-------
of
o
a,
w
H
2
700
600
500
400
300
w 200
100
UPSTREAM OF PACKING (AVG. OF 4)
AT EXIT (SEE FIG. 6)
012
9 10 11
CO
E-I
Oi
<
Oi
H
I—!
m
1.0
0.5
0.2 0.4 0.6 0.8
RADIAL FRACTION FROM CENTER
1.0
Figure 22 . Test 3, gas temperature in entrainment separator
from thermocouple arrays.
Figure 23. Test 3, velocity profile in
scrubber.
-------
high a temperature, the fiberglass structure housing the de-
mister was not damaged. As with previous tests, the only
damage within the system was to the polyethylene Tellerette
packing and the fiberglass packing supports. These were melted
slightly along the top of the demister. It was evident that
the added sprays had no noticeable effect on the afterburning.
Velocity Profile at End of Scrubber Duct-
The total pressure relative to atmospheric at the end of
the scrubber duct is shown in Figure 23. The profile is simi-
lar to that of the second test.
Performance of Liquid Injectors -
As in the previous test, the first angle iron shield for
the liquid injection pipe burned through. The burned-out gap
was narrower than the previous, only about 5 cm (2 in). The
pipe itself was still intact, as were the other two angle iron
and pipe injectors. More holes should be drilled in the first
pipe, near the flow centerline, to try to get more heat trans-
fer away from the center of the angle iron.
Test No. 4 - November 19, 1976
The final test in the series duplicated the conditions
and results of the third test. One difference was an attempt
to grab gas samples through tubes inserted into the entrain-
ment separator section, just after the packing. Also, three
thermocouples were inserted inside the entrainment separator
section to monitor temperatures inside the section. Previous
downstream temperatures were measured at the exit.
Pressure and Temperature Data-
The pressure transducer and thermocouple data are shown
in Figure 24-26. It is important to note that the temperature
at the exit (thermocouples 1 and 2) reached 650-850°C while
those inside, just after the packing (thermocouples 6, 7, and
53
-------
m
o
1
w
«
o.
CO
01 2 34 5 6 7 8 9 10 11 12 13 14 15
Figure 24 . Test 4, rocket chamber pressure
100
1- RIGHT SIDE, MID-DUCT
2- LEFT SIDE, MID-DUCT
01 2 3 4 5 6 7 8 9 10 11 12 13 14 15
TIME, s
Figure 25. Test 4, scrubber duct wall temperature.
54
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Ln
Cn
UPSTREAM OF PACKING
(AVERAGE OF 5)
DOWNSTREAM OF PACKING:
1,2,3- AT EXIT (SEE FIG. 6)
6,7,8- SEE FIG.6
0 1 2 3 4 5.6 7 89 10 11 12 13 14 IS
7 8 9 10 11 12 13 14 15
0.01
012345
Figure 26. Test 4, gas temperature in entrainment separator from
thermocouple arrays.
Figure 27. Test 4, pressure near rocket nozzle exit (diffuser
-------
8), reached only 200-300°C. Contact with outside air seemed
important to the temperature producing mechanism, which must
surely be afterburning of hydrogen and carbon monoxide.
A new figure (27) is shown to illustrate the pressure
felt at the inlet to the scrubber (or rocket nozzle exit). A
differential transducer was used. Start-up and shut-down
transients were quite apparent and could have an effect on
thrust, depending on the area experiencing these pressures.
Only one transducer was used in the test so that spatial vari-
ations in pressure were not measured. The data indicate posi-
tive pressure peaks at shut-down, which mean that short dura-
tion leakage was occurring. These leaks were seen in films
of the tests but probably represented a very small amount-
Velocity Profile at End of Scrubber Duct -
The total pressure relative to atmospheric across the
end of the scrubber duct is shown in Figure 28. The two cen-
ter pressure taps were overpressurized to the extent that the
manometer fluid was blown out. This indicated poor mixing
in the duct which had not been experienced to such an extent
in previous tests. A possible reason was that the rocket
chamber pressure reached a higher level in this than in pre-
vious tests, causing a higher nozzle velocity.
56
-------
0.5
-1.0
(Fluid blown out of mano-
meters at 0 and 0.1)
0
0.2 0.4 0.6 0.8 1.0
RADIAL FRACTION FROM CENTER
Figure 28. Test 4, velocity profile in
scrubber.
57
-------
CONCLUSIONS FROM AFRPL TEST PROGRAM
Much was learned about the basic design and practical as-
pects of the AFRPL gas-atomized scrubber. Most important, it
worked and was not destroyed by the solid rocket exhaust.
However, only tests of 10-second duration were made and there
are serious doubts that the entrainment separator could with-
stand longer durations mainly because of afterburning effects.
Also, the solid rocket motors tested had end burning propellant
grains which produce more gradual thrust start and tail-off
transients. Whether the scrubber could have withstood some of
the rapid starting motors is questionable.
Scrubbing Efficiency
The aluminum oxide particles were collected with greater
than 99% efficiency in the scrubber based on filter sampling and
the gas exiting the scrubber had no discernible opacity. Since
the gas sampling system was unsuccessful, the HC1 removal effi-
ciency can only be inferred from other observations. Chief
among the other observations were the particulate collection ex-
periments and visual opacity. These observations indicate that
there was good contact between the rocket exhaust and the injec-
ted liquid which would also mean that good conditions for HC1
absorption also existed.
Further reason to assume good HC1 absorption was that the
temperature and pressure measurements showed that mixing of the
exhaust with the injected liquid was good. Theory of HC1 ab-
sorption, which will be discussed later, indicates that the
absorption will occur in a few meters provided the drops are tho-
roughly mixed with the gas.
Liquid Injection System
The angle-iron protected pipe injectors performed their
function but could be improved. The first angle burned
through during each test, which required that time be spent
replacing it. Some of this maintenance time could be saved
if a more heat resistant angle was used. Possibly a stain-
less steel angle, coated with an ablative or a ceramic, could
be used.
58
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The amount of liquid used (85-115 £/s) was adequate, as
theory predicted. However, not enough tests were conducted
to determine the optimum rate. Also, the 3$ by weight KOH solu-
tion seemed adequate, but may not be optimum. Because of the
complexity that use of the KOH solution required it was ques-
tionable for an operational scrubber. As will be discussed later,
the use of a basic solution is not essential for the mass trans-
fer of the HC1 into the scrubbing liquid. The neutralization
could perhaps be more easily accomplished in the sump and the
drying pond.
Diffuser
The "diffuser" served to direct the exhaust gases into the
scrubber and to protect the front part of the scrubber from
radiation heat damage. It suffered heat damage during each test.
With the present design of the system this was unavoidable since
the spray ring did not supply nearly enough water to keep
the walls cool.
Entrainment Separator
The entrainment separator was very efficient, as evi-
denced by the low emission rate of aluminum oxide particulates.
The particles were collected on drops in the scrubber with an
efficiency greater than 99.9% so that the entrainment separator
efficiency should be the same as the overall particulate ;
efficiency based on the filter sampling. The start-up pressure
transient was effectively relieved by the hinged hatches so that
overpressurization was not experienced. Since the entrainment
separator was so efficient it is suggested that the packing could
be reduced in thickness somewhat. The reduction in packing
thickness would partially alleviate the pressurization of the
entrainment separator section.
Afterburning
A major finding of the solid rocket test program was the
occurrance of afterburning. The high temperatures sensed by
the exit thermocouples and the damage done to the sampling
59
-------
apparatus are clear evidence that the carbon monoxide and hydro-
gen exhaust gases were afterburning. This afterburning had some
affect on the entrainment separator section. The packing and
packing supports were slightly melted and charred. While the
afterburning itself may not have been preventable its conse-
quences could have been reduced by a different design. There
was no basic design reason for the converging section after the
packing. This convergent section was only there for the pur-
pose of providing a smaller and more uniform section for sampling.
It provided a region where air could collect, circulate and re-
act with the combustible gases in close proximity to plastic
and fiberglass surfaces. Had there been no convergent section
any afterburning would have taken place at some distance above
the structure, causing much less damage.
In order to minimize afterburning effects the gas di-
rection leaving the entrainment separator should be vertical
rather than horizontal. Also, the exiting gases should have
as high a velocity as possible so that any afterburning will
occur as far above the packing and structure as possible.
Finally, the separator should be designed so that when the
rocket is ignited all the air is forced out with the first
blast of exhaust gas. There should not be any pockets of
air left while the system is operating.
Sampling System
Sampling of the effluent gas was difficult because the
afterburning melted the sampling lines and filters, and the
velocity was not uniform. The array of filters for particu-
late sampling worked for the first test but holes were burned
in the filters during the second test. Sampling with filters
was abandoned after the second test. Gas sampling with evac-
uated bottles was unsuccessful because the sampling lines
burned through. The afterburning, in effect, has made gas
sampling practically impossible in the present type of system.
60
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Sampling of the effluent liquid and sump could be done in
a thorough fashion. A number of instruments could have been
used to monitor the flow rate, turbidity, pH, and chloride
ion concentration to provide a better indication of the scrub-
bing efficiency than was accomplished with the gas sampling
system. Another possibility would be to collect all the ef-
fluent liquid, including the post-test rinse of the entrain-
ment separator packing, in a tank for post-test analysis. For
the 10-second tests an 800 liter (210 gallon) tank would be
adequate.
Coupling Effects
The effect of the scrubber on the rocket test was not
measured directly because thrust was not measured. One pres-
sure transducer located beside the rocket nozzle exit did
indicate ignition and burn-out transient peaks which could
have affected thrust. The peaks were short duration so that
their effect should be small. An array of transducers would
be needed to determine the pressure distribution on the rocket
nozzle in order to estimate thrust effects.
61
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SECTION 5
AIR POLLUTION CONTROL EQUIPMENT ALTERNATIVES
INTRODUCTION
The purpose of this section is to review alternative air
pollution control devices and present background information
on equipment which has been used in similar applications. The
most suitable approaches for the present application will then
be presented.
SELECTION OF AIR POLLUTION CONTROL EQUIPMENT
Many types of equipment can be used for controlling emis-
sions from stationary sources. These types include filters,
electrical precipitators, cyclones, mechanical collectors,
scrubbers, adsorbers, and combustors. In the case of the
rocket exhaust the source is a supersonic stream of extremely
hot gas. The pollutant of main concern is hydrogen chloride
gas. Thus, the type of control equipment must be capable of
collecting a very hot, very corrosive gas. Filters, electri-
cal precipitators, cyclones, and mechanical collectors are
primarily designed to collect particulate matter. Combustors
are used mainly for oxidizing gaseous contaminants to non-
toxic gases such as water and carbon dioxide. Scrubbers and
adsorbers are the two types of equipment used most often for
removing gas phase contaminants. Adsorbers retain contaminant
gases on the surface of porous particles around which the car-
rier gas flows. Scrubbers introduce liquid into the collector
to dissolve or react with the contaminant gas.
Adsorbers are not suited for the control of rocket ex-
haust for many reasons. Adsorption usually works best for
adsorbates which are dilute, dry, and cool. In addition,
particles of the adsorbent must be used to provide adequate
surface area., creating the task of removing particulates which
may be toxic. The only major industrial use of adsorption for
62
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removal of a gas similar to hydrogen chloride is the closed
loop system for recycling hydrogen fluoride from aluminum
smelting fumes (Cochran, 1974). The mechanism in this use is
chemisorption on calcined alumina (A1203). Injection-type
dry sorption using limestone to remove S02 from power plant
boilers has been extensively tried, but found to be ineffi-
cient (Slack and Hollinden, 1975). Some adsorption of HC1
will occur on the alumina particles generated by the rocket
motor but this does not appear to accomplish more than a
small fraction of the required transfer.
Liquid scrubbing processes are more attractive for re-
moving rocket exhaust gases because water must be used to
cool the gas and the mass transfer is good at the tempera-
tures involved. Like the dry scrubbing process, if the
scrubbing medium is sprayed into the gas stream, provision
must be made to remove the added matter (drops).
Various types of liquid (wet) scrubbers are available
for mass transfer. The types are plate, massive packing,
.fibrous packing, preformed spray, gas atomized spray, cen-
trifugal, impingement and entrainment, moving bed, and com-
binations (Calvert et al., 1972).
A plate scrubber consists of a vertical tower with plates
mounted transversely inside. Gas enters at the bottom of the
tower and must pass through perforations, valves, slots, or
other openings in each plate before exiting the top. Liquid
is usually introduced at the top plate and flows successively
across each plate as it moves downward to the liquid exit at
the bottom.
Massive packing scrubbers consist of towers containing
manufactured or natural elements. Liquid is usually intro-
duced at the top and trickles down through the packing. The
gas stream should not be too heavily loaded with particles
or the packing will become clogged. Fibrous packing scrubbers
are similar in principle to massive packing scrubbers except
that fiber beds with very large void fractions are used.
63
-------
A pre-formed spray scrubber collects particles or gases
on liquid droplets and uses spray nozzles for liquid droplet
atomization. The sprays are directed into a chamber suitably
shaped to conduct the gas through the atomized liquid drop-
lets. Centrifugal scrubbers with spray manifolds are a type
of pre-formed spray scrubber that impart a spinning motion to
the gas passing through them. This configuration reduces
droplet carryover due to entrainment because the droplets are
impacted upon the scrubber walls by centrifugal force.
Gas-atomized spray devices use a moving gas stream to
atomize liquid into drops, and then accelerate the drops.
High gas velocities (60-120 m/s) are used to promote particle
collection and finely atomize the liquid which is introduced.
Entrainment separators must usually be used.
Impingement and entrainment scrubbers are configured so
that the entering gas must pass over a reservoir of liquid
at a speed and direction which causes the gas to atomize and
entrain the liquid. These devices usually have an entrain-
ment separator built into the exit duct.
Moving bed scrubbers are like the packing scrubbers ex-
cept the packing is usually spheres and these spheres move
around during operation. Gas velocities are high to make
the bed turbulent enough to keep the packing clean. Thus,
.this type is suitable for particulate as well as gas removal.
The efficiencies of the various scrubbers all depend on
a number of factors and each can be designed to any desired
efficiency. The primary factors affecting efficiency are
liquid-gas contact surface area and contact time. Contact
types of packing, plates, or spray atomizers. Contact time
is regulated by packing height, number of plates, and spray
chamber length.
EQUIPMENT FOR REMOVAL OF HC1 and HF
Hydrogen fluoride is encountered more often than HC1 as
a serious industrial air pollutant in the stack gases from
64
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phosphate fertilizer plants, aluminum plants, and calcium
phosphate furnaces. The high absorptivity of both HC1 and
HF in water has made scrubbing the most widely used removal
means. Kohl and Riesenfeld (1960) review several scrubbers
used for HC1 tail gas removal. Most use a packed tower with
water as the absorbent. Scrubbers for HF removal have been
reviewed by Magill et al. (1956), Kohl and Riesenfeld (1960),
and Teller (1967). They describe a number of spray, packed
tower, and venturi scrubbers that have been used. Magill
also describes HF removal by passage through beds of lump
limestone to produce calcium fluoride in fine particulate
form.
In the recent literature Kempner et al. (1970) tested
several packed, plate, and spray tower hydrogen chloride
scrubbers. Tomany (1969) describes a moving bed scrubber
used on an aluminum processing plant. Rust et al. (1973)
also describe a number of scrubbers for use on aluminum
smelters.
In recent years hydrogen fluoride has been recovered
during aluminum smelting by chemisorption on calcined alum-
inas .(Cook et al., 1971). Since solid propellant rockets
produce alumina (A1203) this may be a possible removal mech-
anism for both HF and HC1. The problems with this process
for rocket exhaust scrubbing are that the reaction occurs
at low temperatures and aluminum fluoride dust is produced
(Cochran, 1974). Although at present the technology is not
directly applicable to the rocket exhaust, further study and
development of the process could make it attractive.
SCRUBBERS USED ON HIGH ENERGY EXHAUSTS
Aside from rockets the major producer of high tempera-
ture, high speed exhaust gas*is the jet engine. While the
jet engine does not produce much toxic gas the scrubbing of
its particulate emissions involves similar means. In the
following a number of scrubbing facilities will be described
65
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which are applications similar to the solid rocket exhaust.
While a number of such scrubbers are known to have existed,
only those for which a reasonable amount of information was
obtained will be discussed.
Naval Air Rework Facility
Jacksonville, Florida
This facility operates a jet engine test cell that has
a pollution abatement system for particulate removal. The
control system consists of a quencher and a cross flow packed
scrubber designed by Teller Environmental Systems, Inc. The
entrainment separation was accomplished by an additional sec-
tion of dry packing. The system is sized so that the super-
ficial velocity is in the range of 2.5-5 m/s. According to
tests performed in a similar model system, considerable par-
ticulate removal occurs in the quencher section as the gas
is cooled. The conditions and efficiencies for two jet en-
gines at military power were as follows (Kemen, 1976):
J52 Engine
Thrust 37,000 N (8,300 Ibf)
Volumetric Flow Rate 154 m3/s at 15°C
Exhaust Temperature ~800°C
Heat Release 38 MJ/s
J79 Engine
Thrust 49,000 N (11,000 Ibf)
Volumetric Flow Rate 135 m3/s at 15°C
Exhaust Temperature ~800°C
Heat Release 47 MJ/s
66
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Particle Scrubbing Efficiency
Engine J52 J79
Quench, QL/QG, £/m3 at 15°C 0.29 0.33
Gas Velocity in Scrubber, m/s 3.6 to 4.1
Particle Loading, mg/m3 at 15°C 23 55
Efficiency, % 74 81
Particle size data were not available.
Arnold Engineering Development Center
Arnold Air Force Station, Tennessee
The AEDC is a complex of wind tunnels, propulsion test
cells, aerospace chambers, and hyperballistic ranges. Their
Engine Test Facility has a number of test cells for testing
rockets at simulated altitudes and Mach numbers. The high
altitude (low pressure) is accomplished primarily by steam
ejector-diffusers during prefire and by the rocket exhaust
gas ejector action during the firing. High velocities are
obtained through a combination of ejector-diffusers, air
supply compressors, and exhaust compressors. Two of the
larger test cells (J-4 and J-5) use spray chambers to cool
and dehumidify the exhaust gases. In cooling the exhaust
gases these cells also effectively scrub particulates and
soluble gases. Performance of these test cells as scrubbers
has not been reported, however.
The test cells at AEDC are not designed for low altitude
(sea level pressure) testing. Low altitude testing requires
that the cells be maintained at local atmospheric pressure.
This could be accomplished by opening up the cells. However,
the increased gas density at the higher pressure would put
too high a load on the exhaust system. Either the compressors
would overload or the ducting system would be overpressurized.
The problem would be particularly severe at rocket ignition
when the large mass of air resident in the system has to be
moved at a very high rate as the rocket plume enters the ex-
haust system.
7 67
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Toxic Attitude Propulsion Research Facility
Air Force Rocket Propulsion Laboratory
Edwards Air Force Base, California
The TAPR facility is designed to allow safe ground-
testing of toxic propellants in simulated high altitude envir-
onments. Steam ejector-diffusers are used to pump the system
to the desired pressure. Spray chambers to cool the rocket
exhaust are also incorporated in the facility design. The
facility is described in Aviation Week (1967).
As with the AEDC test cells, the spray chamber acts as a
scrubber for particulates and soluble gases. It was estimated
to be 95% efficient on soluble effluents but some unpublished
data showed that it may be less efficient. A thorough study
of its scrubbing efficiency has not been published.
The facility, like those at AEDC, is not designed for low
altitude (high pressure) testing. The problems that would be
encountered during atmospheric pressure testing have been
briefly described previously.
Jet Propulsion Laboratory - Edwards Test Station
Edwards Air Force Base, California
JPL-ETS operates a rocket engine toxic exhaust scrubber
facility described by Frank C. Brown (1969). The scrubber was
designed to operate on a 1,000-second duration, 9,000 N (2,000
Ibf) thrust liquid rocket. The oxidizer and fuel were OF2 and
B2H6, and the reaction products were HF, EOF, and H20.
The scrubber uses a two-step alkaline water solution
starting at the exit of a diffuser which ducts the rocket ex-
haust gases from the rocket nozzle to the spray scrubber.
Sodium hydroxide is used in the first step during rocket oper-
ation, and calcium hydroxide is used in the second step either
during or after rocket operation. During the second step the
sodium hydroxide is reconstituted and fluorinated and boron-
ated calcium compounds are precipitated out and later removed
physically.
68
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The gas-liquid contact occurs in a 0.91 m diameter by
9.1 m long horizontal, co-current spray duct which exhausts
downward through an elbow into the sump. The gas must then
turn upward to flow through a vertical packed tower. The
tower has a wet stage for more scrubbing and a dry stage for
entrainment separation. The liquid flow rate in the spray
chamber is 151 £/s and in the tower it is 63 2,/s. An axial
fan is used at the tower exit to maintain a negative system
pressure.
The system is designed to reduce the concentration of
fluorine or boron in the effluent to 3 ppm or less. Con-
sidering that almost all the combustion products need to be
removed this represents a very high efficiency. However,
such a high efficiency is probably not unrealistic since the
flow rate of scrubbing liquid was about 90 times the mass
flow rate of the rocket propellant. This compares to the
10-15 factor used for the design of the AFRPL pilot scrubber.
