scramjet combustors

7/27/2019 scramjet combustors

1/8

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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3/8

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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4/8

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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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


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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7/8

BOGDANOFF:

TECHNIQUES

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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