea of the
fully expanded fuel
jet can
easily
be
calculated
to be 0.0605
times
th e
combustor channel
area. In Fig. 8, we
have rather arbitrarily taken the height of the fuel
channe l
(perpendicular to the
fuel
flow) to be equal to the height of
th e
pylon,
an d
havetaken
th e
half-angle sweep
of the pylon
(viewed normal to the end of the pylon) as equal to the ramp
deflection
angle of the
r a m p
injectors . W e assume that the
same mixing
an d
combustion
efficiencies can be
obtained
us-
in g the pylon injection technique or ramp
injectors.
The mainadvantageof the pylon
technique
is that the pylon
fo rward
dragsurfaces
only
obstruct
0.0605
of thechanne l
area
vs 0.2229
of the
channe larea
for the
10.3-deg deflection sur-
face of the
r a m p
injectors. The
es timated pressure
on the
forward drag
surfaces of the pylons is
essentially e qua l
to that
on the
10.3-deg surface
of the
ramps. There will
be some
reduction in the average pressure on the
fo rward
surfaces
of
the pylons due to expansion
waves
emanating from the top
of
the pylons. To
estimate
th e
reduction
in
C'
T
due to
pressure
dragforces
on the pylons , we mak e the
sameassumption
used
to
estimate this reduction
for the
r a m p
injectors.
This
is, the
effective pressure difference
between the forward an d
rear
surfaces of the
pylon
istaken to beequal to two-thirds of the
pressure
calculated for the forward
surfaces with
no
allowance
m a de for the
relieving
effects of
expansion waves .
T he
result
of this
calculation
isth a t , for injectors on one side only of the
c om bus tor channel, the reduction in
C'
T
due to pressure drag
J 3
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BOGDANOFF:
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FOR
SCRAMJET COMBUSTORS
189
on the injectors is 4.4% of the
available
C'
T
for the pylon
injection technique vs 16% for the r a m p injectors, a very
significant
improvement. If, as discussed previously for the
curvedcombustorchann el, it isrequiredt odoublet he n u m b e r
of
injectors
to
assure good
mixing and
combustion
efficiency,
the se numbe r s would increase
to 8.8 and
32%,
and the
dif-
ference
between the
r a m p
an d
pylon
injection
technique
be-
comes even more significant.
In the above comparison of the ramp and pylon
injection
techniques,
we have
ignored
th e
question
of
differences
be-
tween
th e
degree
of
recovery
of fuel je t
m o m e n t u m
for the
tw o
techniques. The angle of injection for the
r a m p
injectors
(10.3 deg) is m or e favorable than that for pylon injection (30
deg). O n the
other h a n d ,
th e
fuel
je t
expansion
is
more com-
plete
(t o
Mach
3.51)
for pylon
injection
than for r a m p
injec-
tion
(t o Mach
1.7).
The latter favors pylon injection. The
overall effect
is to
slightly favor pylon
injection. It may be
argued
that the
r a m p
injector
fuel
channels could be
altered
to
provide
expansion of the fuel to Mach
3.51
a nd still to have
th e
favorable
r a m p injector injection angle. On the other
hand, it may also be argued that th e
pylon
injectors an d fuel
channelscould perhaps
b e
raked downstream
at a angle smaller
than
30 deg to the
flow direction,
or
that
curved pylons
(e.g.,
see Fig. 5b) could be used to allow improved recovery of the
fuel
j e t m o m e n t u m .
From
th e
current discussion,
it
appears
that th e degree of
fuel
je t
m o m e n t u m
recovery is roughly
equal
for the
r a m p
and pylon injection
techniques .
Hence, in
the preceding paragraph, this rough equality was implicitly
assume d.
W e now make anassessment of the
performance gains
which
might be achieved usingthe pulsating
injector
technique. The
performance gains are assumed to be m a de because
better
mixing and higher combustion efficiencies would be obtained
with
pulsating injectors. As
mentioned earl ier,
in the
exper-
imental
investigation
of Ref . 24,
using
r a m pinjectors at equiv-
alence ratios of
1.0,
combustion
efficiencies
were
estimated
to be50-60%. Higher combustion efficiencies would be de-
sirable. Using the
scram
je t
engine cycle solverdescribed
in
R e f . 53, we
have calculated
that fo r flight at Mach 15 and a
fuel equivalence
ratio
of 1.0,C
T
increases
by 0 . 0 3 fo reach
20 % increase in combustion efficiency. Thus, if the use of a
pulsating injector
ca n increaset he combustion efficiency from
60% to,
say,80%, C'
T
would be
increased
by
0.03.
Starting
from our baseline C'
T
value of
0.147,
this would represent a
20 %
increase
in
C '
T
,
a very significant
i m p r ove m e nt .
T he
following
important
point
must be
made regarding
th e
preceding discussion. It is not intended above to in any way
imply
that the new injector/mixing concepts necessarily will
exhibit better
performance
than
a
straight c om bus tor with
r amp
i n j e c to r s t h e latterinjectors
are proven and are
k n o w n
to be
quite effective. Furthermore,
th e flows
being
compared
ar everyc o m p l e x t h eyare three-dimensional, unsteady, tur-
b u len t ,a n d
contain
separatedflowr e g i o n s h e n ce ,t he actual
performance of the new
injector/mixing concepts (e.g.,
re -
garding mixing
an d
combustion
efficiency)
ar e very
difficult
to predict. The new injector/mixing concepts look
sufficiently
good when compared to a
benchmark
ramp
injector
com-
Three new advanced mixing techniques were presented.
The first was a
c om bus tor ,
curved so
that
buoyancy
forces
will aid in the
penetrat ion
of the
fuel
across th e
combustor.
Th e effect of two
fuel
plume geometries was analyzed. The
curvature
an dc om bus tor
length
required for the
fuel
to cross
the combustor were
assessed.
The
second
was pulsation of
th efuel injectors
to increase
penetration
an d
mixing.
A
fluidic
technique,
a
modified Hartmann-Sprenger tube,
w as
identi-
fied
as a strong candidate to generate the pulsations. The
Hartmann-Sprenger
tube must be
optimized
to produce strong
pulsations without producing excessiveheat
transfer.
The third
w as th e injection behind pylons to allow deep penetration
into
th e
airstream. This technique
is likely to
produce high
base pressures on the injector structure. Curved or
slanted
pylons
can be used to increase the recovery of
fuel
jet mo-
m e n t u m .Tapered
pylons
can be used to optimize fuel distri-
bution.
Control of the fuel jet can be improved by partially
trapping it in the curved
rear
of the pylon. It may also be
possible to
increase
th e
base pressure
on the pylon by delib-
erately diverting a fraction of the
fuel
fo r base bur ning on the
pylon.
A
preliminary assessment
of the potential of the three ne w
mixing
techniques to
increase
scramjet engine performance
has been made. Combustors
using
the new techniques were
compared
with
a benchmark c om bus tor
with
swept ramp in -
jectors. In these
assessments,
the new
mixing
techniques
looked
sufficiently
good compared
to the benchmark
case
to warrant
further
investigation.
Acknowledgments
This work w assupportedby the
Eloret
InstituteunderGrant
NCC2-487.
The
CFD
solution for the
curved
combustor was
compute d
by S. Polsky of the
Eloret Inst itute.
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