Other Facilities
A few other rocket exhaust scrubbing facilities have been
briefly described in the Phase I report (Calvert and Stalberg,
1975). They represent efforts by several contractors to re-
duce the emission of primarily hydrogen fluoride gas or beryl-
lium oxide particles. Most designs use the nozzle spray or
gas-atomized spray design quencher/scrubber followed by an
entrainment separator. As usual, reports on the efficiency
of these scrubbers have not been published.
POTENTIALLY SUITABLE SCRUBBERS
Because of the requirement for gas absorption a wet
scrubber is the type of equipment that should be used. No
other type of device has been used so extensively and suc-
cessfully for gas removal. Also, scrubbers that are highly
efficient for gas absorption are usually highly efficient
for particle collection. All the types of scrubbers can be
69
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designed to meet the required efficiency. Thus, selection of
one particular type will depend primarily on economic factors.
Most designs for rocket exhausts, or similar processes, have
been a spray chamber followed by a packing for extra scrubbing
and entrainment separation (mist elimination). This spray
scrubber design is probably the least expensive approach since
all the designs will require a quencher and the quencher can
be an integral part of the spray scrubber. The economic
selection of the scrubber will be detailed in the next section.
70
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SECTION 6
DETAILED DESIGN
DEFINITION OF THE SCRUBBING PROCESS
The scrubbing process required for the solid rocket exhaust
is primarily high efficiency gas absorption. Particle collection
will also be accomplished at an efficiency which depends on the
type of scrubber chosen. While the particulate pollutant emissions
in question were not considered a problem of concern in the desert
air basin portion of San Bernadino County, there might be a signi-
ficant restriction in the case of other pollutants and/or other
jurisdictions. Therefore, attention is given to both gas absorp-
tion and particle collection in this section.
The problems associated with the very high energy, high vol-
ume flow rate, and short duration of the rocket exhaust distin-
guish this absorber from those found throughout the process indus-
try. Conventional configurations can be utilized but the pecu-
liarities of "rocket scrubbing" lead one to consider the possi-
bilities for unconventional approaches.
Simplified Process Flow Sheet
The general nature of the control process for the rocket
exhaust is shown in Figure 29. The four main components shown
are:
1. Quencher
Here the rocket exhaust is cooled and slowed by massive
amounts of injected water. The gas leaving the quencher
is saturated with water vapor and the velocity is low
enough to be more amenable to particle and gas scrubbing.
Some pollutant collection will take place in the quencher.
2. Sc_rub_b_er
In this section the gaseous pollutants are absorbed and
71
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Rocket
Exhaust
Air
Quenching
Water
Quencher
Neutralizing
Wash
Exhaust to
Atmosphere
^
I
7
Scrubber
Gas
Liquid
rv
\_
L
^V.
i
i
Entrainment
Separator
^
7
/
Drying Pond
Figure 29. Schematic of rocket exhaust gas scrubbing process
-------
particulate pollutants are collected. In some scrubbers,
such as the gas-atomized design, this section could be
integral with the quencher.
3. Entrainme'nt Separator
Liquid drop carry-over from the scrubber is collected
here. If a neutralizing agent is used in the scrubber
this section is not absolutely necessary under the local
APCD rules (see Section 3), but several other factors
could justify its use.
4. Drying Pond
The scrubber and entrainment separator liquid effluent
should generally be collected in a drying pond rather
than be allowed to pollute the ground water.
Quenching Equilibrium
Basic to the design of a scrubber for solid rocket exhausts
is the calculation of flow conditions at the inlet to the scrub-
ber. The first and simplest calculations are the equilbrium
balances of mass, momentum, and energy in the quencher. The
required liquid injection system and mixing length in order to
achieve these equilibria will be discussed later.
As will also be discussed later, two scrubber designs,
which have different types of quenchers, will be proposed. The
equilibrium conditions presented here will be appropriate for each
of the types of quenchers.
Ideal Gas Law -
The ideal gas law is adequate to define P.V.T. relationships
for the gases of the rocket exhaust, air, and water vapor at near
atmospheric pressure. With the subscripts "v" and "g" represen-
ting water vapor and non-condensing gases, respectively:
n R T
P = v u (11)
v Q
(12)
73
-------
P • Pv + pg (13)
where P = pressure, N/m2 (Pa)
n = mole flow rate, kgmol/s
Ru = universal gas constant, 8,314 J/kgmol-°:K
T = temperature, °K
Q = volume flow rate, m3/s
Since "P" is known or can be easily approximated and "n " is
o
known from the composition of the rocket exhaust and the amount
of entrained air these equations are rearranged to give:
n R T
Q - -f-f- (14)
V ~ FV
P n
"v '
"Pv" can be found from the vapor pressure relationship for
water.
Vapor Pressure of Water Solutions -
To assure quenching, enough water is assumed to be used
to saturate the gas at the equilibrium pressure and temperature.
The following equations are based on the assumption that some
liquid water is present.
The vapor pressure of pure water is related to temperature
in accordance with the following equation:
log10(PVjpure) = a - ^— (16)
where a = 5.84191
b = 1668.21
c = -45.
for 333° < T < 433°K
74
-------
T = - - - - c (17)
a - Iog10 (Py, pure )
Dissolved species cause a depression o£ the vapor pressure
At a given temperature the depression may be represented as:
(18)
P
v
where d and c2 = constants
X = concentration of the salt solution
Thus:
T = - b - . c (19)
a - loglo (Pv d')
Mass Equilibrium -
The total mass flow rate fm) must equal the mass flow rates
of the rocket exhaust, m , the entrained air, rn , and the in-
' r d
jected liquid, rn^.
m = mr (1 + Ra + Rw) (20)
where m = mass flow rate, kg/s
R = mass ratio, kg/kg
The volume flow rate of gases is:
Q = u A (21)
where u = velocity, m/s
A = cross sectional area of mixing section, m
2
75
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The volume of liquid and solid particles is neglected. The
velocity is then:
u = 2. (22}
A
Momentum Equilibrium -
The momentum balance requires that the vector sum of the
forces acting on a control surface equals the vector change in
momentum within the volume enclosed by the control surface.
For the case of an open quencher, with all pressures being
equal, the only external force is the resistance force due to
obstacles and wall friction. Thus:
F = FD + m u (open) (23)
where F^ = drag (resistance) force, N
F = rocket thrust, N
The flow was assumed to be one dimensional and the injected
liquid was assumed to contribute no momentum in the flow direc-
tion. The velocity of the air was also assumed to be negli-
gible. For the closed, ducted quencher pressure forces are
present. In this case the resistance forces are usually small
in comparison with the other forces and can be neglected. So:
F + Pin (A-Ae) = m u + P A (closed) (24)
where Pin = pressure at quencher inlet, N/m2
Ae = rocket nozzle exit area, m2
Solving for "P. ":
6 in
p = m u + P A - F
in A - A
76
-------
"P. " is the pressure at the quencher inlet due to the ejector
action of the rocket. It is not as well defined as "P". The
equilibrium pressure, P, is the pressure above atmospheric by the
amount equal to the scrubber/entrainment separator pressure drop.
Composition and Chemical Reaction Balances -
The gas composition may change in the quencher because of
reactions of hydrogen and carbon monoxide with oxygen in the
entrained air and reactions of the hydrogen chloride or hydro-
gen fluoride with bases in the injected liquid (if any).
CO and Hz react with the oxygen in the entrained air:
CO + H2 + 02 -» C02 + H20
Because of the assumed absence of catalysts for the reverse
reactions (e.g., the "water-gas" reaction), the amount of CO
and H2 can only reduce.
The oxygen is assumed to react completely. Thus:
n(02) = n(02)a - % n(CO)r + n(H2)r
(26)
n(CO) = n(CO)r - n(02) (27)
n(H2) = n(CO)r + n(H2)r + 2 n(02) -n(02)a - n(CO) (28)
The notation "n(species) is the molal flow rate of the species
in parentheses.
Energy Balance -
The static enthalpy of the rocket exhaust products is:
i (@Tr) - hi(@298°K) + Ah£i(@298°K)
(29)
where "i" refers to the various components of the rocket exhaust,
and:
h = enthalpy per mole, J/kgmol
Ah_p = enthalpy of formation per mole, J/kgmol
77
-------
The enthalpy of the air is assumed to be zero since the temper-
ature is close to 298°K. The enthalpy of the injected liquid is
assumed to be the enthalpy of liquid water at 298°K-
The final equilibrium enthalpy is just:
? n- h.(@T)
j J J
where "j" represents all species present at equilibrium.
The energy balance must account for the change in kinetic
energy as well as the change in enthalpy, so:
n h (@T ) + ng hg (@298°K)+%m u 2 = Z n, 'h. (@T)+%m u2 (30)
rrr >o x/ j.1 -jjj
Solution of Equilibrium Equations -
A Fortran computer program has been written to provide a
solution to the equilibrium equations and is presented in the
appendix "A".
Comparison with Garrett
As a check on the equilibrium predictions made by the com-
puter program described above, a comparison was made with
a previous prediction by Garrett et al. (1972). The prediction
is for a 22 kN (5,000 Ibf) solid rocket exhausting into a scrub-
ber duct that is 0.914 m (3 ft) in diameter, similar to the
AFRPL pilot scrubber. The parameters used in the prediction
are presented in Table 9. The velocity and temperature pre-
dictions are shown in Figures 30 and 31. The two velocity predic-
tions are very close and the temperature predictions are dif-
ferent by only 2 to 5 degrees Celsius. The slight temperature
difference may be due to different methods of calculating the
equilibrium energy balance.
The present prediction program has several features that
are not present in Garrett"s program. Among these are a more
precise method of calculating the equilibrium energy balance,
provision for heat release to form the salt, and provision for
vapor pressure depression due to dissolved salts.
78
-------
TABLE 9. PARAMETERS FOR COMPARISON WITH GARRETT (1972)
ROCKET EXHAUST COMPOSITION AND STATIC ENTHALPY
Component Mole Fraction Enthalpy Relative to 298°K
kj/gmol kcal/gmol
CO
. C02
H20
HC1
A1203(S)
N2
H2
0.
0.
0.
0.
0.
0.
0.
2047
0212
1652
1688
0756
0841
2531
73.
118.
94.
70.
230.
72.
68.
22
25
39
53
90
46
38
17.
28.
22.
16.
55.
17.
16.
50
26
56
86
18
32
34
Total Enthalpy = 182 kJ/gmol (43.42 kcal/gmol)
Total Molecular Weight = 29.0
Specific Impulse = 2,571 m/s (262.2 s)
Mass Flow Rate = 8.655 kg/s (19.08 Ibm/s)
Thrust = 22.241 kN (5,000 Ibf)
Diffuser Exit Area = 0.0323 m2 (50 in2)
Scrubber Area = 0.6567 m2 (7.07 ft2)
Scrubber Pressure = 0.94 atm (13.8 Ibf/in2)
Air Inbleed = 0.1 kg air/kg propellant
No heat of reaction to form KC1
No depression of water vapor pressure due to dissolved salts,
79
-------
100
00
o
Present
Calculations
024 6 8 10 12 14 16 18 20
R , kg WATER/kg PROPELLANT
Figure 30. Comparison of velocity prediction with Garrett,
et al. (1972).
100
95
03
90
W
Present
Calculations
85
80
0 2 4 6 8 10 12 14 16 18 20
RW, kg WATER/kg PROPELLANT
Figure 31. Comparison of temperature prediction with
Garrett, et al. (1972).
-------
Quenching EquiljJjrjLum Predictions
Closed Duct -
The computer program described in Appendix "A" was run
to predict the quenching equilibrium conditions for a closed,
ducted quencher. The equilibrium (exit) pressure was assumed
to be 0.90 a tin (13.2 lbf/in2) and the rocket composition and
conditions were those of the composite propellant rocket de-
scribed in Section 2. Air was assumed to be entrained by the
ejector action of the rocket exhaust in the amount of ten per-
cent of the rocket mass flow rate. This air inbleed amount is
somewhat arbitrary and will be discussed later. The hydrogen
chloride gas in the rocket exhaust was assumed to dissolve in
the liquid water, with consequent heat release and depression
of the water vapor partial pressure.
The 2 meganewton (450,000 Ibf) rocket was used in the com-
putation. However, all the predictions, which will be shown,
except the volume flow rate hold for any size rocket provided
that the propellant composition and other parameters not
related to size are the same, and the following ratio is used:
Thrust = 40>000 N/m2 (16>50o lbf/ft2)
Duct Cross Section
This ratio was selected to provide a water vapor saturated
condition and a velocity of about 100 m/s trading off liquid
usage and duct diameter. It also provides a near maximum:
pressure rise (or suction) in the scrubber (Figure 25). A
summary of the imput parameters for the predictions is given
in Table 10.
The ratio of quenching water mass flow rate to rocket
propellant mass flow rate was varied from about 3.5 (the mini-
mum)to 20. The predictions are shown in Figures 32 to 37. The
temperature in Figure 32 rises above the normal saturation tem-
perature (97°C at 0.9 atm) because of vapor pressure depres-
sion caused by dissolved HC1. The pressure rise shown in Figure
34 illustrates the strong ejector action of the rocket exhausting
81
-------
TABLE 10. PARAMETERS FOR 2 MEGANEWTON ROCKET
QUENCHING PREDICTION
Rocket Exhaust Composition and Static Enthalpy
- As given in Tables 1, 2, and 4
Total Enthalpy = 194 kJ/gmol (46.4 kcal/gmol)
Specific Impulse = 2,590 m/s (264 s)
Mass Flow Rate = 772.2 kg/s (1,702 Ibm/s)
Thrust = 2 MN (450,000 Ibf)
Exit Area* = 1.53 m2 (16.5 ft2)
Duct Cross Section = 50 m2 (538 ft2)
Equilibrium (Exit) Pressure =0.9 atm (13.2 lbf/in2)
Air Bleed Fraction = 0.10.kg air/kg rocket exhaust
HC1 Heat of Solution = 74,800 kJ/kgmol (48,900 kcal/kgmol)
*Based on a thrust coefficient of 1.68 which is
appropriate for a chamber pressure of 10,300 kPa
(1,500 lbf/in2) and a ratio of specific heats of 1.15.
82
-------
6,000
CO
OS
m
E-H
02 4 6 8 10 12 14 16 18 20
R^ kg WATER/kg PROPELLANT
Figure 32, Flow conditions in a closed quencher
02 46 8 10 12 14 16 18 20
Rw, kg WATER/kg PROPELLANT
w;
Figure 33. Gas volume flow rate of 2 MN rocket
closed quencher.
-------
0.3
0
0 2 4 6 8 10 12 14 16 18 20
R , kg WATER/kg PROPELLANT
Figure 34. Pressure rise in quencher.
0.3
0.2
tq
0.1
0.0
1,000
10,000
F/AD, N/m2
100,000
Figure 35. Pressure rise versus thrust to duct area
ratio.
84
-------
OO
u~\
o-
(—(
>-J
a
S3
2
m
ex
0.1
o 0.05
3
U,
1.0
0.01
5 10 15
RW, kg WATER/kg PROPELLANT
20
Figure 36. Concentration of HC1 absorbed in quencher
liquid.
0.5
0 24 6 8 10 12 14 16 18 20
RW, kg WATER/ kg PROPELLANT
Figure 37. Water vapor volume fraction.
-------
into the duct. It shows that a higher back pressure than 0.9
atm (absolute) is allowable before the rocket exhaust would
spill around the duct. Figure 35 is a plot of the pressure
rise for other than design thrusts based on Rw=10 and Ra=^'^
to determine the range of thrusts allowed for a certain design
duct size. When the pressure rise is too low the rocket will
begin to be a poor ejector and spillage will occur. Figure 35
is shown for a water ratio of 10 but will be similar for other
ratios between 5 and 15. The HC1 concentration abosrbed in the
scrubber liquid (neglecting solids) is shown in Figure 36. It is
important to note that for water/propellant mass ratios below 5
the liquid is very concentrated with HC1. The high concentration
would tend to negate the assumption that no gaseous HC1 exists
in the quencher outlet because of the higher vapor pressure.
However, the fraction of HC1 gas would be very small since so
much water vapor is present. The water vapor fraction is shown
in Figure 37 and the gas composition is shown in Table 11.
TABLE 11. EQUILIBRIUM QUENCH COMPOSITION
Dry Basis Total Gas
Gas Mole Fraction Mole Fraction
§ Rw = 10
CO
C02
HC1
H2
H20
N2
02
Molecular Wt.
*The mole fraction of HC1
values of R (R <81 .
W W " J
0.267
0.054
0 *
0.421
0
0.258
0
17.9
would not
86
0
0
0
0
0
18
actually be
.036
.007
0 *
.058
.864
.035
0
.0
zero at low
-------
Open Channel -
A slightly modified computer program was used to compute
quenching equilibrium conditions in an open quencher. Constant
pressure throughout the system was assumed so that the flow area
became a predicted variable. The parameters were the same as
used for the ducted quencher (Table 10). The prediction of flow
area, velocity, temperature and density are shown in Figure 38.
The water vapor fraction and HC1 concentration are very similar
to those for the ducted quencher. The volumetric flow rate can
be obtained from the area and velocity curves.
This prediction is based on several rough assumptions. Air
is assumed to mix completely with the rocket flow and only in
the amount equal to 10% of the rocket mass flow. Estimation of
the amount of entrained air and the rate of mixing is very tedious
except for completely open systems. Models for plume mixing with
air only are given in CPIA Publication 263 (1975) but these are
not applicable here because of the water addition, and because
walls and other boundaries may be present.
Afterburning Considerations
Solid rockets produce and emit considerable amounts of hydro-
gen and carbon monoxide. These gases will react with oxygen in
the air to produce water and carbon dioxide under the proper
conditions of concentration, mixing, velocity, and temperature.
Table 12 presents the accepted flammability limits and spon-
taneous ignition temperatures. If the combustion occurs in an
enclosed structure the violence of the reactions may cause explo-
sions. Normal, open-air, rocket firings are accompanied by almost
complete afterburning of hydrogen and carbon monoxide since very
little CO or H2 is detectable in the vicinity of the rocket.
The use of an enclosed design for an exhaust gas scrubber
required consideration of the possibility of combustion. Methods
for controling the potential combustion are:
87
-------
TABLE 12. COMBUSTION PROPERTIES OF H2 and CO
IN AIR AT STANDARD CONDITIONS
Stoichiometric Flammability
£as Mixture Limit, % by Vol.
% by Volume Lower Upper
H2 29.50 4.0 74.2
CO 29.50 12.5 74.2
Spontaneous
Ignition
Temp. °C
571
609
88
-------
OS
w
E-i
H
2.2
2.0
1.8 !l
1.6
1.4
1.2
1.0
0.8
0.6 ;;
0.4
0.2
0
0 2 4 6 8 10 12 14 16 18 20
R , kg WATER/kg PROPELLANT
w
Figure 38. Flow conditions in an open quencher
89
-------
1. Limiting the amount of air entrained into the system
so that combusion, if any, is slight.
2. Keeping the gas mixtures too cool to spontaneously
ignite.
3. Removing all spark or charge producing mechanisms which
would ignite the gases by grounding, etc.
4. Purposely burning the hydrogen and carbon monoxide
with afterburners under controlled conditions.
The following discussion explains these four considerations.
Entrained Air -
The rocket itself acts as an ejector, which is a type of
pump. The gas (air) that surrounds the nozzle exit is en-
trained by the nozzle exhaust and pumped into the scrubber.
Charts for the amount of pumping by single-stage ejectors,
such as are given in section 5 of Perry § Chilton (1973), may
be used to estimate the amount of entrained air. The ejector
charts overestimate the entrainment because complete mixing of
the entrained and pumping gases is assumed. In most scrubbers
the ratio of the diffuser duct area to the rocket nozzle
throat area would be between 15 and 20. For this range of
area ratios the amount of entrained air is between 10 and 20
percent of the rocket exhaust. Since these percentages re-
present an overestimate, the 10 percent value is the more
correct figure. This 10 percent value was used by Garrett (1972)
and for the baseline case presented in this report.
Limitation of this small amount of entrained air by
closing the gap between the rocket nozzle exit and the dif-
fuser of the scrubber is not a good idea. For very small gaps
the pressure in the gap would be much lower than atmospheric
because the air would have a high velocity due to the flow
constriction. The pressure distribution on the outside of
the rocket nozzle would then be other than uniformly atmos-
pheric and cause incorrect thrust measurements. The possibil-
ity of direct attachment of the rocket nozzle to the scrubber
diffuser also exists. The seal for such an attachment would
90
-------
have to be unable to transmit axial loads to the rocket nozzle
but be able to withstand the high temperature environment.
Also, the seal would be subject to the wall friction forces
from the rocket gases. Such a seal may be possible but the
design is not obvious.
Cooling -
The second means of reducing the possibility of combus-
tion is to keep the gases below the temperature of spontaneous
ignition. In a wet quencher the injected water enters directly
into the supersonic exhaust stream so that cooling is initiated
while the flow is too fast for combustion. By the time the flow
has slowed to subsonic velocities the evaporating water has cooled
the flow to below the boiling point of water. The boiling
point of water is considerably below the spontaneous ignition
temperature of the gases.
