CTD A U.S. Environmental Protection Agency Industrial Environmental Research
    f^ Office of Research and Development Laboratory
                       Research Triangle Park, North Carolina 27711
                                EPA-600/7-78-057
                                      4t\-ro
                                MSfCh 1978
DESIGN CRITERIA FOR
ROCKET EXHAUST SCRUBBERS
Interagency
Energy-Environment
Research and Development
Program Report

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                                                EPA-600/7-78-057
                                                       March 1978
              DESIGN  CRITERIA FOR
       ROCKET  EXHAUST SCRUBBERS
                                by

                    Harry F. Barbarika and Seymour Calvert

                       Air Pollution Technology, Inc.
                      4901 Morena Boulevard, Suite 402
                        San Diego, California 92117
                         Contract No. 68-02-2145
                           ROAP21ADL-101
                       Program Element No. 1AB012
           Project Officers: Dale L. Harmon (EPA) and Luciano Sedillo (AFRPL)

                  Industrial Environmental Research Laboratory
                    Office of Energy, Minerals and Industry
                     Research Triangle Park, N.C. 27711
                            Prepared for

U.S. ENVIRONMENTAL PROTECTION AGENCY  and  U.S. AIR FORCE ROCKET PROPULSION
    Office of Research and Development                 LABORATORY
        Washington, D.C. 20460                 Edwards Air Force Base, CA
                                               93523

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                            ABSTRACT

      An engineering study and design of methods of scrubbing
the exhaust of experimental rocket propulsion systems have been
performed.  The study included an evaluation of the cost and
technical feasibility of scrubbing the rocket exhausts.  The ex-
haust products of major concern were hydrogen chloride and, to
a lesser extent, hydrogen fluoride gases which result from the
combustion of solid propellant rockets.  The best process for
removing these similar gases was found to be a gas-atomized spray
scrubber which used the power supplied by the rocket to atomize
the scrubbing liquid.  Four tests in the 5,000 pound (22 kN)
pilot scale rocket scrubber at the Air Force Rocket 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 which did not use pressure piping nor a ducted spray chamber.
Cost analyses were made for both types of scrubbers for rockets
with thrusts between 4,500 pounds (20 kN) and 450,000 pounds (2 MN)
                               ill

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                            CONTENTS

Abstract	iii
Figures	vi
Tables	x
Abbreviations and Symbols	xii
    1. Introduction	1
          Objectives 	  2
          Approach 	  2
          Requirements 	  3
    2. Background Information	g
          Solid Propellant Rockets 	  g
          Supersonic Flow	7
          Static Test Firings	8
          Combustion of Solid Rocket Exhaust 	  9
          Flow Conditions of Solid Rocket Exhaust	9
          Aluminum Oxide Particles 	  ^7
          Liquid Propellants	-.„
    3. Emission Reduction Objectives 	  20
          Required Efficiency	20
          Constraints of Control Method	21
    4. AFRPL Pilot Scrubber Tests	22
          Introduction 	  22
          Solid Rocket Exhaust Conditions	22
          Scrubber Design	23
          Design Modification - Liquid Injection System. .  .  26
          Instrumentation	31
          Gas and Particulate Sampling	33
          Momentum Reduction Experiments 	  33
          Discussion of Tests	35
          Conclusions from AFRPL Test Program	58
                               IV

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                      CONTENTS (continued)
    5. Air Pollution Control Equipment Alternatives	62
           Introduction	62
           Selection of Air Pollution Control Equipment. .  .  62
           Equipment for Removal of HC1 and HF	64
           Scrubbers Used on High Energy Exhausts	65
           Potentially Suitable Scrubbers	69
    6. Detailed Design	71
           Definition of the Scrubbing Process 	  71
           Performance Requirements	96
           Conventional Scrubber Design	98
           Conventional Scrubber Entrainment Separation. .  .127
           Conventional Scrubber Costs 	 135
           Unconventional Scrubber Design	149
           Unconventional Scrubber Entrainment Separation.  . 157
           Unconventional Scrubber Costs 	 164
           Scrubber Operation Costs	 174
    7. Conclusions and Recommendations 	  ..... 176