Static Electricity -
The third means of reducing the possibility of igniting
the exhaust gases is to remove spark or charge producing mech-
anisms from the scrubber. The solid aluminum oxide particles
acquire charges because of the extreme temperatures in the
rocket chamber where they were formed, and by friction with
the surrounding gas ions and particles. Water droplets ac-
quire charges because of their contact and friction with the
high speed, hot rocket exhaust. The charging of the particles
due to high temperatures and friction is unavoidable. These
particles flowing in the scrubber may induce electrostatic
fields in the scrubber walls and be attracted to the walls.
Under severe charge concentration conditions the particles
could discharge at the walls with an accompanying arc or spark.
These severe conditions were not thought likely to occur because
the size of particles required to hold large enough charges to
discharge with an arc is much larger than is expected from
the rocket exhaust. Additionally, both the particles and the
scrubber walls would have to have high resistivities so that
91
-------
the charges would not be easily conducted away. Experience in
the AFRPL pilot scrubber (Section 4), however, has shown static
electricity to be very probable.
Use of grounded conducting (metal) materials for the scrub-
ber walls ensures that sparking will not occur no matter what
the charge or resistivity of the particles may be. The chance
of sparking between particles is very remote because the parti-
cles carry so little charge and are usually like-charged .anyway.
The use of non-conducting scrubber wall material may allow
static electricity to build up enough on the wall to cause
sparking. This chance of sparking should be reduced by wetting
the walls, which would normally occur in a wet scrubber.
Controlled Afterburning -
The final means of reducing the possibility of combustion
of fuel gases in the scrubber is to purposely burn these gases
under controlled conditions. The use of after-burners would
require consideration of the following factors:
1. Problem of injecting air or oxygen into the hot, fast
rocket exhaust stream.
2. Problem of obtaining adequate mixing of the fuel gases
and the oxygen to ensure complete combustion.
3. Problem of the increase in flow enthalpy (above the al-
ready extremely high value of the rocket exhaust) due
to the after-burning.
Since the purpose of the after-burning is to eliminate the
possibility of unwanted combustion in the scrubber, the after-
burner would have to be located upstream of the low velocity
regions in the scrubber where unwanted combustion is most likely
to cause explosions. In a wet scrubber system the air injec-
tion apparatus would be best located just downstream of the
water injectors because the flow would be cool enough there to
permit stable combustion. An additional bank of water injec-
tors would then have to be used after the after-burning region to
cool the gas, reduce the volume, and slow the flow again. The
mixing duct portion of the scrubber would necessarily be longer.
92
-------
Equilibrium calculations - To make the preliminary calcu-
lations simple, the processes were assumed to occur at one at-
mosphere pressure and achieve chemical equilibrium. The only
reactions of importance were:
CO + % 02 -* C02
H2 + h 02 -*• H20
There are about twenty-five reactions of importance kine-
tically in CO/H2/air systems according to Edelmen et al. (1975).
However, these reactions need only be considered when determining
small amounts (parts per million) of various reaction products,
where we are interested in the major products, measured in parts
per hundred (percent).
The computer program described in Appendix "A" was run for
different amounts of injected air. The effect on velocity for
the closed duct quencher, using the same parameters as before
is shown in Figure 39.
The amount of air required for complete combustion of the
H2 and CO, for a typical composite propellant, is found from
the following equation:
ra -;„ mr /mu i mn \
_£i£ = 11.5 _C + 34.5 [-£ - \ -2\ (31)
mR mR ^mR 8 mR^
where the subscripts refer to:
air - air
R - total rocket
C - carbon (in CO)
H - hydrogen (H2)
0 - oxygen (in CO)
From Table 1 for the composite propellant, this equation re-
duces to:
m
R =
air = 1.15 (32)
a mR
This ratio represents an extremely high volume flow rate of air.
93
-------
200
150
>H
H
O
nj
W
100
so
2 46 8 10 12 14 16 18
RW, kg WATER/kg PROPELLANT
20
Figure 39. Effect of entrained air on equilibrium
velocity.
94
-------
The other variables such as temperature and pressure are
affected also, but not as drastically as velocity. The velo-
city is increased 74% at R^ = 10 from the 10% air ratio of the
stoichiometric air ratio (115%). Thus, if a 115% mass ratio of
air to rocket propellant were used the cross sectional area of
the scrubber would have to be increased 74%, in order to keep
the velocity below about 100 m/s. The blowers and related equip-
ment would also increase the complexity and cost of the system.
Conclusions
Afterburning is practically unavoidable either within the
scrubber system or at the exit. The addition of enough air to
completely burn the fuel gases (CO and H2) within the quencher
is too expensive an alternative. The assumption that a 10%
mass ratio of air to rocket propellant will be entrained and
burned has some basis in ejector theory. Because of static
electricity or local hot spots combustion will likely occur
where the exhaust gases contact and mix with air at the exit
of the scrubber system. The scrubber must be so designed that
pockets of combustible gases cannot form in the system and the
system exit region must be relatively unaffected by afterburning
in close proximity.
95
-------
PERFORMANCE REQUIREMENTS
The primary performance requirement of the scrubber system
is the removal of hydrogen chloride gas with the efficiency
specified in Section 3. Although collection of aluminum oxide
particulates is not necessary their removal efficiency can be
predicted.
Number of Transfer Units for Gas Absorption
Both HC1 and HF are extremely soluble in water so their
absorption rate is gas-phase controlled. Consequently, the
number of overall gas phase transfer units for hydrogen chlo-
ride or hydrogen fluoride gas absorption in an aqueous solu-
tion of a base is:
/ E \
NQG = - In (l - 4J (33)
where Nfir = number of transfer units (NTU)
ED = required efficiency, I
K
Thus, for an efficiency of 99.6%, 5.52 transfer units are
required.
Particulate Removal Efficiency
Since particulate removal is not a primary objective a
detailed analysis of removal efficiency will not be presented.
A recent discussion of the latest performance prediction
techniques can be found in Calvert, et al (1972) and Yung, et al
(1976). The "Scrubber Handbook" by Calvert, et al covers all
types of scrubbers. Yung presents prediction equations for
venturi scrubbers which can be adapted to a gas-atomized rocket
scrubber in which category the quencher falls.
A typical operating condition for the rocket quencher/
scrubber would be a liquid to rocket mass flow rate of ten
(Rw = 10) and an air inbleed mass ratio of 1/10 (R = 0.1)
as discussed previously in this chapter. The corresponding
velocity and QL/QG are 84 m/s and 0.015 m3/m3, respectively.
The mass median diameter of particulates is usually between
96
-------
2 and 10 ym with a geometric standard deviation of about two.
Using Yung's model and the assumptions, operating condi-
tions, and a mass median particle diameter of 5 ymA an over-
all efficiency of particulate removal of greater than 99.9%
is predicted. This figure would be slightly smaller for
smaller mass median diameters. The 50% efficiency size
(cut diameter), based on the operating conditions is about
0.33 ymA, so the high efficiency would remain so long as the
mass median diameter were greater than about 2 ymA. For
alumina particles with a density of about 3.7 g/cm3 the
aerodynamic diameter is approximately twice the actual (physical)
diameter for diameters above 1 ym.
97
-------
CONVENTIONAL SCRUBBER DESIGN
In Section 5 potentially suitable designs were discussed.
Since gas absorption was the primary objective, a spray packed
column, or plate column type of device would be the most prac-
tical choice. Of the spray types the gas-atomized type seemed
to be the cheapest since energy was available in the gas stream
to provide atgrnization of the liquid at high relative velocity.
In this subsection the scrubber types are narrowed down
to one type on the basis of economic considerations. Prelim-
inary cost estimates of the three types show the gas atom-
ized spray scrubber to be the least costly. A detailed pro-
cess design is then made for the spray scrubber which primar-
ily involves determining the length required for the mass
transfer to take place.
Preliminary Sizing of Spray and Column Scrubbers
Preliminary sizing for cost comparison purposes is based
on the large, 2 meganewton (450,000 Ibf) thrust, rocket. The
scrubber has to operate on the gas leaving the quencher. A gas
velocity of 100 m/s was selected as reasonable as a basis for sizing
the scrubber. The following table summarizes the flow conditions:
TABLE 13. SCRUBBER INLET CONDITIONS FOR PRELIMINARY SIZING
Metric English
Volume flow rate 5,000 Am3/s 10.6xl06ACFM
Gas density 0.53 kg/m3 0.033 lbm/ft3
Although these conditions correspond to a quench water
ratio, RW, of 5.6 kg water/kg propellant, which is a little
above the minimum required for quenching only, we assume that
no HC1 has been absorbed in the quencher water.
98
-------
Spray Scrubber Diameter -
The gas-atomized spray scrubber is mechanically the same as
quencher. Operating equations are the same, so that the duct
diameter is as previously stated:
D . /4 Thrust,N
c ^TT 40,000 ' C
Plate Column Diameter -
Calvert et al. (1972) present an equation for plate col-
umn diameter based on the allowable superficial gas velocity.
The allowable velocity is based on empirical correlations for
the onset of priming and/or entrainment of the liquid. The
equation for column diameter is:
Dc =
TTE \PL-PG
Qr
G ' u ' (35)
where "a" is an empirical constant with dimensions of velo-
city. This empirical constant is given in the following table:
TABLE 14. EMPIRICAL CONSTANTS FOR EQUATION (35)
Type of Plate Column a, m/s
Bubble Cap Tray 0.043 >
Sieve Tray 0.057
Valve Tray 0.072
Packed Column Diameter -
Packed column diameter can also be calculated by a method
presented in Calvert et al. (1972). The allowable velocity is
limited by the onset of flooding and the design superficial
velocity is usually set at about 75% of the flooding velocity.
This flooding velocity is available from charts given in Calvert
et al. (1972) or section 4 of Perry § Chilton (1973). The flood-
99
-------
ing velocity depends on the type and size of the packing. An
average "packing factor" for 5 cm diameter packing material is
about 160 m2/m3. The charts for flooding velocity have "x"
and "y" axes which are related to column diameter, gas mass
flow rate and liquid mass flow rate for the present conditions
by:
mT
x = 0.023 ~ (36)
m,-,
(3
y = 0.050 -\ (37)
uc
Comparison of Diameters -
For the 2 meganewton [450,000 Ibf) thrust rocket and con-
dition described in Table 13 the following table compares dia-
meters :
TABLE 15. COMPARISON OF SCRUBBER DIAMETERS
Diameter, m
Type L/G= 1 &/m3 L/G = 2 £/m3
Gas Atomized Spray 8.0
Bubble Cap Plate 58.4
Sieve Plate 50.7
Valve Plate 45.1
5 cm Packing 43.8 46.0
Spray Scrubber Length -
The length required will be detailed later in this sec-
tion. It will be shown that only a few meters are required
for mass transfer. For comparison purposes we can assume that
5 meters is adequate.
Plate Column Height -
To make a preliminary performance estimate for a system of
this kind one can assume a plate efficiency of about 75%.
100
-------
Since 5.5 transfer units are required a total of 8 plates are
needed. The plate spacing in industrial usage is about 0.5
meters, so a height of about 4 meters is needed. Including
ends, a total of about 5 meters would be adequate.
Packed Column Height -
As a rough rule the height of a transfer unit is between
0.5 and 1 meter. Using the mid-range value of 0.75 meter, a
total column height of about 4.2 meters is required. Adding
ends, about 5 meters would be the total requirement.
Auxiliaries -
Auxiliary equipment includes the caustic tank, pumps,
piping, quencher, sewer, and drying pond. These items are com-
mon to all the scrubber types and need not be costed for com-
parison purposes. It is assumed that the spray scrubber needs a
low velocity entrainment separator, while the other types can
operate at low enough velocities that the amount of entrainment
they generate is acceptable for discharge.
It will be seen that even with this disadvantage the gas-
atomized spray is cheaper than the others so the foregoing
assumption is acceptable. Since so -much power is available to
the system from the rocket exhaust a cyclone type separator, as
described later, is cheapest for the gas-atomized spray scrubber,
For the 2 MN rocket the entrainment separator would consist of
7 cyclones, each 8 m in diameter and 24 m high.
Preliminary Cost Estimates for Spray and Column Scrubbers
Cost estimates were made based on a number of sources,
including Chilton (I960), Calvert (1968), Peters and Timmerhaus
(1968), Popper (1970), Guthrie (1974) and Lee Saylor (1976).
A consensus summary of the cost estimates for the 2 MN rocket
is presented in Table 16 . The first three figures presented
are for the plate or packed column scrubbers while the last
figure represents the ducted spray. These figures are at best
rough estimates and are presented for comparative studies only.
There are two reasons for this; first the designs for each
process can at best be considered conceptual. From a compar-
ative economic standpoint this is not serious since the over-
all processes evaluated are essentially similar and differ
101
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TABLE 16. CAPITAL COST ESTIMATES FOR LARGE ROCKET SCRUBBER
(in millions of dollars, mid-1976)
Scrubber Type
Scrubber Cost
Auxiliaries*
Total
Sieve
Plate
13.7
2.7
16.4
Bubble
Plate
15.5
2.7
18.2
Packed
Plate
12.4
2.7
15.1
Spray §
Cyclone
2.2
2.7
4.9
*Auxiliaries include: cooling duct, deflector, waste treatment,
piping, caustic facilities, and sewerage.
only in the type of scrubber being used for each system. The
second reason which limits application of the figures for ab-
solute purposes is the nature of the job being estimated. The
rocket scrubber process is unique from the viewpoint that no
systems are in current operation from which to base economic
values.
In large part, dollar values have been derived from basic
cost estimation fundamentals. This can often give considerable
errors of magnitude in the estimates made for each process com-
ponent. In general these errors tend to be smoothed out on the
summation of the item costs for total process cost purposes.
In consideration of these limitations the figures presented
for the capital cost estimates may be taken as plus or minus
30%.
The estimates may therefore be considered significant
because they give positive information regarding the selection
of the final process design.
The figures presented are total complete project costs
for a mid-1976 base. Note that the total process costs have
been split into two parts, namely for the scrubbing section
and for the auxiliaries. "Auxiliaries" in this case refers
to the cooling duct, deflector, waste treatment, piping,
caustic facilities and sewerage. The auxiliaries costs are
102
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identical for each process and have been separated from the
total cost figure to highlight the effect of scrubber costs
on the project.
Where materials of construction need to be specified ,
carbon steel 0.64 cm (0.25 in) thick is used. The scrubber sizes
in some cases were so large that estimating techniques were
difficult. In these cases costs are based on a number of
smaller diameter units with an equivalent total flow area.
It is evident that the ducted spray scrubbing
system is superior to the other units from the capital cost
standpoint. At $4.9 million its cost is approximately one-
third of its closest rival, the packed column.
Even allowing for the magnitudes of error inherent in
the costing process there is sufficient spread in the capital
cost numbers to conclude that the spray is the most promising
device for cleaning the rocket exhaust gases.
Process Design for Spray Scrubbing
The process design for the gas-atomized spray scrubber
involves specification of the liquid flow rate, liquid com-
position, duct cross-section, and mass transfer length. The
equilibrium conditions, assuming an infinite length scrubber,
are calculated with the same equations as were used for the
quencher. There is, in fact, no difference between the quen-
cher and the gas-atomized spray scrubber other than that
more mass transfer can occur in the scrubber. Thus, the
only new design equations to be presented will concern the
mass transfer.
Equilibrium Conditions -
For the specification of liquid flow rate, liquid com-
position, and duct cross-section the quencher and scrubber
can be considered one unit. The liquid supply would come
from one source rather than two to reduce costs. The duct
would be one section of uniform area. The liquid supply
would, however, have to be injected at two different distances
103
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from the duct entrance since the prediction of the gas-atomized
drop size is very difficult based on injection into a hot,
supersonic stream. This two staged liquid injection will not
affect equilibrium calculations. The conservation equations
were described in the subsection on quenching equilibrium. The
air entrainment rate is assumed to be 0.1 kg air/kg propellant
as for the previous quencher calculations.
The criterion for quenching design was to provide a satu-
rated gas at about 100 m/s (328 ft/s) velocity, such as com-
monly occurs in a venturi scrubber using an economic tradeoff
between liquid flow rate and duct area. The duct area selected
was:
= Thrust 2 [38)
D 40,000 '
where the thrust is in newtons. This criterion is valid for
the scrubber section also. The quench water mass flow rate
required is about 5 times the rocket mass flow rate based on
the figures previously presented. The required scrubber li-
quid flow rate will be developed in the mass transfer analysis.
The scrubber liquid composition should be a basic solu-
tion so that evaporation of the liquid after it leaves the
scrubber will not cause vaporization of the acid gases and to
reduce corrosion. The very high solubilities of HC1 and HF in
water minimize the need for a basic solution to improve mass
transfer rate. Table 17 shows the costs of four commonly used
basic chemicals, based on vendors' quotes for mid-1977 in the
Los Angeles area.
TABLE
Chemical
Ca(OH)z
NaOH
KOH
Na2C03
17. BASIC
Cost per
1,000 kg
$105
528
655
148
CHEMICAL COSTS, MID-1977.
Cost for 10 Large Rockets
(99,150 kg HC1)
$10,570
57,430
99,750
21,340
104
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Besides cost difference these four chemicals have different
solubilities, so that slurry moving equipment and piping are
required for some. The solubilities are (based on Weast, 1971):
TABLE 18. CHEMICAL SOLUBILITIES
Chemical Solubility, g/100 cm3 solution
30°C 1QQ°C
Ca(OH)2 0.153 0.077
NaOH 119 347
KOH 126 178
Na2C03 38.8 45.5
Garrett et al (1972) objected to Ca(OH)2 because of the
insolubility of calcium fluoride, CaF2, but that would seem
to be a minor point for the solid rocket. So much insoluble
aluminum oxide (A1203) will be present that the CaF2 would be
insignificant. For the solid rocket the entrainment separator
and plumbing must be able to handle an appreciable amount of
undissolved solids.
The cost of the systems required to handle either a
slurry of Ca(OH)2 or a solution of Na2C03 must be weighed
against the chemical costs. This tradeoff depends primarily
on the use frequency and projected system lifetime. Based on
frequency of 10 tests per year and a 10-year life Na2C03 is
recommended.
The minimum concentration of the base should be a stoichio-
metric balance with the amount of HC1 and HF present. The
maximum concentration should be the smaller of twice stoichio-
metric or 50% by mass, which is about the highest slurry
concentration that can be pumped and piped economically. The
stoichiometric mass flow ratio with HC1 is:
g Na2C03 _ 106 a -i 4q (39)
—-2(36.46)
105
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For the composite propellant rocket the amount of HC1 emitted
is 17.1 mole %, or 21.4 mass % of the total rocket exhaust.
Thus, the stoichiometric amount of Na2C03 is:
TOT, «. • T, = 0.31 m (40)
Base, stoich r
and the mass concentration of Na2C03 is:
C
Base, stoich = -- (41)
where "R" is the scrubber liquid to rocket mass flow rate
ratio. At double the stoichiometric amount of base a 50%
concentration corresponds to R =0.81, which is a much smaller
ratio than will be required for the scrubber.
Equilibrium Mass Transfer -
HC1 and HF produce a temperature increase when dissolved
in water and this causes their solubilities to decrease. While
it is possible to account for the heat effect on the mass trans-
fer rate, it requires a complex computation. For simplicity it
is assumed that the interfacial liquid temperature is 100°C,
which is about the maximum possible. This is a conservative
assumption which is later shown to be acceptable. The equi-
librium line on Figure 40 is based on data in section 3
of Perry § Chilton (1973) for 100°C.
For co-current physical absorption (i.e., no base used)
the operating line shown in Figure 40 is based on the following:
Inlet: x. = 0.0
in
yin = 0.024
Outlet: y = (0.004) y = 9.6 x 10"5
where x = mole fraction HC1 in liquid
y = mole fraction HC1 in gas
and "yin" is based on figures presented in the quencher analysis
106
-------
0.0020
iH
U
DC
H
C_J
W
J
s
0.0015
§ 0.0010
0.0005
0
Operating Line
^37622x^0.024
0
0.01 0.02
x, MOLE FRACTION HC1 IN LIQUID
Figure 40. Physical absorption of HC1.
107
-------
section for R - 5.6, but assuming no HC1 absorption in
quencher. The intersection of "yout" with the equilibrium
line determines "xout":
xQut = 0.066
The slope of the operating line is the minimum liquid-to-gas
mole ratio:
/M
I—-I =3.62 moles liquid/moles gas
\ ^r /
X G/ min
Based on the figures presented in the section on quencher
analysis, the liquid-to-gas mole ratio has been calculated as
a function of the scrubber liquid flow rate and is shown in
Figure 41 . The zero point on this figure was taken at a
quencher "R " of 5.6, so that the total system minimum water
requirement for a scrubber liquid-to-gas mole ratio of 3.62 is
as follows:
RT * 5.6 + 8.8 = 14.4 kg water/kg propellant
Since the phase equilibrium was based on a conservative 100QC,
an RT = 15 should represent an adequate operating flow rate of
water. The mass transfer efficiency for a basic solution
would be better than that for physical absorption in water.
The mass concentration of Na2C03 , based on a double stiochio-
metric ratio and equation ( 41) should be about 4% in the
total liquid.
Mass transfer coefficient - In a steady state process of
absorption, the rate of mass transfer through the gas film may
be expressed by:
N = kG (pG - p.) (42)
108
-------
10
03 8
w
1 6
i
0 2 4 6 8 10 12 14 16 18 20
R , kg SCRUBBER WATER/kg ROCKET PROPELLANT
•5
Figure 41. Liquid to gas ratio in scrubber.
109
-------
where "N" is the molar flux of transferring component
(kgmol/s-m2) , "k " is the gas phase mass transfer
coefficient based on conditions at the interface (kgmol/s-m2
atm) and "p " and "p." represent the partial pressure of the
diffusing component in the bulk of the gas and the interface,
respectively. It is normally more convenient to rewrite the
rate expression in terms of the overall gas phase coefficient
based on equilibrium conditions, i.e.:
N = KG (pG - pe) (43)
where "p " is the equilibrium partial pressure.
c
If the mass transfer process is gas phase controlled,
i.e., kp-Kr replacement of the film coefficient with the over-
all coefficient to express the mass transfer rate is justified.
Calvert (1968) suggests the system to be gas phase controlled
if the Henry's Law constant is less than 0.2 atm/ (kgmol/m3) ,
and liquid phase controlled if Henry's Law constant is greater
than 200 atm/ (kgmol/m3) . In our case of dilute HC1 at 0.9 atm
pressure the lower value corresponds to y/x=0.004. The Henry's
Law constant corresponds to the slope of the "equilibrium line"
at the operating point. This slope is y/x=0.024, based on
Figure 40, which is much closer to the gas phase control cri-
terion than to the liquid phase control criterion.
Gas film transfer coefficient (k.J - Gas film transfer
u
coefficient (kg) is calculated using the semi-theoretical equa-
tion of Frb'ssling in Calvert et al . (1972):
k R T dd
u b 0. , i
J Sc^ (44)
where Re =
Sc =
dd ur P
JL
U *1
yg
P D
G G
— L, . -r U . J J ^.
DG
= Reynolds number
= Schmidt number
110
-------
Rp = ideal gas law constant
T = absolute temperature, °K
d-, = drop diameter, m
DG = gas phase diffusivity , m2/s
ur = relative velocity between drop and gas, m/s
P = density, kg/m3
yG = viscosity, kg/m-s
Subscript "z" refers to point conditions at distance
"z" from the point of atomization
The effect of drop velocity on the mass transfer coeffi-
cient is apparent in the above expression. For the rocket
scrubber, this implies constantly changing values for "kg"
o
with time (or drop travel distance) as the atomized liquor
accelerates along the plume path. This point will be returned
to later in discussing the scrubber length required to com-
plete the necessary mass transfer.
Atomized drop diameter, d,, was estimated from the Nuki-
yama and Tanasawa equation in Calvert et al. (1972):
(0 V'5
in (45)
G /
at standard conditions
where Q^ = liquor flow rate, m3/s
QG = gas flow rate, m3/s
Up = gas flow rate, m/s
The instantaneous drop velocities were computed from the
equations of motion for drops accelerated into a gas stream
according to Dallavale (1948) :
d ud Pr CR 1 d? u ud
ix = G R 4 d r x (46)
dt 2m
111
-------
(47)
- = P i i — i i
dt
where u
-------
For the co-current flow conditions found in the gas-
atomized spray scrubber, the required absorber length, z, may
be expressed by the relationship:
n Jo KG a ^ = y y } (49)
G ° ' e/lm
where y t § y- = the mol fractions of HC1 in the gas phase
out in j. j.i_ 11- • 6,..,
at the scrubbing section outlet and inlet
respectively
yg = the mol fraction in the gas phase at equil-
ibrium with the liquor
(y ~ ye) im = ^e l°garithmic mean value of I y - ye) for
' the inlet and outlet conditions
This assumes a linear equilibrium relationship over the
concentration range of interest. As can be seen in Figure 40
this is a justified approximation over the range from "y. " to
"v "
7 out '
The expression on the right-hand side of equation (49) is
commonly referred to as the number of transfer units required
for the operation (NOG).
For the present HC1 absorption conditions, the number of
transfer units required for the process is 5.52. The gas mole
flow rate, n,-,, may be expressed in terms of linear gas flow
b
rate, ur, according to the following expression:
nr = i-—(50)
( G M
where M = the molecular weight of gas
For pG=0.53 kg/m3 and P=0.9 atm, equation (49) reduces to
K~ a dz = 0.2 ur (51)
b b
Both "kG" and "a" vary with position along the drop tra-
jectory. Moreover, it has been shown that for the case of gas
113
-------
phase limiting, kG, the film coefficient, may be used to rep-
resent KG, the overall gas coefficient. Thus, equation (51)
may be written:
Dr ^ ^
KG = -—H_ = 2 + 0.552 Rez2 Sc * (52)
The mass transfer area, a, may be related by the following
expression:
az = ^[^1(^1 (53)
LJ
Equations (<.5) , (46), (47), (52), and (53) afford a solu-
tion to the mass transfer length as a function of gas velocity
as expressed by equation (51). This is best accomplished by
computer programming.
The procedure adopted was to obtain an expression for the
product of "Kra" as a function of drop travel distance, z, for
a number of gas velocities ranging from 10 m/s to 100 m/s. The
program was set up to compute the drop size according to equa-
tion (45). Instantaneous drop velocities, uj and u^ , were
x . y
then calculated from equations (46) and (47) using the method
of finite differences for successive increments of time , At.
Knowing the drop velocity at each time interval, it is possible
to compute instantaneous values for "K " and "a" from equations
(52) and (53), respectively.
The next stage was to obtain an expression for the product,
K_a, as a function of drop travel, z. It was found that the
results could be expressed by the form:
KQa = ClZC2 (54)
where Ci and C2 are constants.
The expressions obtained for each gas velocity were found
to represent the calculated data with an exceptionally high
index of fit (approximately 1.0 in all cases).
114
-------
Reduction of the expressions for "Kga" to the form pre-
6
sented in equation (54) simplifies calculation of the drop
travel required for mass transfer. Thus, substituting the
expression for "KGa" in equation (51} gives a simplified ex-
pression for the transfer length required for mass transfer:
/ C1(z)C2 dz = Q.2 ur (55)
Jo G
Note that the constants "Ci" and "C2M are independent of
mass transfer length, z. However, they are dependent on gas
velocity, Up.
Using the expression presented in equation (55) it is now
possible to compute the plume transfer length as a function of
gas velocity. This is shown in Figure 42,
The important point brought out in Figure 42 is the small
plume length required for mass transfer for the practical gas
flow rates anticipated for the rocket. Even at the exception-
ally low velocity of 10 m/s, travel length is only 3.5 meters.
Moreover, for velocities 30 m/s to 100 m/s the required length
is relatively consistent at 1.75 m.
What this Implies in practical terms for the rocket study
is that the plume length required for mass transfer is not
the critical design parameter. The simplifying assumptions
which were made in the computation are acceptable because they
do not have a significant influence on the required contact
length for mass transfer. More important is the length required
for .momentum and heat transfer in slowing down the gases to a
velocity suitable for particle and drop separation and cooling
the hot gases to the scrubbing liquid boiling point.
Conclusions - The plume length required to complete the
mass transfer of HC1 gas to scrubbing liquid has shown that a
distance of 3.0 meters is adequate to handle all practical gas
velocities likely to be encountered in designing the rocket
scrubber.
115
-------
H
CD
2:
W
3.0
2.0
1.0
40 60 80
GAS VELOCITY, u, m/s
100
Figure 42. Effect of gas velocity on plume length required for
HC1 mass transfer to scrubbing liquor at 0.9 atm.
116
-------
Coupling Effects between Scrubber and Rocket
From the point of view of the rocket test engineer it is
very important that the exhaust scrubber not affect the rocket
performance. The effects of the scrubber on the rocket were
discussed briefly in Section 2 and the first part of this section
(6). It was pointed out that because of the supersonic
nature of the rocket exhaust the scrubber could not affect
the rocket chamber conditions of pressure, temperature, and
and propellant mass flow rate during normal operation. It
is possible that the pressure on the exterior of the rocket
nozzle expansion cone and consequently the measured thrust
will be affected by the scrubber.
As discussed in the subsection on definition of the pro-
cess the rocket nozzle acts as an ejector pump to entrain air
into the scrubber. In order to limit the amount of entrained
air the gap between the nozzle exit and the scrubber entrance
should be small. The smaller the gap the higher the velocity
of the entrained air past the nozzle exit. And, according to
Bernoulli's equation, the lower the static pressure on the out-
side of the nozzle exit. An estimate of the amount this ambient
pressure is decreased based on the assumed mass influx of air
of 10% of the rocket mass and a co-planar arrangement yields
the equation: (Bernoulli's equation using parameters of the large
rocket. )
- 1 - , (56)
where p '= reduced ambient pressure
^a
p = ambient pressure
3-
f = ratio of scrubber entrance diameter to rocket
nozzle exit diameter
For a 10% pressure reduction f = 1.16 which for the large (2MN)
rocket means that .the difference in radii (gapj between the
scrubber entrance and the nozzle exit would be 11 cm (4.4 in).
117
-------
This lowered pressure (Pa) affects the thrust as seen in
the thrust coefficient equations for an ideal rocket from Hill
and Peterson (1965):
where
/
/2Y2 /
Y-l ll
Y =
pe =
po =
pa =
e =
2 \ (Y+I)/(Y i)
H-l]
r /„ \ CY-D/Y
/pe \
i -I e )
-1- i „ /
\ P /
ratio of specific heats
exit pressure,
N/m2
chamber pressure, N/m2
ambient pressure, N/m2
nozzle area expansion ratio
a
(57)
and
e =
Y
2 *
,1 ^p
2
Y-l
e \ Y
J \
r/T> \ x
/pe\ -
[Uo/
2(Y-D
-Y 1
Y _-,
(58)
The obvious effect of the reduced ambient pressure is in the
reduction of the term in equation (57):
which increases the thrust coefficient, C^. The other effect
r
is that, because of the reduced pressure on the nozzle outside
surface the exit pressure, p , is effectively increased. Put
C
another way, the net pressure force acting on the rocket in the
direction of the thrust (which creates the thrust) is reduced.
This effective exit pressure increase can also be considered
an effective nozzle exit area and length decrease. This length
decrease would be to the location on the nozzle exterior surface where
the pressure was equal to the ambient pressure. This could be
determined experimentally by locating static pressure taps at
various distances forward from the nozzle exit plane.
118
-------
To illustrate the effects, equation (57) has been plotted
for y = 1.15 in Figure 43. An example of the effect of a 50%reduced
effective exit area and a 50% reduced ambient pressure is shown.
The net effects cancel, so that the thrust coefficient, Cp, is
not changed. The 50:50 relationship between reduced effective
nozzle area and ambient pressure is strictly for illustrative
purposes and would have to be determined experimentally. It
should be noted that if the design point were located at a
greater than optimum expansion ratio the effects would not
cancel and a higher thrust would result.
Conclusion -
For most experimental rockets which are underexpanded the
effect of the scrubber on thrust should be small,, e.g. less
than 5%. This coupling effect should not be considered the
only impact that a scrubber has on a rocket test. Two examples:
1. Thrust vector control tests - The rocket nozzle is moved
side to side various degrees of arc and it must not touch the
scrubber. The scrubber entrance would also have to be extra
large to capture the exhaust when the nozzle is canted to one
side. 2. Vertical upward exhaust tests - This configuration
would require precise control of the quench and scrubbing
liquid shut-down so that parts of the rocket would not be
wetted which could cause damage. During start-up the nozzle
would have to be sealed to keep the igniter and propellant
from getting wet.
119
-------
2.0
1.8
1.6
1.4
Optimum (Pa=PP)
1.2
Shifted due to 50% reduction in
exit area and ambient pressure
1.0
10 20
NOZZLE EXPANSION RATIO, e
Figure 43. Rocket thrust coefficient.
50
120
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Design Details
So far only the overall process has been considered. In this subsec-
tion the various details of the design will be discussed. Follow-
ing this discussion of design details the entrainment separator
will be described.
Total Quencher/Scrubber Length -
Garrett, et al. (1972) and the experimental results presen-
ted in Section 4 suggest that a total length to diameter ratio
of 10 is adequate. This is subject to the constraint that the
length of the section aft of the scrubber liquid injection point
has to be 3 meters (10 ft). A conservative design would be to
make the L/D for the quench section 10 and then add a 3 meter
spray scrubber section.
. In order to keep the length to reasonable values the system
should be made of a bundle of small diameter round or rectangular-
sections as shown in Figure 44.
Quench Liquid Injection -
The radial injector designed by Garrett, et al. (1972) for
the pilot scale scrubber was not adequate, as discussed in Sec-
tion 4. The tips of the injectors burned off and the liquid did
not break up the supersonic core. The simple pipe with an angle
iron protector as used in the tests in the AFRPL pilot scrubber
tests did prove adequate. It is fully described in Section 4.
The one improvement to this design would be to coat the angle
iron protectors with a trowelable or castable insulation which
would increase their life.
Evaporating Pond -
The waste liquid flows by gravity to an evaporating pond.
The pond is lined with PVC to prevent seepage of the salt into
the water table. The design is based on a 10-year storage of dry
material and sufficient area for evaporation. The evaporation
rate at AFRPL is about 2.96 m/year and the rainfall averages
about 10.4 cm/year. The surface area is based on the volume
of liquid waste generated during a test divided by the evaporation
121
-------
Liquid Injection
at Inlet to Duct
Figure 44. Sketch of gas-atomized scrubber section.
122
-------
occurring during the period between tests. The depth is based
on the amount of solids expected to build up during the life of
the pond or between clean-out periods.
Water Storage Tank -
The tank should be sized to hold twice the amount of water
required during a test to allow for misfirings, pump malfunction,
and other contingencies. For the 2 MN rocket the volume should
be about 1,390 m3 (49,100 ft3=367,000 gal).
Piping to Scrubber -
Because of the low frequency of use and the dry climate conven-
tional carbon steel pipe can be used. The pipe should be sized
to minimize the line pressure loss, which will depend on the
piping length. A few expansion joints will be needed because
of the temperature variation in the desert climate.
Piping of Scrubber Waste Liquid -
Because of the solids content and lower available pressure
head precast concrete pipe should be used. In many applications
an open concrete culvert will be adequate.
Caustic Tank and Mixer -
Depending on topography the caustic tank may or may not be
elevated. The size is relatively small and will not significant-
ly effect overall costs whether or not it is elevated. The tank
should be sized to hold a 30% by weight solution or suspension of the
required base. It should be stainless steel. A pump (or multiple pumps)
will be required to inject the concentrated base into the mixer
with the fresh water. The mixer need not be powered but should
have baffles to create turbulent mixing. The caustic pump will
create most of the mixing action.
Materials of Construction -
Because of the low frequency of testing, especially for the
large size rockets, carbon steel is adequate for the quencher,
scrubber, and entrainment separator. The material should be 0.64
cm (0.25 in) thick to allow for corrosion. If the use is relative-
ly frequent stainless steel could be justified.
123
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Power Requirements -
Very minimal power is required. Unlike other pollutant
sources, all the gas moving power is supplied by the source it-
self. Also, with the use of hilltop or elevated water tanks the
liquid is moved by gravity. The water can be pumped to the
tanks by natural underground pressure or by low-power small
pumps because of the length of time between tests. The major
power user is the caustic pump which will use about 0.2 watts
per newton of rocket thrust (0.001 HP/lbf). While this is a
large amount of power for the large rockets, its duration is so
short (1-2 minutes) that the energy usage is slight. Other power
users in the system would be the power actuated valves.
Sampling and Analysis -
The discharges from the rocket scrubber have to be monitored
in order to determine the amount of pollutants escaping into the
atmosphere. Elaborate and expensive instruments for continuous
monitoring are either available or can be adapted for use in the
rocket scrubber application. These instruments include gas chroma-
tographs, IR and UV spectrophotometers, coulometers and colori-
meters. The need for continuous monitoring, however, is not
great because the duration of operation is usually only about one
minute. Also the operation will be steady and the rocket exhaust
is at a constant mass rate for over 90% of the burning time. Thus,
because of the short duration and the steady operation, less ex-
pensive average sampling can be used. A description of the sug-
gested sampling and analysis methods follows.
Gaseous Emissions - The gaseous pollutants are hydrogen
chloride, hydrogen fluoride, carbon monoxide, and possibly,
nitrogen oxides.
HCL - Collect by fritted glass absorbers or impingers and
analyze solution using a chloride specific ion elec-
trode.
HF - Collect by fritted polypropelene absorbers or impin-
gers and analyze solution using the fluoride specific
ion electrode method (see EPA Method 13 B).
124
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CO - Collect a gas sample and analyze by non-dispersive
infrared spectrometry (see EPA Method 10).
N0x - Collect a gas sample in an evacuated flask containing
an abosrbing solution. The amount is then determined
by colorimetric analysis (see EPA Method 7).
If a mass spectrometer is available the gas could be simply
"grab" sampled and taken to the laboratory for analysis.
Water Mist - All the pollutants produced by the rocket (HC1,
HF, CO, N0x, fluoride salts and A1203 particles) may be found in
the water droplets escaping from the mist eliminator. The chem-
ical composition of the mist can be determined by collecting the
drops in impingers and analyzing the solution by chemical and
specific ion electrode methods.
Solid Particles - The mass loading of the solid particles
may be determined by collection on high efficiency filters (see
EPA Method 5) and refer to Section 4 for a typical setup. The
size distribution is determined by using cascade impactors or
multiple filters. The primary solid particulates should be A1203,
but chemical analysis should be made to determine if HC1 or HF
have been adsorbed on the AlaOs and if any fluoride salts
are present.
Waste Water - The waste water should be analyzed to serve
as a check on the results of the gas stream analysis.
Two important considerations must be kept in mind when de-
signing instrumentation for the rocket scrubber. The first is
that the sampling must be done remotely. Remote sampling requires
special flow measuring instrumentation and use of solenoid valves
and switches to turn the sampling equipment on and off at times
corresponding to the start and end of the motor operation. The
second consideration is that the flow from the scrubber exit may
be non-uniform and hot due to afterburning. Because of non-
uniformity, a number of sampling probes should be used simul-
taneously. Afterburning will probably ruin any conventional
sampling apparatus. The scrubber exit should be designed to
provide a region of uniform flow to keep the number of sampling
probes required to a minimum and reduce afterburning effects.
125
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Labor and Maintenance -
Construction labor costs are factored into the overall con-
struction costs later in the report. Maintenance of the scrubber
consists mainly of a thorough flush and clean water wash of all
system components to remove residual caustic or unneutralized HC1.
The caustic pumps, especially, should be flushed after each run.
Additional maintenance is required for instrumentation set-up and
check-out before each test.
Control and Monitoring -
Control of the entire system would be integrated with the
main rocket firing control panel as a sub-panel which would ini-
tiate the water supply system and caustic release all in a related
sequence to the rocket firing. All the valves and motor starters
would have to be power actuated so that they can be operated
by the control computer. The important flow rates, pressures,
and temperatures need to be monitored to determine if the correct
process conditions were being met.
Safety -
The use of caustic chemicals in the scrubber liquid requires
handling precautions not normally encountered by rocket testing
personnel. Standard precautions are available from the manufac-
turer and consist mainly of care in avoiding contact by splash-
ing of the liquid or by inhaling of the vapors.
The maintenance crew should take precuations when flushing
the system to avoid contact with the residuals. These residual
solids and liquids may contain unreacted caustic and acid. Flush-
ing with fresh water may also cause heat evolution and in addition
to the temperature hazard the heat could cause emission of toxic
gases, such as HC1.
Another hazard to the maintenance crew would be pockets of
toxic and combustible gases in the system. Time must be allowed
for these gases to disperse before entering the system. Depen-
ding on how well ventilated a particular portion of the system is
the crew may even need to use self-contained breathing apparatus
when first entering the system.
126
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CONVENTIONAL SCRUBBER ENTRAINMENT SEPARATION
The conventional rocket scrubber design requires an entrain-
ment separator because the carrier gas is moving at too high a
speed to permit quick settling out of the droplets by gravity.
The effect of having no entrainment separator would be the de-
position of alumina and salts for up to a kilometer aft of the
scrubber. This deposition may be tolerable under certain condi-
tion unless fluoride salts are present and plant or animal damage
may result.
In order to design an entrainment separator the gas condi-
tions, drop size distribution, drop composition, gas loading,
and allowable pressure drop must be determined.
Gas Conditions
To serve as a baseline a total injected water mass flow
rate to rocket mass flow rate (R ) of 15 is used. The rate
of air bleed (R ") is assumed to be 0.1 . The gas conditions,
using the figures presented in the quencher analysis section
are summarized in Table 19.
TABLE 19. . GAS CONDITIONS AT SCRUBBER OUTLET
R = 15 kg water/kg propellant
R = 0.1 kg air/kg propellant
3.