References	180
Appendices	184
    A. Programs for Calculating Quencher Equilibrium ..... 184
    B. Conversions	197
                                 v

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                          FIGURES


Number

  1   Rocket exhaust flow rates	    12

  2   Rocket exhaust power	    15

  3   AFRPL 22kN pilot scrubber	    24

  4   Injector	    27

  5   Location of manometer probes and thermocouples ....    32

  6   Sketch of location of thermocouples in entrainment
      separator section	    34

  7   Location of filter holders in the entrainment
      separator exit .....  	    35

  8   Test 1, rocket chamber pressure	    37

  9   Test 1, scrubber duct wall temperature	    37

 10   Test 1, scrubbing liquid flow rate	    38

 11   Test 1, sampling pump vacuum pressure	    38

 12   Photograph of AFRPL scrubber exit after test 1 ....    39

 13   Test 1, size distribution data	    44

 14   Test 2, rocket chamber pressure	    45

 15   Test 2, scrubber duct wall temperature	    45

 16   Test 2, gas temperature in entrainment separator from
      thermocouple arrays	    46

 17   Test 2, sampling pump vacuum pressure	    46

 18   Test 2, velocity profile in scrubber	    48

 19   Photograph of liquid injectors after test 2	    50

 20   Test 3, rocket chamber pressure	    51


                                 vi

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                       FIGURES (continued)


Number                                                       Page

  21   Test 3, scrubber duct wall temperature	51

  22   Test 3, gas temperature in entrainment  separator from
       thermocouple arrays	52

  23   Test 3, velocity profile in scrubber	52

  24   Test 4, rocket chamber pressure	54

  25   Test 4, scrubber duct wall temperature	54

  26   Test 4, gas temperature in entrainment  separator from
       thermocouple arrays. ... 	  55

  27   Test 4, pressure near rocket nozzle exit  (diffuser
       inlet)	•	55

  28   Test 4, velocity profile in scrubber	57

  29   Schematic of rocket exhaust gas scrubbing process.  .  .  72

  30   Comparison of velocity prediction with  Garrett, et  al.
       (1972)	80

  31   Comparison of temperature prediction with Garrett,  et   80
       al  (1972)	

  32   Flow conditions in  a closed quencher	83

  33   Gas volume flow rate of 2 MN rocket closed quencher.  .  83

  34   Pressure rise in quencher	84

  35   Pressure rise versus thrust to duct area ratio ....  84

  36   Concentration of HC1 absorbed in quencher liquid  ...  85

  37   Water vapor volume  fraction	85

  38   Flow conditions in  an open quencher	89

  39   Effect of entrained air on equilibrium velocity.  ...  94

  40   Physical absorption of HC1	107

  41   Liquid to gas ratio in scrubber	109
                                 Vll

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                       FIGURES (continued)

Number                                                      Page

 42   Effect of gas velocity on plume length required for
      HC1 mass transfer to scrubbing liquor at 0.9  atm.  . .   H6

 43   Rocket thrust coefficient	 •   12°

 44   Sketch of gas-atomized scrubber section 	   122

 45   Cyclone with tangential gas  inlet	132

 46   Outline of module cost format	136

 47   Quench duct installed costs  	   143

 48   Atomized spray scrubber costs 	   143

 49   Cyclone costs	144

 50   Water storage tank costs	144

 51   Caustic slurry supply pump costs	145

 52   Caustic storage tank costs	145

 53   Pipeline costs	146

 54   Quench and scrubbing liquor  pump costs for  small
      rocket scrubbers	146

 55   Drainage sewer costs	147

 56   Drying bed costs	147

 57   Total installed costs for conventional scrubbers.  . .   148

 58   Schematic of non-conventional rocket  exhaust  gas
      scrubbing process 	   151

 59   Plan  view of non-conventional scrubbing system.  .  . .   152

 60   Typical cross-section of a non-conventional absorber
      configuration alternative 	  ......   153

 61   Sketch showing arrangement of tube  bank separator  . .   163

 62   Transition piece costs	169

 63   Scrubber excavation costs 	   169
                               Vlll

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                      FIGURES (continued)