QG = 3,350 m3/s (large rocket)
u = 67 m/s
PG = 0.54 kg/m3
Qj = 10.2 m3/s liquid water drops*
* assuming no gravity settling or wall catch
Drop Size Distribution
Determination of the distribution of drop sizes is based
on the Nukiyama-Tanasawa relation given in equation (45) .
The result is a Sauter mean drop diameter of about, d^ =
0.016 cm. Assuming a geometric standard deviation of about,
a = 2, and assuming a log-normal distribution, the geometric
o
mass mean diameter is d, = 0.020 cm based on a relation given
dg
by Orr (1966) .
127
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Drop Composition
The drops are assumed to consist of water containing
dissolved salt and suspended alumina. The complete reaction
of HC1 with the original hydroxide solution should produce
about 4.52 kg mol/s of salt if KC1 or NaCl is produced or
2.26 kg mol/s if CaCl2 is produced for the large rocket. The
amount of alumina collected is equal to the amount that is
collected on the drops, primarily by inertial impaction. Cal-
culations of the collection efficiency using the method pre-
sented previously showed that more than 99.91 of the mass
of the exhaust alumina should be collected. For the large
rocket about 3 kgmol/s alumina would flow out in the liquid.
The total dissolved and suspended solids concentration would
then be about 5.51 by mass (for stoichiometric amount of
Na2C03).
Gas Loading
The amount of entrained liquid in the gas is calculated
using Table 19:
QL
-»- = 3.04 kg water/m3 gas
^G
or, including the dissolved and suspended solids, the liquid
loading is 3.22 kg liquid/m3 gas.
Allowable Pressure Drop
The rocket provides a tremendous amount of power to the
scrubbing process. The pressure drop for a conventional
entrainment separator is accounted for in the governing equa-
tions by raising the mixing chamber downstream pressure by
various amounts until the rocket no longer supplies enough
suction to keep the entrance pressure below local atmospheric
(91.01 k Pa). This pressure drop is available to an entrain-
ment separator. Based on Figure 35 about 0.2 atm or 20 k Pa
(3 lbf/in2) is available which is a great deal more than adequate
for any conceivable type of entrainment separator.
128
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Required Efficiency of the Entrainment Separator
The required efficiency of the entrainment separator will
depend on three factors. The first is the tolerable deposition
of water, alumina, salts, and minute amounts of unreacted acid
and hydroxide in the vicinity of the rocket scrubber. The
second factor is economic, relating to the cost of the entrain-
ment separator and to the cost saving from recycling the scrub-
ber liquor. The final factor is possible future limits on
fluoride salt emissions. Additionally, the entrainment
separator will cause an increase in the gas residence time in
the system to help ensure complete mixing and chemical reaction.
The entrainment separator also is a flow impediment which reduces
the force of "blowback" which occurs when the rocket motor
burns out. Other than cost there are two negative factors.
The separator causes enough of a pressure drop that pressure
relief must be provided for at startup; and it must be care-
fully designed so that pockets of combustible hydrogen and
carbon monoxide do not form.
Tolerable Deposition -
The deposition of alumina and salt would be tolerable if
the usage were infrequent and the environmental impact on the
surroundings neglibible. Proximity to metal surfaces which
may be corroded by the salt and to agricultural land would
make the deposition intolerable. The effects on plants are
described by Lerman (1976).
Cost -
The cost of entrainment separators will be detailed later,
but, as a general rule, they are relatively expensive. They
are more than twice as costly as the scrubber shell itself,
so their elimination from the system would greatly reduce costs.
The costs saved by recycling the scrubber liquor depend on the
local water costs and the frequency of scrubber use. A small,
frequently used system could probably benefit from recycling
the scrubber liquid. However, recycling would be limited to
129
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some extent by the buildup of chloride ion (Cl~) which would
reduce the solubility of HC1.
Future Limits on Fluoride Salts -
The U.S. Environmental Protection Agency has proposed
emissions standards for total fluorides only for primary
aluminum reduction plants and phosphate fertilizer plants
(Chaput, 1976). Also, the latest National and California
ambient air quality standards do not single out fluorides as
a significant pollutant (California ARE Bulletin, March 1976).
However, other countries do have air quality standards for
fluorides according to Stern (1971) and the harmful effects
of fluorides on vegetation (Brandt and Heck, 1968) and animals
(Stokinger and Coffin, 1968) have been documented. Thus, it
is possible that emission of fluoride salts contained in the
scrubber liquid drops such as CaF2 would have to be controlled
to a certain limit.
Types of Separators
A number of types of entrainment separators could be used
on the rocket scrubber since the allowable pressure drop is
not a problem. The flow velocity for most separators such as
mesh, packed bed, tube bank, and zig-zag baffles must be be-
low 10 meters/second in order to keep reentrainment to a mini-
mum. To reduce the flow velocity from about 100 m/s to 10 m/s
would require an increase in area of 10 times. This increase
would mean that the flow area of the entrainment separator
would have to be about 500 m for the large rocket. The in-
ternals for such a cavernous structure would be expensive as
compared to the empty volume of a cyclone separator. Addi-
tionally, mesh and packed bed separators are susceptible to
plugging due to the solids present in the liquid drops. The
alternative design would be a cyclone separator which can use
the high energy available from the rocket to its advantage.
130
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Cyclone Entrainment Separator Design
The cyclone causes the streamlines of the flow to curve
so that a centrifugal force acts on the drops. The centrifu-
gal force causes the drops to be deposited on the sides of the
cyclone and eventually drain away as liquid effluent. In order
to prevent reentrainment of liquid from the cyclone walls the
gas inlet velocity must be kept below about 40 m/s. A typical
cyclone is pictured in Figure 45.
The equations for the efficiency of cyclones are given by
Calvert, et al. (1975).
The overall efficiency of the cyclone can be predicted by
integrating the penetration equation over the drop size dis-
tribution. Calvert, et al. (1972) (Figure 8.2-4) have performed
this integration for log-normal distributions.
The dimensions of the cyclone have to allow for vortex
development and the inlet must be sized to keep the velocity
below the 40 m/s. The following equations satisfy these cri-
teria:
ab =
De =
h =
D =
c
b =
a =
QG/40
(QG/42)1/2
6 D
e
2 D
e
0.7 D
e
1.5 D^
where all dimensions are in meters. The dimensions of "a" and
"b" may be a little greater than usually encountered, so that
the calculated efficiency will be optimistic. The pressure
drop through a cyclone has been found by Calvert, et al . (1975)
to be:
AP = 4.96 x 10'6 P 2.8 (59)
131
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I J
Figure 45- Cyclone with tangential gas inlet
132
-------
where AP = pressure drop, a.tm
PG = gas density, kg/m3
QG = volumetric flow rate, m3/s (each, cyclone)
a = cyclone inlet height, m
b = cyclone inlet width, m
De = cyclone exit diameter, m
Design for Large Rocket -
Thrust F = 2 MN
QG = 3,350 m3/s
It is obvious that a number of cyclones are required
to keep the diameter within reason. Chose,
D = 8 m
c
then,
h = 24m
D = 4 m
e
a = 6 m
b = 2.8 m
QG = 40 ab = 672 m3
so, 5 cyclones are required.
The pressure drop is,
AP = 0.013 atm
The efficiency is about 99.41, which may be a slight
overestimate. However, a specific required efficiency has
not been set so that this may or may not meet future require'
ments. Higher efficiencies could be achieved by increasing
the cyclone volume.
133
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Relief Ports
Provision for pressure relief hatches at the top of the
cyclone separator is required. These relief ports should
allow the air present in the system at rocket start-up to
exhaust. The instantaneous character of the rocket ignition
will cause a compression wave to be generated which could rup-
ture the scrubber duct or cyclone if not relieved. A typical
design would be a heavy hinged gate which would open when a
certain pressure is reached and close by gravity when the opera-
ting pressure is obtained. More costly spring-held relief ports
are also feasible. The gas emitted through these relief ports
would not affect the emission of pollutants since it consists
primarily of air.
134
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CONVENTIONAL SCRUBBER COSTS
Introduction
In a previous subsection preliminary cost estimates were
made to determine the least expensive type of scrubber for rocket
exhaust cleaning. Here the costs are detailed for the selected
gas-atomized scrubber and applied to rockets of any thrust be-
tween 20 kilonewtons (4,500 Ibf) and 2 meganewtons (450,000 Ibf).
The procedure adopted in preparing the total project costs
follows Guthrie (1970). In essence the total project costs are
prepared by summing individual bare module costs. Modules in our
case are of three types, two direct and one indirect.
Modules
The direct cost modules may be classified as the chemical
processing module and the offsite facilities module. In general,
the former covers all those items which would normally appear
in a typical chemical process flow sheet such as pumps, vessels,
exchangers, etc., while the latter encompasses those items situa-
ted external to the process battery limits such as sewerage, waste
treatment, water distribution, etc. The direct costs cover the
costs for equipment, material and labor associated with the com-
plete installation in the field. The total bare module cost for
a particular item is obtained by adding to the direct module cost
an indirect module cost. The latter comprises all costs associa-
ted with engineering, office overheads, freight, taxes, etc. and
is estimated by applying a percentage figure to the direct costs.
Figure 46 summarizes the general format for a typical chemi-
cal process module. It will be noted that the individual cost
items making up the total material and labor costs are based on
applying percentage figures to the basic equipment FOB cost.
For example, the costs for piping required to tie in a piece of
equipment to the process is 32.0% of the FOB cost. The total
material cost (FOB equipment and installation material) is ob-
tained by multiplying the FOB equipment cost by a material factor
135
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1
Mat
Fob. equipment
Piping
Concrete
Steel
Instruments
Eleclrical
Insulation
, Paint
erial
Direct Direct
material. M lobor.L
(E + m)
100.0^
32.0 VV
3.9 XX
1.7 \ N
7.3 V
!H Labor
n* factor
ao ' (xO.58)
I
v
>\
Direct
cost
factor*
(i«2.20}
i
factor I I
("1.
62)
1
I
* Field installation
i AO n 1 4. I co n 1 -
1 QjC. / | * 1 jo, U [ -
1 L/M ratio If
= 0.36
(M&L)
Indirect
factor
T
Total bore moduls — — *] 295, 1 1
Figure 46 . Outline of module cost format (factors
presented are examples only).
of 1.62. Similarly, labor costs for equipment installation is
0.58 times the FOB costs. Total direct costs are obtained by sum-
ming the direct material and direct labor cost factors and applying
this to the FOB cost to obtain the total indirect module cost.
As mentioned earlier, the indirect module is obtained as
a percentage of the direct module cost. This is indicated in the
figure at 341 (indirect factor = 1.34). Summing the direct and
indirect module costs gives the total bare module costs.
It should be noted that the numbers presented for the cost
factors are examples only. These vary from module to module
depending on the type of unit process being costed.
Total costs associated with the offsite facilities are ob-
tained in essentially the same manner as those for the chemical
process. The main difference is the absence of a process equip-
ment FOB cost. In this instance the total direct costs are ob-
tained by itemizing and summing labor and material costs to ob-
tain the direct module cost. From this step on the procedure is
identical to the chemical process module.
136
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The grand total for the project was obtained by adding per-
centage figures to the summed module costs to account for con-
tingency and fee. Examples of the complete costing procedure
are presented in a latter part of this section.
Equipment Costs
A knowledge of the equipment FOB cost is fundamental to the
module approach to cost estimating. In the case of field erected
items, the total direct labor and material costs are needed. These
were taken in most part from the literature. In some instances
vendor quotes were solicited to either confirm the published
data or to obtain updated figures.
The 2.0 meganewton rocket size was chosen as the basis from
which to scale all other rocket sizes. Scaling was performed in
most cases by using literature cost exponents. The exceptions
are noted in the text which follows.
Figures 47 through 56 present the costs for the individual
process components listed in Table 20 . The data have been pre-
sented on a cost versus rocket thrust format to simplify the
total cost estimating job. Note that the figures are for total
installed costs. These include all equipment costs, direct
material and labor costs required for field installation and
project indirects. Each of these separate items may be obtained
by applying the cost factors noted on the graphs. The total
project costs may be estimated by summing the module costs and
adding figures for the contingency and fee. A figure of 18% has
been recommended by Guthrie (1970) for this purpose.
The following summarizes the derivation of the data presen-
ted in Figures 47 through 56.
A) Quench duct, atomized spray scrubber and cyclone cost,
Figures 47 through 49,
Costs for the quench duct, the atomized spray scrubber and
cyclone costs were based on assuming field assembly of shop pre-
fabricated units. Shop costs were based on material weight and
fabrication manpower as presented by Calvert et al. (1972).
137
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TABLE 20. LIST OF PROCESSING UNITS MAKING UP THE
TOTAL CONVENTIONAL SCRUBBING PROCESS
Quench duct
Spray scrubber
Cyclones (separator)
Water storage tank
Caustic pumps
Caustic tank
Pipeline (water supply)
Liquor supply pump
Drainage sewer
Drying bed (waste treatment)
138
-------
Cost exponent, labor and material factors were taken from
Guthries "norm" for chemical process modules. Indirects were
taken at 60% of the directs as opposed to the 34% "norm". This
was to allow for the non-conventional nature of these processing
items which would required higher investment.
B) Water storage tank costs, Figure 50.
Water storage tank costs have been presented for the purpose
of making cost estimates for a complete grass roots project. The
current set-up at the test base could use the existing water
storage tank located at the top of Haystack Butte and this would
eliminate this cost for the existing set-up.
The costs presented in Figure 50 were based on Guthrie
(1970) and assume field erection throughout.
C) Caustic supply pump, Figure 51.
The maximum pump size considered for the current application
was taken at 2,000 gal/min. Scrubbing units requiring flows
greater than this were assumed to use multiple pump units to meet
the required capacity. The figures presented in Figure 51 were
based on the data in Guthrie (1970) for centrifugal, motor-driven
pumps.
D) Caustic storage tank costs, Figure 52.
Caustic storage tank costs assume field erection of shop
fabricated vessels. The numbers are based on Guthrie (1970) for
API conical vessels.
E) Pipeline Costs, Figure 53.
The pipeline costs presented in Figure 53 assume the quench,
and scrubber liquor water would be gravity fed from storage tanks
situated on top of Haystack Butte. For large rockets this is
the least expensive route. However, as the rocket size is lowered
a trade-off point is reached where the cost of piping from the
hill to the scrubber (182 m) equals the cost of pumping from a
locally situated storage tank. This point was taken at 0.1
MN (22,500 Ibf) in this work and was arrived at from consideration
of pipings, pumps and power costs. A similar tradeoff would be
needed for other test sites.
139
-------
Current material costs were obtained from vendor quotes.
Field installation labor hours were taken from Guthrie (1970)
and charged at the average crew labor rate of $10 per hour. A
figure of 20.0 percent was added to the sum of these costs to
allow for valves and other piping auxiliaries such as flanges,
bends, etc. These costs were escalated 50% to allow for the
unusual field conditions found at the test area. Indirect costs
were based on Guthrie (1970) and taken at 30%.
F) Scrubbing liquor supply pumps, Figure 54.
As mentioned earlier, for rockets smaller than 0.1 MN
it may be more economical to pump the scrubbing liquor from a
locally situated storage tank. Costs for these pumps are presented
in Figure 54. The maximum pump size was chosen at 2,000 gal/min.
All costs and cost factors were based on Guthrie (1970) for motor
driven centrifugal pumps.
G) Drainage sewer costs and drying bed costs, Figures 55 and
56.
Drainage sewer costs and drying bed costs were based on in-
formation taken from Lee Saylor (1976). Sewer costs presented
in Figure 55 were prepared from estimates made for a number of
rocket sizes. The figures include all costs associated with
sewer construction and include pipe costs, excavation, shoring,
backfill, etc. Drying bed costs shown in Figure 56 include con-
struction costs for excavation, grading, compacting, etc. Linear
costs were obtained from vendor quotes. A figure of 40% was added
for indirects per Guthrie (1970) .
Total Installed Costs
Figure 57 shows a summation of all the cost modules for
total cost versus rocket thrust relationship. Two lines are
shown to differentiate between gravity feed through a 182 m
pipeline from the hilltop or pumping from a local tank. The
breakpoint of 0.1 MN thrust is somewhat subjective and based on
the point at which operating (power) costs become significant. The
contingency and fee factor of 18% is included in the figure.
140
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Scrubber Costs - Worked Examples
Two worked examples of cost estimates made for conven-
tional gas atomized scrubbers are presented in Table 21.
These are for three levels of rocket thrusts. It can readi-
ly be appreciated that the costing procedure is straight
forward and needs little by way of explanation. In essence,
total costs are estimated by summing the installed costs for
the components listed in the Table with the individual item
costs taken from Figures 47 through 56. To the sum is added
a percentage to allow for contingency and fee.
It is worth reiterating that at Haystack Butte rockets
smaller than 0.1 meganewtons show no pipeline costs since scrub-
bing liquor is supplied via pumps from locally situated storage
tanks. Conversely, rockets greater than this size show pipeline
costs. These units are assumed gravity fed from storage tanks
in a situation similar to that existing at Haystack Butte. Total
project installed costs for 2.0 meganewton (450,000 Ibf), 0.22
megamewton (50,000 Ibf), and 0.022 meganewton (5,000 Ibf) rockets
are $4.26 million, $1.16 million, and $208,000 respectively, based
on December, 1976 dollars.
141
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TABLE 21. WORKED EXAMPLES OF CONVENTIONAL GAS
ATOMIZED SPRAY SCRUBBER
Rocket thrust, meganewtons 2.0 0.22 0.022
Rocket thrust, pounds-force 450,000 50,000 5,000
Cost Item Cost $1,000
Quencher 210 54 13
Spray Scrubber 280 74 23
Cyclones 1,400 370 92
Water Tank 112 28 7
Caustic Pump 140 13 2
Cuastic Tank 23 11 5
Pipeline (water supply) 1,220 400
Liquor Supply Pump -- -- 23
Drainage Sewer 155 27 10
Drying Bed 72 8_ 1
Installed Cost $3,612 $985 $176
Contingency + Fee @ 18% 650 177 52
Total Installed Cost $4,262 $1,162 $208
142
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ROCKET THRUST - POUNDS X 10"
20 50 100 200
500
100
H
W
O
asis: Dec.'76, M§S =
Field Installation
Material Factor - 1.62
Labor Factor =0.58
Indirects = 0.6
Cost Exponent = 0.6
Material: Carbon Steel
mm '
ROCKET THRUST - POUNDS x 10
10 20 50 100 200
500
1,000
asis: Dec.'76, M§S =
Field Fabrication
Material Factor = 1.62
Labor Factor = 0.58
Indirects = 0.60
Cost Exponent =• 0.6
Material: Carbon Steet
100
Figure 47.
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Quench duct installed costs (figures include shop and
field materials, labor and indirects).
0.02
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Figure 48. Atomized spray scrubber costs (figures include field
materials, labor and indirects).
-------
ROCKET THRUST - POUNDS x 10
20 50 100 200
500
1,000
o
u
Basis: Dec.'76,
Field Fabrication
Material Factor = 1.62
Labor Factor » 0.58
Indirects = 0.60
Cost Exponent = 0.6
Material: Carbon Steel
ROCKET THRUST - POUNDS x 10"
20 50 100 200
500
100
100 ! = =M
.01
| Basis: Dec.'76,
Indirects = 0.36
Cost Exponent = 0.63
is Material: Carbon steel
0.1
ROCKET THRUST - MEGANEWTONS
1.0
Figure 49. Cyclone costs (figures include field materials, labor
and indirects).
0.01 0.1 1.0
ROCKET THRUST - MEGANEWTONS
figure SO. Water storage tank costs.
-------
ROCKET THRUST - POUNDS x 10 3
10 20 50 100 200
500
100
H
W
o
Basis: Dec.'76, M§S - 480
Material Factor =1.72
Labor Factor = 0.70
Indirects = 0.40
Material: Carbon Steel
ROCKET THRUST - POUNDS X 10"
10 20 50 100
200
500
Shop Fabrication
Material Factor = 1
Labor Factor = 0.34
Indirects =0.33
Cost Exponent =0.30
Material: Carbon Steel
0.1
ROCKET THRUST - MEGANEWTONS
1.0
0.01
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Figure 52. Caustic storage tank costs.
- API conical
(Figures include shop and field materials,
labor and indirects).
Figure 5.1. Caustic slurry supply pump costs.
-------
ROCKET THRUST - POUNDS x 10
5 10 20 50
- 3
1,000
o
o
o
H
CO
O
100
ROCKET THRUST - POUNDS x 10"3
50 100 200
500
Basis: Dec.'76, M§S = 480i
Indirects =0.30
Line Length = 182 m
Material: Carbon Steel
Exponent =0.5
o
o
o
100
H
en
o
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Figure 53. Pipeline costs.
Basis: Dec.'76, M§S = 480
Material Factor =1.72
Labor Factor = 0.70
Indirects = 0.40
Material: Carbon Steel
0.01
1.0
ROCKET THRUST - MEGANEWTONS
Figure 54. Quench and scrubbing liquor pump costs for
small rocket scrubbers.
-------
100
ROCKET THRUST - POUNDS x 10
10 20 50 100 200
8
Basis: Dec.'76, M§S = 480
Indirects - 0.40
0.01 0.1
ROCKET THRUST - MEGANEWTONS
Figure 55. Drainage sewer costs.