Number                                                       Page

 64   Costs for scrubber excavation retaining walls,
      gunite, drag and deflector units for large rockets . . I70

 65   Entrainment separator costs	1^0

 66   Scrubbing liquor viaduct and manifold costs	171

 67   Water storage tank costs	171

 68   Caustic storage tank costs	172

 69   Drainage sewer costs 	 172

 70   Drying bed costs	173

 71   Total installed costs for unconventional scrubber
      compared with conventional design   	 173
                                 IX

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                            TABLES

Number                                                      Page

  1   Composition of solid rocket  exhaust  .........   10

  2   Flow conditions of solid rocket  exhaust  -  composite
      propellant ......................   13
  3   Estimated extremes  of the  ratio  of  nozzle  exit
      temperature to pressure  ...............   14

  4   Enthalpies of rocket exhaust .............   16

  5   Composition of liquid propellant rocket  exhaust  .•  .  .   19

  6   Required removal efficiencies  ............   20

  7   Composition of test rocket exhaust ..........   23

  8   Predicted AFRPL test rocket flow conditions  .....   23

  9   Parameters for comparison  with Garrett  (1972)  ....   79

 10   Parameters for 2 meganewton rocket  quenching predic-
      tion .........................   82

 11   Equilibrium quench  composition ............   86

 12   Combustion properties of H2 and  CO  in air  at standard
      conditions ......................   88

 13   Scrubber inlet conditions  for preliminary  sizing.  .  .   98

 14   Empirical constants for equation 35 .........   99

 15   Comparison of scrubber diameters ........... 100

 16   Capital  cost estimates for large rocket  scrubber.  .  . 102

 17   Basic  chemical costs,  mid-1977 ............ 104

 18   Chemical solubilities ................ 105

 19   Gas  conditions at scrubber outlet  .......... 127

 20   List of  processing  units making  up  the  total conven-
      tional scrubbing process ........  ....... 138

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                      TABLES (continued)

Number                                                      Page

 21   Worked examples of atomized spray scrubber costs .  .  .  142

 22   Equilibrium conditions in unconventional scrubber
      without baffles	155

 23   List of components making up the total unconventional
      scrubbing process - large rockets  	  165

 24   Worked examples of non-conventional scrubber costs  .  .  168

 25   Scrubber operation plan	174
                                XI

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                      LIST OF SYMBOLS



Latin



  a    = speed of sound,  m/s



  a    = cyclone dimension,   figure 35



  A    = area, cm2 or m2



  b    = cyclone dimension,  figure 35



  c*   = characteristic exhaust velocity,  m/s



  C    = mass concentration, fraction



  Cn   = thrust coefficient, dimensionless
   r


  Cp   = specific heat at constant pressure,  cal/g-°K



  CD   = resistance coefficient,  dimensionless
   K


  G.J.   = nozzle thrust coefficient, dimensionless



  d    = diameter,  cm or  m



  d    = particle diameter,  ym,  ymA,  or  cm



  D    = diameter,  m



  D    = cyclone dimension,  figure 35




  D    = cyclone dimension,  figure 35
   c


  Dp   = gas  phase  diffusivity,  m2/s



  ED   = required efficiency,  %
   K


  f    = fraction



  F    = thrust,  N



  h    = convective heat  transfer coefficient,  cal/s-cm2-°K



  h    = cyclone dimension,  figure 35



  k    = conductivity,  cal/s-cm-°K




                               xii

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                 LIST OF SYMBOLS (continued)





Latin


  k     = mass transfer coefficient, kgmol/s-m2-atm



  Kg    = overall mass transfer coefficient, kgmol/s-m2-atm



  L/G   = liquid to gas volume ratio, £/m3 or m3/m3



  m     = mass flow rate, kg/s



  M     = mach number, dimensionless



  M.W.  = molecular weight, kg/kgmole



  n     = mole flow rate, kgmol/s



  N     = molar flux, kgmol/s-m2



  NQG   = number of overall mass transfer units



  p     = pressure, N/m2 (Pa)