1.0
ROCKET THRUST - POUNDS x 10"
10 20 50 100
10
Basis: Dec.'76, M§S = 480 ;
Indirects =0.40
Cost Exponent =1.0
0.01
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Figure 56. Drying bed costs (figures include total materials,
installation and indirect costs).
-------
2
O
-3
h-l
s
*» 1,
H
C/D
O
U
O
10
ROCKET THRUST - POUNDS x 10
20 50 100 200
_ 3
500 1,000
PUMP FEED (HIGHER OPERATING COST)
0.1
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Figure 57. Total installed costs for conventional scrubbers
(December 1976).
148
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UNCONVENTIONAL SCRUBBER DESIGN
Introduction
Although any scrubber designed for rocket exhaust cleanup
is unusual, the design previously called "conventional" is a
widely used gas-atomized spray scrubber. Because of the high
cost of such a system for infrequent use, an alternative de-
sign is proposed. The major criterion for this alternative,
or unconventional, design is to achieve the least possible
cost for a system that could reasonably be expected to ade-
quately scrub the rocket exhaust. The major emphasis is placed
on a design for the largest rocket, where the largest cost
savings could be realized.
Ideas Considered
Several ideas were investigated and rejected on either
feasibility or economic grounds. One idea was to air-drop a
neutralizing solution on the ground cloud. The National Aero-
nautics and Space Administration (NASA) was considering this
idea for launches of the Space Shuttle. For the AFRPL applica-
tion it seemed that there were too many problems with this
idea. Major among these problems was the probable scrubbing
efficiency. A slight miscalculation of the trajectory could
result in a partial miss of the exhaust cloud. Missing the
cloud could result in rain-out of caustic on the surroundings
as well as a poor scrubbing efficiency. It was also felt that
since NASA was already pursuing this idea that another line
should be followed in this study.
Another idea was to construct a network of spray manifolds,
extending out into the desert, that would encompass the rocket
exhaust plume. As envisioned this concept would use high pres-
sure spray nozzles to distribute the neutralizing liquid through-
out the plume, with no walls required. This idea was not that
inexpensive because of the length of pressurized piping needed.
Also, it was susceptible to inefficiency, whenever the wind
would blow the plume out of the spray network.
149
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Selected Design
A.P.T. has conceived an unconventional design which takes
advantage of the topography of Haystack Butte at AFRPL to
create an open spray scrubber. The scrubbing liquid is gravity
fed through open culverts to the open scrubber channel to avoid
the high cost of the piping. A wall is used on the downhill
side to reduce wind effects and contain the exhaust plume and
spray. Figure 58 presents a schematic of the concept while a
plan view sketch of the proposed system is given in Figure 59.
The rocket exhaust is first led through the deflector to
the transition piece which consists of a simple section of
ductwork constructed from materials capable of withstanding the
high temperatures of the rocket gas. The purpose of the transi-
tion piece is to take the gas from the deflector to the beginning
of the quenching and scrubbing excavation. The cross-sectional
area of the transition piece was taken at 9.0 square meters and
covers a distance of about 35 meters.
Cooling of the hot gas and mass transfer of the contami-
nants to the spray liquor takes place in the cooling and scrub-
bing section. It is envisioned this would be constructed by
excavating the hillside. One side and the bottom
of the scrubber cross section are stabilized with gunite. The
remaining side is made from corrugated sheet steel suitably
supported while the top remains open. Figure 60 presents a view
of a typical cross section illustrating these principles. The
cross-sectional area of the excavation expands from an inlet of
25 square meters to an outlet of 600 square meters over a length
of approximately 50 meters. The exit dimensions were chosen
to provide a gas velocity of 5 m/s for entry to the entrainment
separator.
To facilitate slowing down the gas, horizontal and vertical
baffles are attached to the base and side of the excavation, the
aim being to remove momentum through drag. These are illustra-
ted in the sketch of Figure 60.
150
-------
Water
Tank
Deflector
Rocket
Exhaust
Caustic
Tank
Transition
Piece
Quencher +
Scrubber
Excavation
Entrainment
Separator
V
t
7
Drying Pond
Figure 58. Schematic of non-conventional rocket exhaust gas scrubbing process
-------
Deflector
Elevation
Contour Lines
Transition
Piece
Scrubbing
Section
Water Supply from
Storage Tank
50
100
Scale - 1" = 50
Figure 59 .Plan view of non-conventional scrubbing system,
152
-------
Retaining
Wall
Supply
Aqueduct
Figure 60. Typical cross-section of a non-conventional
absorber configuration.
153
-------
Water required for* cooling and scrubbing is gravity fed
via an open culvert from the water storage tank on Haystack
Butte. Base required for neutralization is also gravity fed
at a point close to the start of the water run. The scrubber
liquid is then discharged into an open spray manifold designed
to distribute the liquor along the length of the excavation.
As the liquor traverses the gas plume, momentum will be lost to
the liquor thus helping to slow down the gas. Simultaneously,
the liquor will be atomized providing the surface area required
for mass transfer from the gas to the liquid and the drop for
collecting particles. Some of the liquid drops fall from the
plume over the trajectory length while the remainder are removed
in the entrainment separator located at the end of the scrubber
excavation.
The entrainment separator is built across the exit of the
excavation and is envisioned as being a tube bank type. The
scrubbing liquor removed in the separator combines with the
liquor fallout from the gas plume and flows under gravity in
an open-conduit to the waste treatment drying bed.
Process Design
Since the source process is the same for the unconventional
as well as the conventional and both use a spray design, the
process parameters are similar.
Equilibrium Conditions -
As with the conventional design the quencher and scrubber
can be considered one unit. However, the equilibrium cross
section is a function of the scrubbing liquid rate since the
scrubber is open to the atmosphere. In the subsection on quen-
ching the area, gas density, temperature, and velocity were
determined as functions of the liquid to rocket mass flow rate
ratio for the quencher. These curves apply to the scrubber as
well.
The equilibrium mass transfer is practically the same as
for the ducted spray scrubber. The previous analysis showed
that the total system (quencher and scrubber) liquid flow rate
154
-------
should be 15 kg water per kg propellant for adequate mass trans-
fer. Based on Figure 38 this corresponds to the following condi
tions:
TABLE 22. EQUILIBRIUM CONDITIONS IN UNCONVENTIONAL SCRUBBER
WITHOUT BAFFLES
RW = 15 kg water/kg propellant
R& = 0.1 kg air/kg propellant
A = 20 m2
PG = 0.54 kg/m3
T = 92°C
u = 160 m/s
Q = uA = 3,200 m3/s
The velocity, u = 160 m/s is too high. A lower velocity is
required so that large drops will settle out in the scrubber
and a low pressure drop entrainment separator can be used. This
lower velocity can be accomplished by using baffles to expand
the plume to a larger cross section (A). The baffles will also
create drag which will reduce the velocity. For the large
amount of water input the effect of velocity on the equilibrium
temperature will be slight so that the volume flow rate will
not be affected by the baffles. The baffles will also serve
to knock out the scrubber liquid drops and create more turbulent
mixing.
The mass transfer length has been shown to be on the order
of a few meters for the spray scrubber, which is much less than
the designed 50 meter quencher/scrubber length.
Coupling Effects Between Scrubber and Rocket
At Haystack Butte the rocket could be affected by the
thrust deflector in the same manner as the quencher inlet of the
155
-------
conventional scrubber described previously. The unconventional
scrubber section itself is located where it will not have any
effect on the rocket.
Thrust Deflector
Because of the vertical configuration of the test rocket
at Haystack Butte a deflector is required to direct the exhaust
into the scrubber. Design and cost information should be available
from the companies who constructed the thrust deflectors used for
the National Aeronautics and Space Administration very high thrust
rocket tests at AFRPL. Those NASA thrust deflectors were capable
of deflecting a downward thrust of 6.7 meganewtons (1.5 x 106
Ibf) to the horizontal.
156
-------
UNCONVENTIONAL SCRUBBER ENTRAINMENT SEPARATION
Introduction
The primary function of the rocket exhaust scrubber is to
remove the halogen acid gases produced. If a basic scrubbing
solution is used and the reaction with the halogen acid within
the scrubbing drops produces a non-toxic salt then there is no
need to remove the drops. In the case of rocket exhausts con-
taining hydrogen fluoride the production of a truly non-toxic
salt in the scrubbing drops may be practically impossible. How-
ever, as of October, 1976, there is no specific rule in Califor-
nia concerning the emission of fluoride salts.
Another argument against the need for an entrainment separa
tor on the open, unconventional scrubber design is based on the
rapid slowdown of the gas after leaving the scrubber. The gas
is not confined to a duct, but is completely free to expand and
diffuse, and will lose its velocity in a short distance. The
drops which have more momentum will carry a little farther but
they will also quickly slow down. Although the detailed cal-
culation for this two-phase wake flow have not been made, the
argument appears to be valid.
There are at least two reasons why an entrainment separator
might be required. First, the local air pollution control dis-
trict may tighten the restrictions. Specific rules on fluoride
salts, for example, may be forthcoming. Secondly, other reasons
may exist which would require the entrainment to be removed
immediately following the scrubber. Proximity to other struc-
tures or sensitive areas where salt spray would cause problems
are reasons for having an entrainment separator.
Entrainment Separation Within Scrubber Section
The unconventional scrubber is designed to have the scrub-
bing liquid carried into it at several locations along its 50
meter length. As the gas traverses the trajectory path it will
157
-------
lose momentum to the scrubbing liquor resulting in the atomiza-
tion of droplets of increasing size. The larger droplets atomized
at the low gas velocities will collect the smaller droplets
formed at the high velocities through inertial impaction. This
results in an overall larger mean droplet size for final separa-
tion. A study of the collection of small droplets via impaction
and definition of a final droplet size for ultimate separation
was performed.
Drop Size From Critical Weber Number -
According to Calvert (1968), the stable droplet size for
water drops in gas streams may be predicted from:
We = pG (UG - ud)2 -^ x $ (60)
2a
where We = Weber number, dimensionless
p = gas density, g/cm3
o
Uj = drop velocity, dm/s
u~ = gas velocity, cm/s
a = surface tension, dynes/cm
d-, = drop diameter, cm
Hidy (1970) suggests a critical Weber number = 6.5.
The maximum stable drop diameter will be predicted when
the term (UG - u^) tends to zero. However, for the present
case the maximum will be predicted when either "un" or "u," is
u d
zero. For this condition "u-," equals the terminal settling
velocity, u. , and for drops greater than 0.15 cm in diameter this
may be expressed:
(\ 0.5
g dd pd\
—^-^ , cm/s (61)
PG /
where g = gravitational constant
p, = drop density, g/cm3
158
-------
Equating (60 ) and (61 ) shows that a stable drop diameter
of about 0.5 cm will be formed at gas velocities less than 16.0 m/s
This drop size corresponds to the terminal settling velocity
of the drop and will not change regardless of gas velocities be-
low this figure. Put in other words, slowing the gas below
16.0 m/s will not affect drop size.
Drop Size from Nukiyama-Tanasawa Relation-
The Nukiyama and Tanasawa equation discussed previously
may be used to predict atomized drop size at the high velocities:
(62)
where d-. = the Sauter mean diameter, cm
UG = gas velocity, cm/s
and (QT/Qr) is dimensionless
The Sauter mean diameter, d, is related to the geo-
metric mean diameter as follows:
In dds = In dd - 0.5 In2 a (63)
where a = geometric standard deviation
o
The standard deviation for gas atomized drops is generally
on the order of 2.0. With this information, drop diameter at the
start of the scrubbing section, where "UG" = 500 m/s, was compu-
ted to be at least 10 ym.
Collection of Small Drops on Large Drops -
The worst condition for design will be to capture those
10 urn drops atomized at the start of the scrubbing section.
The most pessimistic viewpoint would be to consider no drop
growth through impaction as the gas traverses the trajectory.
159
-------
Drop collection would take place therefore only in the last
stage of the scrubbing section,on the large drops (0.525 cm)
formed at the low gas velocities.
The mechanism for drop collection by impaction is described
in Calvert, et al. (1972) and Yung, et al. (1976). The collection
efficiency is primarily a function of collected drop size, liquid
to gas ratio, collector drop size, and relative velocity between
collected and collector drops. Assuming a relative velocity of
10 m/s, a collected drop diameter of 10 m and a collector drop
diameter of 0.525 cm the fractional efficiency of collection is:
E ~1 - exp (-1.2 QL) (64)
where "Qr" refers to the volume flow rate (m3/s) of the collector
drops. The total liquid input to the scrubber section is about 11
times the mass flow rate of the rocket propellant (R = 11). The
collection efficiency would be 601 if the R ratio of the collec-
w
tor drops were 1 and greater than 99% if the R ratio were 5.
Since this is a worst case analysis in that the collected drops
would actually be larger than 10 ym and the collector drops would
actually be smaller than 0.525 cm, the conclusion is that a
high percentage of the mass of the entrainment will be around
0.525 cm diameter.
Separation by Gravity Settling
Entrainment separation may be accomplished by gravity
settling because the drop size predicted by the critical Weber
number concept is sufficiently large.
The removal efficiency of drops falling under gravity in
turbulent flow may be expressed by the equation of Calvert, et
al (1975):
' -u L
E = 1 - exp |~^—} (65)
H ur
(3
160
-------
where u = the terminal settling velocity of the drop, m/s
L = horizontal distance travelled by the drop, m
u = gas velocity, m/s
o
H = total height fallen drop
For a gas plume of square cross section and uniform velo-
city the flow rate is,
QG = ufi H2 , m3/s (66)
Substituting in equation (65) gives:
-u L
E = 1 - exp ' rs
According to Fuchs (1964) , the terminal settling velocity
of 0.525 cm diameter water drops is u =9 m/s. For Q = 3,200
m3/s and u., = 5 m/s the removal efficiency of these drops would
be 99% for a length of L = 65m. Actually, because of the flow
resistance of the air this length would be shorter. Smaller
drops would, of course, require a longer settling length.
Because settling areas with lengths on the order of 65 m or
longer may not be available at the site of the scrubber, separa-
tion by other means should be investigated.
Unconventional Entrainment Separator Design -
There are three inexpensive, low pressure drop designs that
are applicable to the unconventional scrubber mist. One is
a gravity settler which uses a parallel array of horizontal
plates so that the parameter, H, in equation (66) is reduced.
Another type uses a series of horizontal baffles and a third
uses a bank of vertical tubes. The costs of these designs are
comparable, so the choice must be based on other grounds. Accor
ding to design equations given by Calvert, et al. (1975) the
tube bank should be slightly more efficient than the other two
161
-------
and they present some data showing a slightly higher reentrain-
ment velocity for the tube bank design.
Tube Bank Entrainment Separator - Figure 61 shows in
schematic form the arrangement within a tube bank separator. In
principle, liquid drops are separated from the gas stream by
impaction on the tubes. The liquor is then removed from the pro-
cess by gravity flow along the tubes to a suitable collection
manifold. Calvert et al. (1975) present a mathematical model
for tube bank design. Their models were used to determine the
collection efficiency of a tube bank consisting Of 3.34 cm
(1.3 in) tubes at 2.54 cm spacing. The models show a single
stage sufficient for almost 100% collection. Again this is due
to the large drop size being collected. For practical purposes,
a 3-stage unit would be the minimum for consideration and this
has been used in the cost estimates.
Calvert et al. (1975) have also studied reentrainment in
tube banks at the pilot plant scale. Their findings show no
reentrainment for velocities below about 7 m/s. Since our de-
sign velocity is 5 m/s no reentrainment is expected.
The pressure drop in tube banks at 5 m/s is quite small.
Calvert et al. (1975) found it to be on the order of 100 Pa
(1 cm W.C.).
162
-------
Figure 61. Sketch showing arrangement of tube bank
separator.
163
-------
UNCONVENTIONAL SCRUBBER COSTS
Introduction
Cost estimates for the open, unconventional scrubber are
made using the same procedure previously described for the
conventional scrubber. The procedure follows that of Guthrie
(1970) which uses the sum of the individual module costs.
The cost modules are biased to the specific location of the
test area at AFRPL's Haystack Butte and the design for the large
rocket. The cost modules concerning the transition piece, the
scrubber excavation, and the scrubber excavation retaining walls,
etc. would be different than presented for other locations. The
costs for smaller thrust rockets are based on a scale-down from
the large, 2 meganewton thrust, rocket. The emphasis has been
placed on the larger rockets because the cost benefit over the
conventional design is much greater for the larger sizes.
Equipment Costs
Fundamental to the module approach to cost estimating is a
knowledge of the equipment FOB cost or, in the case of field
erected items, the total direct labor and material costs. These
were taken in most part from the literature. In some instances
vendor quotes were solicited to either confirm the published
data or to obtain updated figures.
The 2.0 meganewton rocket size was chosen as the basis from
which to scale all other rocket sizes. Scaling was performed in
most cases by using literature cost exponents. The exceptions
are noted in the text which follows.
Figures 62 through 70 present the costs for the individual
process components listed in Table 23. The data have been pre-
sented on a cost versus rocket thrust format to simplify the
total cost estimating job. Note that the figures are for total
installed costs. These include all equipment costs, direct
material and labor costs required for field installation and
project indirects. Each of these separate items may be obtained
by applying the cost factors noted on the graphs. The total
164
-------
project costs may be estimated by summing the module costs and
adding figures for the contingency and fee. A figure of 18% has
been recommended by Guthrie (1970) for this purpose.
TABLE 23. LIST OF COMPONENTS MAKING UP THE TOTAL
UNCONVENTIONAL SCRUBBING PROCESS - LARGE ROCKETS
Transition piece
Scrubbing liquor viaduct and manifold
Scrubber exvacation, retaining wall deflectors
and drag elements
Entrainment separator
Water storage tank *
Caustic storage tank *
Drainage sewer *
Drying bed *
* Common to both conventional and unconventional scrubbers.
The following summarizes the derivation of the data pre-
sented in Figures 62 through 70 . The cost basis for some items
has previously been presented and are not repeated.
A) Transition Piece and Separator Costs, figures 62 and 65.
Costs for the transition piece and separator were based on
assuming field assembly of shop prefabricated units. Shop costs
were based on material weight and fabrication manpower as pre-
sented in Calvert, et al. (1972). Cost exponent, labor, and
material factors were taken from Guthrie's "norm" for chemical
process modules. Indirects were taken at GuthrieTs "norm" of
0.34 for the transition piece and separator. The indirect fac-
tor for the scrubber was taken at 0.60 to allow for the non-con-
ventional nature of this item.
165
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B) Scrubber Excavation - figure 63.
Scrubber excavation costs were taken from Lee Saylor (1976) .
The figures include costs for grading, leveling, and backfill.
It was assumed 20% of the excavation would require blasting with
the remainder requiring air tool. Indirects were taken at 0.34.
C) Scrubber Excavation Retaining Walls, Gunite, Drag and
Deflector Units, figure 64.
Figure 64 presents the combined costs for the miscellaneous
items making up the excavated scrubber body. Gunite is used on
the excavated rock portion of the scrubber. Essentially this is
confined to the wall on the uphill side of the scrubber and
the base. Costs were based on information taken from Lee Saylor
(1976). The purpose of the retaining wall is to make up that
part of the scrubber walls not formed by the excavation. In most
part, this is confined to the downhill side of the excavation.
The wall supports to some extent act as drag and deflector units.
Since their costing procedure was similar to that for the retain-
ing wall, their costs were lumped together. Material costs were
based on cement material quotes from vendors. Total material
and labor were taken at five times the material cost in constrast
to the normal 3.5 to 4.0 in order to account for the unusual
field conditions. Total indirects were taken at 0.34 per Guthrie's
"norm".
D) Scrubbing Liquor Viaduct and Manifold, figure 66.
Costs were derived from a combined material and labor cost
based on material weight as presented in Calvert, et al. (1972).
Indirects were taken at 0.34 from Guthrie's "norm".
Total Installed Costs
Figure 71 presents a summation of all the cost modules as
a function of rocket size. The 18% contingency and fee factor
is included. The unconventional design costs are location speci-
fic because the use of the typography in the design. The conven-
tional design is slightly location specific because of the depen-
dence on the hilltop location for the water tank and the length
166
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of the pipeline to the test pad. Shown for comparison is the con-
ventional scrubber costs. For the 2 meganewton (450,000 Ibf)
thrust rocket the cost difference is about $2.5 million.
Scrubber Costs - Worked Examples
Three worked examples of cost estimates made for unconven-
tional gas atomized scrubbers are presented in Table 24. It can
readily be appreciated that the costing procedure is straight-
forward and needs little by way of explanation. In essence, total
costs are estimated by summing the installed costs for the com-
ponents listed in the table with the individual item costs taken
from Figures 62 through 70. To the sum is added a percentage to
allow for contingency and fee.