  P     = pressure, N/m2 (Pa)



  P     = Prandtl number, dimensionless



  q     = heat flow rate, cal/s



  Q     = volume flow rate, m2/s



  R     = gas constant, J/°K-kg



  R     = mass flow rate ratio, kg/kg



  R     = Reynolds number, dimensionless
   Q


  S     = cyclone dimension, figure 35



  S     = Schmidt number, dimensionless
   \*


  t     = thickness, cm



  T     = temperature, °K



  u     = velocity, m/s



  x     = mole fraction in liquid



  y     = mole fraction in gas
                                Xlll

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                 LIST OF SYMBOLS (continued)



Greek


  T     = ratio of specific heats,  dimensionless


  Ahf   = enthalpy of formation per mole, J/kgmol


  e     = expansion ratio


  n     = efficiency, fraction or I


  9     = time, s


  U     = viscosity, k.g/m-s


  p     = density, kg/m3 or g/cm3


  a     = geometric standard deviation
   o




Subscripts


  a     = air


  c     = column


  d     = drop


  D     = drag, duct


  e     = exit


  EL    = emission limitation


  f     = mean of free stream and surface


  g     = gas


  gc    = geometric count mean


  G     = gas


  in    = inlet


  L     = liquid


  o     = stagnation


  out    = outlet


                                 xiv

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                    LIST OF SYMBOLS (continued)
Subscripts
  p     = pollutant
  r     = rocket, relative
  R     = required
  T     = total
  u     = universal
  v     = vapor
  w     = water
  x     = x-axis direction
  y     = y-axis direction
  1     = upstream
  2     = downstream
                                 xv

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                           SECTION 1
                          INTRODUCTION

     The Air Force Rocket Propulsion Laboratory (AFRPL) has been
concerned about the environmental effect of their experimental
rocket testing for a number of years.  In recent years the various
efforts undertaken to understand and develop technologies to
reduce the effects have taken the form of engineering design
studies, pilot scale scrubber tests with experimental liquid and
solid rockets, modeling of the atmospheric diffusion and disper-
sion of exhaust plumes, and analyses of HC1 concentrations in
exhaust clouds and at ground level.  The goal of the engineering
design studies has been to determine the technical and economic
feasibility of scrubbing rocket motor exhausts.  The primary
exhaust products of interest being the particulate - A1203,
and the gases - CO, HC1, and HF, which result from experimental
rocket tests.
     This report describes the rocket exhaust scrubber system
design effort which has been performed under a U.S.  Environmental
Protection Agency  (EPA] contract but was instigated by the AFRPL
through an interagency agreement.  The objective of the program
has been to define the technology, feasibility, and cost for
reducing the gaseous and particulate emissions from rocket firings
so that a decision can be made as to what level of atmospheric
pollution reduction is attainable and feasible.
     The exhaust  scrubber system design program began with a
joint program by the AFRPL and the Arnold Engineering Development
Center (AEDC) at Tullahoma, Tennessee.  The result was the 5,000
pound (22 kN) thrust rocket exhaust scrubber system installed
at AFRPL.  A number of liquid and solid rockets were tested in
this pilot facility.  This report contains an analysis of the
four solid rocket tests conducted.

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     The first phase of the present design effort was also per-
formed by Air Pollution Technology, Inc. (A.P.T.) under an EPA
contract, no. 68-02-1328.  The results of this first phase study
to determine what systems might be applicable to cleaning the
exhaust from test firings of solid rockets were published in
the report by Calvert and Stalberg (1975).

OBJECTIVES
     The technical objective of the project has been to prepare
design criteria for scrubbing systems to handle the exhaust from
solid rocket motors ranging in thrust from 5,000 pounds (22 kN)
to 450,000 pounds (2 MM).  The design criteria include tradeoffs
of scalability, cost, scrubbing efficiency, feasibility, and
availability of equipment.  The criteria also include definition
of the processes involved in quantitative terms.  Such important
design specifications as flow rates of coolants, neutralizing
agents, power and fuel requirements,  treatment of the scrubbing
media and coupling effects between scrubber and rocket are consi-
dered.  The development of design criteria has been aided by
testing in the AFRPL pilot scrubber.

 APPROACH
      The approach to accomplishing the  objectives of the  pro-
 gram has been to develop design criteria for the largest
 rocket, 450,000 pounds (2 MN)expected to be tested.  The  ex-
 pected performance of the large rocket  has been used as the
 process baseline for the design criteria.   The largest rocket
 presented the most severe design problems and also caused the
 most pollution.  In developing the design criteria special
 care was taken to present equations, data, graphs, etc. that
 can be used also for designing smaller  scale scrubbers.
      Because cost estimates based on conventional designs

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have reached exorbitant  levels,  design criteria for an uncon-
ventional type of  scrubber  as  well as a conventional type have
been developed.  The  unconventional design represents the least
expensive design that could feasibly scrub the rocket exhaust.
This design  is a higher  risk (i.e., possibly less efficient)
since  it has not been tested.
REQUIREMENTS

      Specific  requirements for the research were set forth in

the  contract under the "Scope of Work".  The work to be accom-
plished  was defined as:

           The  contractor shall assist the EPA in carrying
      out Phase II of Interagency Agreement EPA-IAG-R5-0644
      by  preparing design criteria for scrubbing systems to
      handle  the exhaust from solid rocket motors ranging
      in  thrust from 5,000 pounds to 450,000 pounds.

      The "Scope of Work" listed seven tasks that were used as

guidelines and requirements for preparing the design criteria:

      1.   The  tradeoffs of scalability, cost, scrubbing,
          efficiency, feasibility and availability of
          equipment shall be considered in selecting the
          system components.  The contractor shall gene-
          rate  the information necessary to support their
          selections.

      2.   The  contractor shall determine the flow rates
          for neutralizing solutions and system cooling.
          The strength of the neutralizing agent used for
          removing the pollutants from the exhaust products
          shall be given as a percentage of the total
          neutralizing solution.

      3.  , The contractor shall make an estimate of the
          amount of electrical power required for opera-
          tion  of the motors, pumps, precipitators, etc.,
          used  in the system.

      4.   The contractor shall make an estimate of the
          quantity and type of fuel that may be required
          for any burners.

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      5.   The  contractor shall evaluate and  recommend the
          treatment requirements for the scrubbing media,
          and  any resultant by-products, and its  ultimate
          disposal process to meet ground water  disposal
          standards and/or satisfy Class I disposal re-
          quirements .

      6.   The  Air Force Rocket Propulsion Laboratory
          (AFRPL) will, within the limits of their system
          (5,000 Ib thrust rocket exhaust scrubber) and
          with any minor modification required,  perform
          specific tests or series of tests  requested by
          the  contractor to obtain data which  will aid the
          performance  of this work.

      7.   The  contractor shall study theoretically the
          coupling effects between the rocket  motor and
          the  scrubber system.  The EPA will  choose the
          exhaust scrubber coupling method which  has the
          least  effect on the thrust measurement  of the
          rocket motor and minimizes the exhaust  emissions.
          This recommendation shall also consider cost
          against efficiency.

     The  "Statement of Work" also set forth  other require-
ments :

     8.   At the completion of all of the preceding work,
          the  contractor shall make an oral  presentation
          to the AFRPL to present their findings  and re-
          commendations.   They shall present  supporting
          data  to define the choices of components and
          processes  for the exhaust scrubber.  The contrac-
          tor  shall  present the basis for their  scaling
          factors.   If the existing AFRPL 5,000  Ib thrust
          exhaust  scrubber cannot demonstrate  feasibility
          and  scalability, the contractor shall present
          their  recommendation for an intermediate system.
          The  conceptual  design of the intermediate system
          shall  be  detailed as to design cost  estimate,
          data requirements, reasons for choice  of the
          particulate  thrust level capability  of  the faci-
          lity,  a  preliminary construction cost estimate,
          the  proposed test program and a time phases
          schedule  for events from design start  to comple-
          tion of the  required test program.   The AFRPL
          will,  from this presentation, make  the  decision
         as to whether  design and build contracts  will
         be required  for the intermediate  system  (thrust
         level specified by  the  contractor).