167
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TABLE 24. WORKED EXAMPLES OF UNCONVENTIONAL
SCRUBBER COSTS
Rocket Thrust, meganewtons 2.0 0.22, 0.022
Rocket Thrust, pounds-force 450,000 50,000 5,000
Cost Item Cost $1,000
Thrust deflector* 100 11 1
Transition piece 76 20 5
Scrubber excavation 600 66 7
Excavation retaining walls, etc. 73 24 8
Entrainment separator 220 25 2
Scrubbing liquid viaduct 88 30 10
Water storage tank 112 28 7
Caustic storage tank 23 11 5
Drainage sewer 155 27 10
Drying bed 72 8_ _!_
Installed Cost- $ 1,519 $ 250 $ 56
Contingency + Fee @ 18%- 275 45 1£
Total Installed Cost- $ 1,792 $ 295 $ 66
*Rough estimate
168
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100
ROCKET THRUST - POUNDS x 10
10 20 50_ 100
200
500
10
H
CO
O
Basis: Dec.'76, M^S = 480
Shop and Field Fabrication
Indirects =0.34
Cost Exponent = 0.60
Material: Carbon Steel
0.01
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Figure 62. Transition piece costs.
1,000
ROCKET THRUST - POUNDS x 10"
10 20 50 100 200
100
Basis: Dec.'76,
Indirects = 0.34
Exponent = 1.0
0.1 l
ROCKET THRUST - MEGANEWTONS
Figure 63. Scrubber excavation costs.
-------
ROCKET THRUST - POUNDS X 10
10 20 SO 100
200
100
ROCKET THRUST - POUNDS x 1
-------
ROCKET THRUST - POUNDS x 10"
10 20 50 100 200 500
100
o
o
o
H
03
O
u
10
Basis: Dec.'76, MIS = 480
Indirects = 0.30
Material: Carbon Steel
Exponent = 0.5
0.01
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Figure 66. Scrubbing liquor viaduct and manifold costs.
ROCKET THRUST - POUNDS x 10""
10 20 50 100 200 500
100
o
o
o
H
co
o
Basis: Dec.'76, M§S = 480
Indirects =0.36
Cost? Exponent = 0.63
Materials: Carbon steel
0.01
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Figure 67. Water storage tank costs (figures include total
materials, field installation and indirects).
171
-------
ROCKET THRUST - POUNDS x 10"
10 20 50 100 200
500
Basis: Dec.'76, M£S - 480
Shop Fabrication
Material Factor = 1.13
Labor Factor = 0.34
Indirects =0.33
:3 Cost Exponent = 0.30
Material: Carbon Steel
.01
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Figure 68. Caustic storage tank costs.
100
ROCKET THRUST - POUNDS x 10"
10 20 50 100 200
o
o
o
0.01
0.1
ROCKET THRUST - MEGANEWTONS
Figure 69. Drainage sewer costs.
1.0
172
-------
100
ROCKET THRUST - POUNDS x 10"
10 20 50 100
200
500
O-l
Basis: Dec.'76, M§S = 480
Indirects = 0.40
Cost Exponent
0.01
0.1 1.0
ROCKET THRUST - MEGANEWTONS
Figure 70. Drying bed costs.
ROCKET THRUST - POUNDS x 10"
20 50 100 200
500
1000
0.1
0.01
Based on location
at Haystack Butte
Includes water tank
cost
0.01 0.1 i 5
ROCKET THRUST - MEGANEWTONS
Figure 71. Total installed costs of conventional and unconventional
scrubbers.
-------
SCRUBBER OPERATION COSTS
The operation costs of the scrubber, whether the closed or
open design, consist of four parts:
1. Labor
a. Direct
b. Supervision
c. Overhead
2. Utilities - primarily electricity and water
3. Expended chemicals
4. Replacement parts
Labor
The detailed labor requirement for the large (450,000 Ibf)
thrust rocket scrubber will be considered. A test schedule of
one test per month will be used. The following table (25)
presents the scrubber operation plan.
TABLE 25. SCRUBBER OPERATION PLAN
Task No. Description Time
I. Preparation 2 weeks preceeding run
A. Maintain equipment
B. Check water flow
C. Mix basic solution
D. Fill liquid tanks
E. Check instrumentation
II. Run Part of 1 day
III. Clean-up 2 weeks after tun
A. Clean out scrubber
B. Flush base from piping
C. Check drying pond
174
-------
Scrubber maintenance and operation should require a full-time
operator and a part-time helper. For a large scrubber the helper
may be needed about half-time. The cost of this labor for the
1 1/2 men would be about $46,800 per year using an hourly rate
of $15.00 which includes supervision and overhead.
Other Costs
The utility cost would be very small since the water comes
from wells and the only power is that required to drive pumps
to slowly fill the water tank. The expended chemicals for 12
large rocket tests in a year would be about $25,600 per year
based on Table 17.
It is difficult to estimate the cost of replacement parts
for the scrubber. The possible parts that would have to be
replaced would be some of the liquid injector nozzles and parts
of the quencher which may be overheated.
Total Yearly Operational Costs
Based on the previous discussion the total yearly operational
costs are summarized as follows for the large rocket:
Labor $46,800
Utilities 5,000
Chemicals 25,600
Replacement Parts 10,000
Total $87,400 per year
175
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SECTION 7
CONCLUSIONS AND RECOMMENDATIONS
PURPOSE
The purpose of the effort described has been to design a
scrubbing system for test rocket exhausts. This design was to
be feasible in terms of scrubbing efficiency and cost. While
the designs presented should provide for the required scrubbing
efficiency at the lowest possible cost, the expense may still
be too great.
DESIGN
The exhaust product of concern was gaseous hydrogen chlor-
ide, which we estimated had to be removed with 99.6% efficiency.
The scrubber chosen was the gas-atomized spray of which the
familiar venturi scrubber is a type. The energy required to
achieve the high efficiency is supplied by the rocket exhaust.
The length required for the mass transfer was determined to be
about 3 meters after the exhaust had been slowed to about 100
m/s and cooled to below 100°C in a spray quencher. This
mass transfer length is possible without the scrubbing liquid
being basic because of the high solubility of HC1 in water.
The scrubbing liquid does have to be neutralized before it is
released from the system, however. The amount of water required
for quenching and scrubbing was about 15 times the mass flow
rate of the rocket exhaust.
Early in the design effort it was determined that even
this design would be expensive for large rockets. Thus, an al-
ternative unconventional design is also proposed. It too, is
a gas-atomized spray scrubber which uses the special topography
176
-------
available at Haystack Butte at AFRPL. To save structure costs
the quench and scrubber section can be excavated from the side
of the hill and left open on top. Piping costs are reduced by
using open culverts which deliver the water to the top of the
open scrubber. The operation and efficiency of the open gas-
atomized scrubber have not been demonstrated.
An open gas-atomized scrubber can be built on flat terrain
by either digging a ditch or building side walls to guide the
flow. Because the scrubber operation is so intermittent, it is
advantageous to use a large water storage tank and a small pump
which fills the tank over a long period of time. The alterna-
tive of using a large pump which can handle the entire scrubber
flow rate would involve a much greater equipment cost and the
provision of a large power supply.
A major cost item of either design is the entrainment
separator, which may be superfluous in some installations. The
conventional design uses cyclones and the unconventional uses a
tube bank. The requirement for entrainment separators must be
evaluated at the time installation is considered. While they do
help insure adequate mixing for complete mass transfer, if a
basic scrubbing solution is used they are only removing water
drops containing a salt. Thus, since the water is not being
recycled and the salt laden drops may not be a pollution problem
entrainment separators may not be needed.
DESIGN PROBLEMS
While the proposed designs are fairly complete there are
certain aspects that require additional study. The first is
the structural integrity of the entrainment separator. Mention
has been made in the text of relief ports that would open when
the initial blast from the rocket impacted on the air present
in the system. These relief ports must operate properly or the
entrainment separator will be destroyed.
A second problem is the effect on the structure of a failure
of the rocket - explosion, nozzle ejection, etc. Since the
177
-------
rockets being tested are experimental there is more than a slight
possibility of failure. Thus, provision must be made for either
reducing the effects of failure in the design, or being prepared
to go to the expense of rebuilding. A third problem is the de-
sign and cost of a thrust or flame deflector for rockets tested
in the vertical position. The design and cost was mentioned
only briefly in this report.
A final problem is the design of a "universal" scrubber.
Solid rocket testing may have many purposes. Propellant and
nozzle evaluation, smokeless propellant and plume studies,
thrust vector control evaluation, altitude operation, and missile
nosetip tests are some of the purposes. Of these few types of
tests thrust vectoring presents the major problem of how to re-
direct the exhaust into the scrubber. Altitude testing is
usually conducted in a chamber to which a scrubber could be
easily made an auxiliary. Studies of the characteristics of the
rocket plume will usually preclude the use of a scrubber.
AFRPL PILOT SCRUBBER
Four 10-second duration solid rocket tests were made in
the 5,000 Ibf (22 kN) pilot-scale rocket scrubber at AFRPL
during the program. This scrubber was also a gas-atomized type,
but the entrainment separator was a section of packed Teller-
ettes. Afterburning of hydrogen and carbon monoxide occurred
at the exit of the entrainment separator, which only slightly
affected the scrubber, but destroyed the gas and particulate
sampling systems.
Other instrumentation and visual observations lead to the
conclusion that the scrubber was greater than 99% efficient at
particle (aluminum oxide) removal and probably as efficient at
hydrogen chloride removal. The testing did not proceed into a
series of 30-second duration tests because of the fear that the
afterburning would have more serious consequences for the fiber
glass entrainment separator. It was through that if the system
exit had been vertical upward instead of horizontal the afterburning
178
-------
would have not had a noticeable effect in longer duration tests
COSTS
The installed costs of the conventional and unconventional
designs have been estimated for small and large rockets. The
method used is described so that changes may easily be made as
locations, situations, and costs change. The total installed
cost estimates for the designs for rockets with thrusts ranging
between 5,000 and 500,000 pounds (0.02-2.0 MN) are presented
in Figure 71. For the 450,000 pound (2 MN) thrust rocket the
conventional design installed cost was estimated to be $4.26
million and the unconventional design cost was $1.79 million
(December, 1976).
The yearly operation costs for the large rocket were also
estimated. Consisting of labor, utilities, chemicals, and
replacement parts, the estimated yearly cost is roughly $90,000
if tests are conducted once per month.
179
-------
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Brandt, C.S., and W.W. Heck. Effects of Air Pollutants on
Vegetation in Air Pollution, 2nd Ed. A.C. Stern, ed. , 1968.
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California Air Resources Board. Ambient Air Quality Standards.
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Calvert, S. Source Control by Liquid Scrubbing, 2nd Ed. In:
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Calvert, S., J. Goldshmid, D. Leith, and D. Mehta. Scrubber
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Calvert, S., and S. Stalberg. Evaluation of Systems for Control
of Emissions from Rocket Motors - Phase I. EPA-600/2-75-021-a,
NTIS/PB-245-590, 1975. 48 pp.
V
Calvert, S., S. Yung, and J. Leung. Entrainment Separators for
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1975.
Chaput, L.S. Federal Standards of Performance for New Stationary
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Chilton, C.H., ed. Cost Engineering in the Process Industries.
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Cochran, C.N. Recovery of Hydrogen Fluoride Fumes on Alumina
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Cook, C.C., G.R. Swany, and J.W. Colpitts. Operating Experience
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Coulson, J.M., and J.F. Richardson. Chemical Engineering. Per-
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Dallavalle, J.M. Micromeritics. Pitman Publishing Corp, New
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Edelman, R., J. Boccio, and H. Weilerstein. The Roles of Mixing
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Eisel, J.L., E.W. Price, and E.G. Brown. A1203 Particles Pro-
duced During Solid Propellant Combustion. A1AA Paper No. 74-197,
1974. 6 pp.
Frank C. Brown § Company, Inc.. Manual No. 762-1-M1 for Exhaust
Scrubber Facility, Test Stand C, Building No. E-18, JPL-ETS
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Ridgewood, New Jersey), October 1969.
Fuchs, N.A. The Mechanics of Aerosols. The MacMillan Co,
New York, 1964. 408 pp.
Garrett, J.W., et al. A Design Study for Toxic Rocket Exhaust
Gas Cleaning. AEDC-TR-72-97, AFRPL-TR-72-32, Arnold Engineering
Development Center, Arnold AFS, TN, 1972. 239 pp.
Guthrie, K.M. Capitol Cost Estimating in Modern Cost-Engineering
Technique. H. Popper, ed. McGraw-Hill, 1970. pp. 80-108.
Guthrie, K.M. Process Plant Estimating, Evaluation and Control.
Craftsman Book Company of America, Solana Beach, CA, 1974.
Hill, P.G. and C.R. Peterson, Mechanics and Thermodynamics
of Propulsion, Addison-Wesley Publishing Co., 1965. pp. 563.
Holman, J.P. Heat Transfer, 3rd ed. McGraw-Hill, 1972.
Kemen, R.J. Unpublished results of tests conducted in 1973 and
1974, Naval Air Rework Facility, Jacksonville, Florida, 1976.
Kempner, S.K., E.N. Seiler, and D.H. Bowman. Performance of
Commercially Available Equipment in Scrubbing Hydrogen Chloride
Gas. J. Air Poll. Control Assoc., 20(3) : 139-143, 1970.
Kohl, A.L.,and F.C. Riesenfeld. Gas Purification. McGraw-Hill,
1960. 556 pp.
Kreith, F. Principles of Heat Transfer, 2nd ed. International
Textbook Co., 1965.
Lee Saylor, Inc. Current Construction Costs 1976. Walnut
Creek, CA, 1976.
Lerman, S. The Phototoxicity of Missile Exhaust Products: Short
Term Exposure of Plants to HC1, HF, and A1203. AMRL-TR-75-102,
NTIS/AD-A026837, May 1976. 46 pp.
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Magill, P.L., F.R. Holden, C. Ackley, and F.G. Sawyer. Air
Pollution Handbook. McGraw-Hill, 1956.
Nadler, M.P. Environmental Study of Toxic Exhausts. Naval
Weapons Center, China Lake, CA. AFRPL-TR-76-13, 1976. 105 pp.
Orr, C. Particulate Technology. MacMillan Co., 1966.
Perry, R.H., and C.H. Chilton. Chemical Engineer's Handbook,
5th Ed. McGraw-Hill, 1973.
Peters, M.S., and K.D. Timmerhaus. Plant Design and Economics
for Chemical Engineers, 2nd Ed. McGraw-Hill, 1968.
Philco-Ford Corporation, Aeroneutronics Division, Newport Beach,
CA. Contract F04700-67-C-0583 .
Popper, H. ed. Modern Cost-Engineering Techniques. McGraw-Hill,
1970.
Radke, H.H., L.J. Delaney, and P. Smith. Exhaust Particle Size
Data from Small and Large Solid Rocket Motors. Aerospace Corp-
oration Report No. TOR-1001 (S2951-18)-3, San Bernardino, CA,
1967.
Rush, D., J.C. Russell, and R.E. Iverson. Air Pollution Abate-
ment on Primary Aluminum Potlines. J. Air Poll. Control Assoc.,
23(2):98-104, 1973.
Sedillo, L., Evaluation of the Pilot Rocket Scrubber, Final
Report, U.S. AFRPL, Edwards AFB, CA, AFRPL-TR-77-89, 1978.
Slack, A.V., and G,A, Hollinden. Sulfur Dioxide Removal from
Waste Gases. Noyes Data Corporation, Park Ridge, N.J., 1975.
294 pp.
Stern, A.C. Air and Water Pollution Quality Standards. In:
Industrial Pollution Control Handbook, H.F. Lund, ed., 1971.
Stokinger, H.E., and D.L. Coffin. Biologic Effects of Air
Pollutants. In: Air Pollution, 2nd Ed., A.C. Stem, ed,
1968. pp. 445-546.
Strand, L.D., and G. Varsi. Airborne Measurements of Parti-
culates from Solid Rocket Boosters. In: Proceedings of 8th
JANNAF Plume Technology Conference, Colorado Springs, Colorado,
1974. pp. 141-163.
Sutton, M.P., Rocket Propulsion Elements, John Wiley § Sons,
1963.
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Teller, A.J. Control of Gaseous Fluoride Emissions. Chem. Eng
Progress, 63(3):75-79, 1967.
Tomany, J.P. A System for Control of Aluminum Chloride Fumes.
J. Air Poll. Control Assoc., 19 (6):420-423, 1969.
Weast, R.C., ed. Handbook of Chemistry and Physics, 52nd Ed.
Chemical Rubber Co., 1971.
Yung, S., S. Calvert, and H. Barbarika. Venturi Scrubber Per-
formance Model - Final Report. EPA Contract 68-02-1328, Task
No. 13, July 1976.
183
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APPENDIX A
PROGRAMS FOR CALCULATING QUENCHER EQUILIBRIUM
Closed quencher equilibrium calculation program
FORTRAN IV listing
Closed quencher equilibrium calculation example
Input/Output
Open quencher equilibrium calculation program
FORTRAN IV listing
Open quencher equilibrium calculation example
Input/Output
184
-------
CLOSED QUENCHER EQUILIBRIUM CALCULATION PROGRAM
FORTRAN IV LISTING
100 ALPHA BASEC3)
110 DATA TO/298./
120 XL1NCX1,X2,Y,Y1,Y2>=X1-CX1-X2>/(Y1-Y2)*(Y1-Y>
1 30 TCC(P) = 1668 .21/C7.96681-AL0G10=6.98*T-28498.
150 EH2(T)=6.94*T-2069.
160 E02(T>=7.11*T-2120.
170 ECe2(T)=9.42*T-96864.
180 EH20VCT)=8.1*T-60213.
190 EH20LCT)=18.02*T-73678.
200 EN2CT>=6.97*T-2078.
210 EHCL=0.0*T
230 EAL(T>=21.6*T-391281.
240 INPUT,FC0,FH2,FC02,FH20,FN2,FHCL,FHF,FAL
250 INPUT,X(*UR
260 INPUT,HC0,HH2,HC02,Hh20,HN2,HhiCL,hHF,HAL
270 INPUT,RISP,XMR,AREAE
280 INPUT,AREA2,P2
290 INPUT 49,(BASECI>,!=!,3)
295 49 F0RI*AT(3A4>
300 INPUT,AT0K,HSALT,CO,Cl,C2
310 U1=RISP
320 F=RISP*XMR
330 P2A2=AREA2*P2*101325-
340 PRINT 51
350 51 F0RKAT(/,3X,19HR0CKET C0NSTITUENTS)
360 PRINT 52,FC0,HC0
370 52 F0RfcATOIX,1HF,9X,1HF,/,3X,2HC0, F9.3,F1 0.0)
380 PRINT 53,FH2,HH2
390 53 FeH^AT<3X,2HH2,F9.3,F10.0)
400 PRINT 54,FC02,HC02
410 54 F0R^AT<3X,3HC02,F8«3,F10.0)
420 PRINT 55,FH20,HH20
430 55 F0RMAT<3X,3HH20,F8.3,F10.0>
440 PRINT 56,FN2,HN2
450 56 F0RMATC3X,2HN2,F9.3,F10.0>
460 PRINT 57,FHCL,HHCL
470 57 FeRKATC3X,3HHCL,F8.3,F10.0)
480 PRINT 58,FHF,HHF
490 58 F0RMAT(3X,2HHF,F9.3,F10.0)
500 PRINT 59,FAL,HAL,XHWR
185
-------
510 59 F0RKAT(3X,4HAL0X,F7.3,F10.0,/>3X,9hK0L. WT.=*F4.1>
520 PRINT 60,RISP,XMR,AREAE
530 60 F0RMATC/,2X,11HR0CKET I SP= » F5 .0, 1 X, 3HI*»/S* />
540* 9X>10HKASS RATE=,F6.2*1X,4HKG/S,/,
550& 9X,10HEXIT AREA=,F5 .3*1X,2HM2)
560 PRINT 61,AREAS,P2
570 61 F0RMAT(/,2X,14HSCRUBBER AREA = ,F5.2, 1X,2HI*2,/*
580« 2X,14HBACK PRESSURE = ,F5 . 3, 1 X, 3HATI" >
590 P2=P2*101325.
600 PRINT 62,BASE,HSALT,CO,C1,C2
610 62 F0RN!ATC/,2X,22HNEUTRALIZING AGENT IS
620* 4X,6HHFNET=>F7.0*/*4X,3HCO=>F4.2*2X*
630& F5.3*26X)
640 AM=XMR/XMWR
650 A1=FC0*AM
660 A2=FH2*AK
670 A4=FC02*At^
680 A6=FH20*AM
690 A8R=FN2*Atf
700 A9=FHCL*AM
710 A10=FHF*A^
720 Al 1=FAL*AI^
730 GfS=0.0
740 HHR=A1 *HC0-i-A2*HH2+A4*HC02+A6*HH20+A8R*HN2
750& +A9*HHCL+A10*HHF+Al1*HAL
760 ASALT=0.0
765 AT0N=AT0M
770 IFCAT0M.EQ.O.O) G0 T0 20
780 ASALT=CA9+A10>/AT0I*>
785 IFCBASEC1).EQ."N0NE"> AT0N=0.0
790 A9=0.0
800 A10=0.0
810 20 INPUT>RA
820 IF(RA.LT.O.O) ST0P
830 A3=(RA*XMR>*0.21/28.97
840 A8A=A3*3.76
850 A8=A8R+A8A
860 Y3=A3-(Al+A2)/2.
870 Y3=AMAX1CO.O»Y3)
880 IFCA1.GT.A2) G0 T0 70
890 Y1=A1-A3
900 Yl =AN'AX1 (0.0,Yl )
910 Y2=A1+A2+2.*(Y3-A3)-Y1
920 G0 T0 80
930 70 Y2=A2-A3
940 Y2=AMAX1(0.0/Y2)
950 Y1=AH-A2 + 2.*(Y3-A3)-Y2
960 80 Y4=A1+A4-Y1
186
-------
970 YG=Y1+Y2+Y3+Y4+A8+A9+A10
980 X1=Y1/YG
990 X2=Y2/YG
1000 X3=Y3/YG
1010 X4=Y4/YG
1020 XI*WG=(Y1*28.0H-Y2*2.01 6+Y3*32.+Y4*44.009 + A8*28.01 34 +
1030* A9*36.461+A10*20.0064)/YG
1040 PRINT 63,YG,X1,X2,X3,X4,XOWG
1050 63 F0RMAT(/*2X>40HT0TAL GAS FL0W RATE k/0 WATER VAP0R,NG =,
1070* 2X,6HC02/NG,2X>3HMWG.»/>1X.»4F7.3,F7.1 )
1080 PRINT 64
1090 64 F0RKAT(/»3X*2HRW,4X*2HRA*4X*1HT*5X,4HPV/P*5X*1HQ*
1 100A 7X,1HU,6X,2HP1,3X,6HNWL/NG,/,15X,1HC*12X,4HM3/S*
1110& 5X*3HM/S*5X*3HATM*1IX)
1 120 30 INPUT,RVJ
1130 IF(RW.LT.O.O) G0 T0 20
1150 A5=A6+A7
1160 Y5=A2+A5-Y2
1180 HH1=HHR+A7*EH20L(TO)+A3*E02(TO)+A8A*EN2+A8*EN2CT)+A9*EHCLCT)+A10*EHFCT>+A1 1*EAL(T)
1370« +ASALT*HSALT
1380 E2 = HH2 + XM2*IJ2*U2/8368.
1390 PR=PV/P2
1400 IFCE2.LT.E1) G0 T0 150
1410 TCP=TC
1420 PRP=PR
I 430 Q2P=Q2
1440 U2P=U2
187
-------
450 P1P=P1
460 E2P=E2
470 Y7P=Y7
480 G0 T0 100
490 150 TC=XLIN(TC*TCP*El,E2*E2P>
500 PR=XLIN(PR,PRP,E1,E2,E2P>
510 Q2=XLIN(Q2,Q2P*E1,E2,E2P>
520 U2=XLINCU2,U2P,E1,E2,E2P>
1530 PI=XLIN
1540 Y7=XLINF8.2,3X)
610 ST0P
620 END
188
-------
CLOSED QUENCHER EQUILIBRIUM CALCULATION EXAMPLE
INPUT/OUTPUT
? . 194,.29A» .014, .126*. 087*.171,0, .114
j
29.2
7
-9365,15993,-66380,-35426,16886,-5560,0,-319630
7
2590.,772.2,1.53
?50.,0 .90
?
HCL
?1 ,-17868.,9.99,.029,1 .37
R0CKET C0NSTITUENTS
C0
H2
C02
H20
M2
HCL
HF
F
0. 194
0 .294
0.014
0.126
0.087
0.171
0.
H
-9365.
1 5993.
-66380.
-35426.
16886.
-5560.
o.
AL0X 0.114 -319630. .
M0L. WT.=29.2
R0CKET ISP=2590. K/S
I«ASS RATE = 772.20 KG/S
EXIT AREA=1.530 f2
SCRUBBER AREA=50.00 M2
BACK PRFSSURE=0.900 ATM
NEUTRALIZING AGENT IS HCL
HFNET=-17888.
C0=9.99 Cl=0.02900 C2 = 1.370
189
-------
T0TAL GAS FL0W RATE to/0 WATEH VAH0R*NG = 17.121 KGM0L/S
C0/NG H2/NG 02/NG C02/NG flwG
0.267 0.421 0. 0.054 17.9
HW RA T PV/P Q U PI NkL/NG
C Ni3/S l*/S ATI"
20.0 0.1 90.4 0.777 2541.12 50.8 0.690 47.07
18.0 0.1 91.2 0.800 2848.22 57.0 0.692 41.54
16.0 0.1 91.9 0.820 3170.07 63.4 0.692 35.97
14.0 0.1 92.5 0.837 3503.66 70.1 0.688 30.39
12.0 0.1 93-1 0.851 3846.67 76.9 0.680 24.80
10.0 0.1 93.6 0.864 4197.69 84.0 0.668 19.19
8.0 0.1 94.3 0.874 4556.24 91.1 0.652 13.57
6.0 0.1 95.5 0.883 4925.01 98.5 0.631 7.95
4.0 0.1 102.3 0.890 5351.58 107.0 0.607 2.37
3.0 0.1 108.5 0.885 5158.18 103.2 0.587 0.23
3.5 0.1 108.6 0.892 5511.84 110.2 0.601 1.00
190
-------
OPEN QUENCHER EQUILIBRIUM CALCULATION PROGRAM
FORTRAN IV LISTING
10 C0f"l*0N P2,AT0l*,AT0N,ASA|_T,CO,Cl,C2,Xf>2,El
20 C0MM0N Y1,Y2,Y3,Y4,Y5,Y6,A8,A9,A10,A11
30 C0KK0N RA,RW,HSALT,XMWG
40 DIMENSI0N ADI(10>,GDI(10)»Rkl(10)
50 ALPHA BASEC3)
60 DATA TO/298./
90 EC0=6.98*T-28498.
100 EH2(T>=6.94*T-2069.
110 E02=7.11*T-2120.
120 EC02(T)=9.42*T-96864.
130 EH20V=6.97*T-2078.
160 EHCLCT>=6.7*T+4.2E-4*T*T-24094.
170 EHF(T)=0.0*T
180 EAL=21 .6*T-3/2*A
190 INPUT*FC0*FH2*FC02*FH20,FN2»FHCL»FHF,FAL
200 INPUT*XN)WR
210 INPUT*HC0,HH2*HC02*HH20,HN2»HHCL*HHF>HAL
220 INPUT*RISP*XMR»P2,AE
230 INPUT 49,(BASE(I)*I=1>3)
240 49 F0KI*ATC3A4>
250 INPUT, AT0I*»HSALT>CO»C1 , C2
260 INPUT*NB
270 INPUT,(ADI(I)>!=!»NB)
280 INPUT, (GDI CD, 1 = 1 ,NB)
290 INPUT,CRWICI),1=1,NB)
300 U1=RISP
310 F=RISP*XMR
320 PRINT 51
330 51 F0RMAT(/,3X,19HR0CKET C0NSTITUENTS)
340 PRINT 52,FC0,HC0
350 52 F0RMATU1X,1HF,9X,1HH,/,3X,2HC0, F9.3, F 1 0 .0 )
360 PRINT 53,FH2,HH2
370 53 F0RMAT<3X,2HH2,F9.3,F10.0>
380 PRINT 54,FC02,HC02
390 54 F0RI^AT(3X,3HC02,F6.3,F10.0)
400 PRINT 55,FH20,HH20
410 55 F0R^AT(3X,3HH20,F8.3,F10.0)
420 PRINT 56,FN2,HN2
430 56 F0RMATC3X,2HN2,F9.3,FIO.O)
440 PRINT 57,FHCL,HHCL
191
-------
450 57 F0RMAT<3X*3HHCL*F8.3*F10.0>
460 PRINT 58*FHF*HHF
470 58 F0RMAT(3X*2HHF*F9.3*F10.0>
480 PRINT 59*FAL*HAL*XttkR
490 59 F0RMATC3X*4HAL0X*F7.3*F10.0*/*3X*9HM0L« WT.=*F4.1>
500 PRINT 60*RISP,XMR,P2*AE
510 60 F0RMATC/*2X*11HR0CKET ISP=*F5.0*1X,3HM/S*/*
520& 9X*10HMASS RATE=, F6 .2* 1 X, 4HKG/S* /*
530& 9X,9HPRESSURE=*F5.3*1X*3HATM*/*9X*13HEXHAUST AREA=,
540* F5.3*1X>2HM2)
550 P2=P2*101325.
560 PRINT 62,BASE,HSALT,CO,C1,C2
570 62 F0RI*AT*PX,22HNEUTRALIZING AGENT IS ,3A4*/,
580 & 4X,6HHFNET = ,F7.0,/,4X,3HCO = ,F4.2*2X,3HC1=,F7
590 & F5.3)
600 PRINT 66*NB
610 66 F0Rl^AT(/,2X,18HNUKBER 0F BAFFLES=,12>
620 PRINT 67*(ADI(I)*I=1*NB)
630 67 F0RMATC4X,14HDRAG AREAS* N'2* /* 4X* 10F6 .2 )
640 PRINT 68*(GDI,I = 1 *NB)
650 68 F0HMATC4X*10HDRAG C0EFS*/*4X*1OF6.3)
660 PRINT 69*,I=1,NB>
670 69 F0RN!AT(4X* 12HWATER RATI 0S*/* 4X* 10F6 .2 >
680 AK=XMR/XKWR
690 Al=FC0*Af^
700 A2=FH2*AM
710 A4=FC02*AK
720 A6=FH20*A!«
730 A8R=FN2*A^
740 A9=FHCL*AM
750 A10=FHF*AM
760 A11=FAL*AM
770 HHR=Al*HC0+A2*l"iH2+A4*HC02+A6*HH20+A8R*HN2
780* +A9*HHCL+A10*HHF+A11*HAL
790 ASALT=0.0
800 AT0N=AT0Ki
810 B=F
820 IFCAT0i«.EQ.O.O) G0 T0 20
830 ASALT=(A9+A10)/AT0M
840 IFCBASEC1).EQ."N0NE"> AT0N=0.0
850 A9=0.0
860 A10=0.0
870 20 INPUT*RA
880 IF(RA.LT.O.O) ST0P
890 A3=*0.21/28.97
900 A8A=A3*3.76
910 A8=A8R+A8A
920 Y3=A3-(Al+A2)/2.
192
-------
930 Y3=AMAX1 (0 .0,Y3>
940 IFCA1.GT.A2) G0 T0 70
950 Y1=A1-A3
960 Yl =Af*AXl (0 «0,Y1 )
970 Y2=A1+A2+2.*CY3-A3)-Y1
980 G0 T0 80
990 70 Y2=A2-A3
1000 Y2=AI"AX1 CO «0*Y2>
1010 Y1=AH-A2 + 2.*CY3-A3)-Y2
1020 80 Y4=A1+A4-Y1
1030 YG=Y1+Y2+Y3+Y4+A8+A9+A10
1040 X1=Y1/YG
050 X2=Y2/YG
060 X3=Y3/YG
070 X4=Y4/YG
080 XMtoG=(Yl*28.01+ Y2*2 .016 +Y3*32.+Y4*44.009+A8*28.0134 +
090* A9*36.461+A10*20.0064>/YG
100 FD=0.0
110 PRINT 63>YG,X1 * X2, X3*X4,X(^kG
120 63 F0KMATC/,2X,40HT0TAL GAS FL0W RATE W/0 WATER VAP0R»NG
30* F7.3*1X*7HKGM0L/S*/*3X*5HC0/NG*2X,5HH2/NG*2X,5H02/NG,
40& 2X>6HC02/NG,2X,3HI«WG,/,1X,4F7.3.»F7.1 )
50 PRINT 64
60 64 F0RMA1(/,3X,2HRW>4X*2HRA,4X*1HT*5X,4HPV/P*5X>1HQ»
70& 7X,1HU>6X,2HA2*3X,5HRH0GV*3X,3HM0K,/,1 5X, I HC, 1 2X,
180& 5X*3HI^/S*5X»2HM2*3X*5HKG/N3*4X, 1HN)
190 RU=0-0
200 D0 200 J=l*NB
210 AD=ADI(J)
220 CD=CDICJ)
230 RU! = RW + RWI CJ)
240 B=B-FD
1250 A7=RW*XMR/18.02
1260 A5=A6+A7
1270 Y5=A2+A5-Y2
1280 XM2=XKR*C1.+RW+RA)
1290 HH1=HHR+A7*EH20L(TO)+A3*E02(TO)+A8A*EN2(TO)
1300 El =HHH-XKiR*Ul*Ul/8368.
1310 IF(RWI(J) .EQ.0.0) G0 T0 150
1320 CALL CALCCB,RH0GV*U2>
1330 150 CONTINUE
1 340 FD=0.0
1350 IFCAD.EQ.O.0) G0 T0 200
1360 FD=CD*AD*RH0GV*U2*U2/2.
1370 BD=B-FD
1380 CALL CALCU2)
1390 200 C0NTINUE
1400 ST0P
193
-------
1410 END
1420 SUBROUTINE CALCY2,Y3*Y4,Y5,YG>A8,A9,A10,A11
1450 C0M0N RA,Rk,HSALT,XMWG
1 460 XLINCX1,X2,Y*Y1, Y2 > =X1 - /< Y 1-Y2>*(Y1-Y>
1470 TCCCP>=1668.21/(7.96681-AL0G10>-228.
1480 EC0(T)=6.98*1-28498.
1490 EH2(T)=6.94*1-2069.
1500 E02(T)=7.11*1-2120.
1510 EC02(T)=9.42*1-96864.
1520 EH20V(T)=8.1*1-60213.
1530 EH20L(T)=18.02*1-73678.
1540 EN2CT)=6.97*1-2078.
1 550 EHCL(T) = 6.7*T+4.2E-4*T*T-24094.
1560 EHF>
1610 Y6=PV*YG/(P2-PV)
1620 Y7=Y5-Y6+AT0N*ASALT
1630 IF(Y7 .LI .0 .0) G0 T0 100
1640 GP.S=55.49*ASALT/Y7
1650 Gf S=AKiINl (GKiS»CO)
1660 APV=PV/133.322
1670 IF(ATei^.NE.O .0) AP V=PV*< 1 . + C1 *GMS**C2 )/I 33. 322
1680 TC=TCCCAPV)
1690 T=TC+273.15
1700 Q2=YG*8314.*T/CP2-PV>
1 750 HH2 = Yl*EC0CT)+Y2*EH2CT)+Y3*E02(T)-»-Y4*EC02(T) +
1760& Y7*EH20L+AI1*EAL(T)
1770& +ASALT+HSALT
1780 E2 = HH2 + XI^2*U2*U2/8368.
1790 PR=PV/P2
1800 IF(E2.LT.E1> G0 T0 150
1810 TCP=TC
1820 PRP=PR
1830 Q2P=Q2
1850 E2P=E2
1860 G0 T0 100
1870 150 TC=XL1N(TC>TCP>E1,E2>E2P)
1880 PR=XLIN(PR,PRP,E1,E2,E2P)
1890 Q2=XLINF7.3*F9.2*F7.1 *F7 .2*F 7 . 3*E10 . 3)
1980 RETURN
1 990 END
194
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OPEN QUENCHER EQUILIBRIUM CALCULATION EXAMPLE
INPUT/OUTPUT
?.194*.294*.014,.126*.087*.171*0*.114
7
29.2
?
-9365,15993*-66380*-35426*16886,-5560,0*-319630
7
2590.*772.2*0.90*1.53
7
HCL
7
1*-17888.*9.99*.029*1.37
?9
70*0*0*0*0*0*0*0,0
70*0*0*0*0*0*0*0*0
74,2,2*2*2*2*2*2*2
R0CKET C0NSTITUENTS
C0
H2
C02
H20
N2
HCL
HF
F
0. 194
0.294
0.014
0.126
0.087
0.171
0.
H
-9365.
1 5993.
-66380.
-35426.
168&6.
-5560.
0 .
AL0X 0.114 -319630.
U'T.=29.2
R0CKET ISP=2590. M/S
I"ASS HATE = 772.20 KG/S
PHESSUHE=0.900 AIM
EXHAUST AREA=1.530 K2
NEUTRALIZING AGENT IS HCL
HFNET=-1 7888.
C0 = 9.99 Cl=0. 02900 C2 = 1.370
NUI^BEK 0F BAFFLES= 9
DRAG AREAS* M2
0. 0. 0. 0. 0. 0. 0. 0.
DRAG C0FFS
Q. 0. 0. 0. 0- 0. 0. 0.
WATER RAT I 0S
4.00 2.00 2.00 2.00 2.00 2.00 2.00 2-00 2.00
195
-------
TOTAL GAS FL0W RATE W/0 LATER VAP0R,NG
C0/NG H2/NG 02/NG C02/NG MVvG
0.267 0.421 0. 0.054 17.9
= 17.121 KGI*0L/S
RV-
4.0
6.0
8.0
10.0
12.0
14.0
16.0
18.0
20.0
Rt
0.
0.
o.
0.
o.
o.
o.
0.
o.
* T
c
99 .8
95.1
94.1
93.4
92.9
92.4
91 .8
91 .0
90 .2
PV/P Q U A2 RH0GV MB*
K'3/S I*/S M2 KG/M3 N
0.882 4930.13 507.8 9.71 0-529 0.200E+07
0.876 4648.17 364.8 2.74 0.536 0.200E+07
0.868 4349.40 264.6 5.28 0.538 0.200E+07
0.858 4034.83 233.3 7.29 0.539 0.200E+07
0.846 3713.78 197.7 8.78 0-539 0.200E+07
0.832 3392.30 171.5 9.78 0.540 0=200E+07
0.815 3074.86 151.5 20.30 0-541 0.200E+07
0.794 2765.68 135.6 20.40 0.542 0.200E+07
0.770 2469.03 122.7 20.11 0 . 543 0 .200E+07
196
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APPENDIX B
CONVERSIONS
joule (J) x 0.2390 = calorie (cal)
joule (J) x 9.478 x 10'1* = BTU
liter (£) x 0.2642 = gallon (gal)
meter (m) x 3.281 = feet (ft)
m3 x 35.31 = ft3
newton (N) x 0.2248 = pounds (Ibf)
pascal (N/m2) x 1.450 x ID'" = lbf/in2 (psi)
pascal (N/m2) x 4.019 x 10"3 = inch water column (in. W.C.)
pascal (N/m2) x 1.020 x 10"2 = centimeter water column (cm W.C.)
watt (W) x 1.340 x 10 ~3 = horsepower (HP)
watt (W) x 3.4.12 = BTU/hour
197
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TECHNICAL REPORT DATA
(Please read Instructions on the reverse before completing)
. REPORT NO.
EPA-600/7-78-057
2.
3. RECIPIENT'S ACCESSION NO.
. TITLE ANDSUBTITLE
Design Criteria for Rocket Exhaust Scrubbers
5. REPORT DATE
March 1978
6. PERFORMING ORGANIZATION CODE
7. AUTHOR(S)
Harry F. Barbarika and Seymour Calvert
8. PERFORMING ORGANIZATION REPORT NO.
9. PERFORMING ORGANIZATION NAME AND ADDRESS
Air Pollution Technology, Inc.
4901 Morena Boulevard, Suite 402
San Diego, California 92117
10. PROGRAM ELEMENT NO.
1AB012; ROAP 21ADL-101
11. CONTRACT/GRANT NO.
68-02-2145
12. SPONSORING AGENCY NAME AND ADDRESS
D PERIOD COVERED
*EPA, Office of Research and Development
Industrial Environmental Research Laboratory
Research Triangle Park, NC 27711
13. TYPE OF REPORT AND PERIOD C(
Task Final; 12/75-12/77
14. SPONSORING AGENCY CODE
EPA/600/13
15.SUPPLEMENTARY NOTES jERL-RTP task officer is Dale L. Harmon, Mail Drop 61, 919/
541-2925. (*) Cosponsor is the U.S. Air Force, Edwards AFB.
The report gives results of an engineering study and design of methods for
scrubbing the exhaust of static-tested solid rockets. Pollutants of major concern
were hydrogen chloride and hydrogen fluoride gases. The best process for removing
these gases was found to be a gas-atomized spray scrubber, using the power sup-
plied by the rocket to atomize the scrubbing liquid. Four tests in the 22 kN pilot-
scale rocket scrubber at the U.S. Air Force Propulsion Laboratory were analyzed
to aid in the design. Two types of gas-atomized scrubbers were designed: one was
a conventional design similar to a venturi; the other was a low-cost unconventional
open type, using neither pressure piping nor a ducted spray chamber. Cost analyses
were made for both types for rockets with thrusts between 20 kN and 2 MN.
17.
KEY WORDS AND DOCUMENT ANALYSIS
DESCRIPTORS
b.IDENTIFIERS/OPEN ENDED TERMS
c. COSATI Field/Group
Air Pollution
Solid Rocket Fuels
Combustion
Exhaust Gases
Scrubbers
Static Tests
Hydrogen Chloride
Hydrogen Fluoride
Atomizing
Spraying
Air Pollution Control
Stationary Sources
Gas-atomizing
13 B
211
21B
07A
14B
07B
13H
13. DISTRIBUTION STATEMENT
Unlimited
19. SECURITY CLASS (ThisReport)
Unclassified
21. NO. OF PAGES
214
20. SECURITY CLASS (Thispage)
Unclassified
22. PRICE
EPA Form 2220-1 (9-73)
198
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