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The contractor shall provide  a  preliminary cost
estimate for construction  of  the  facility to be
designed.

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                          SECTION 2
                   BACKGROUND INFORMATION

     The background information needed to design a scrubber
 consists of a complete description of both the process pro-
 ducing the effluent and the effluent itself.  This section
 includes a description of solid propellant rockets and their
 static testing.  A brief discussion of supersonic flow is
 given since this type of flow is not usually encountered in
 industrial applications of scrubbers.  Finally a description
 of the flow conditions and composition of the solid rocket
 exhaust is presented.

 SOLID PROPELLANT ROCKETS
     The characteristics and design of solid propellant rockets
 are presented in several standard textbooks.  Among them are the
 texts of Sutton (1963) and Hill and Peterson (1965).   The in-
 formation presented in the following paragraphs was derived pri-
 marily from these two references.
     The propellant of solid rocket engines is contained in the
 combustion chamber itself.  The propellants burn on exposed sur-
 faces to produce hot gases which produce a reactive force when
 expanded through the rocket nozzle.  Most solid rocket propel-
 lants contain all the materials necessary for sustaining com-
bustion.
     Propellants are usually distinguished as being homogeneous
or heterogeneous.   Homogeneous propellants contain fuel and oxi-
dant within the same molecule.  Double-base propellants (nitro-
glycerine-nitrocellulose) with small amounts of other additives
are the most common examples of homogeneous propellants.
     Heterogeneous mixtures of oxidizing crystals in an organic
fuel binder are called composite propellants.  The most common
oxidizer is ammonium perchlorate.  The fuel or binder is usually
 an organic polymer such as a polybutadiene or a polyurethane. Fine
                              6

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aluminum powder  is  often  used  in  propellants  because  of  its  high
heat of reaction with  oxygen and  its  relatively  high  density.
Additives  or  special binders containing  forms  of fluorine  have
also been  used to  increase  the energy of the  propellant.
     The most important performance characteristic  of a  rocket
is  the  thrust.   The following  equation describes the  three fac-
tors affecting the  thrust:
                        F = m  c*  CT                          (1)

where   F  =  thrust, N
        m  =  mass flow  rate  of  propellant,  kg/s
        c*  =  characteristic  velocity,  m/s
        CT  =  nozzle  thrust coefficient

Mass flow  rate is  a primary factor which determines the magni-
tude of the  thrust. The  characteristic  velocity is dependent
upon combustion  chamber properties, such as ratio of  specific
heats,  molecular weight,  and combustion  stagnation  temperature.
High temperatures  and  low molecular weights produce high values
of  "c*".   The thrust coefficient  is a function of nozzle geo-
metry only.   The thrust coefficient is a maximum when the  pres-
sure in the  exhaust plane of the  nozzle  is the same as the
ambient pressure.   Since  the nozzle throat is  choked,  the  flow
in  the  rocket nozzle is supersonic, which allows pressures
other than ambient  to  exist in the nozzle.

SUPERSONIC FLOW
     Flow  of  gases  at  speeds greater  than those  at which small
disturbances  (waves) are  propagated through a  compressible fluid
is called  supersonic.  The  speed of these small  disturbances  is
called the speed of sound,  a,  which,  in  an ideal gas,  is:
                          a = /yRT                          (2)
where   y = ratio of specific heats
        R = gas constant
        T = static temperature

                              7

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The Mach number, M, is the ratio of the fluid speed, u, to the
speed of sound:

                           M = H                            (3)
                               a
Supersonic flow in ducts has different characteristics than are
normally encountered with subsonic flow.  The primary difference
is that pressures within a supersonic duct are not influenced by
downstream disturbances since pressure waves cannot propagate
upstream in supersonic flow.  Thus, one can put any disturbance
whatsoever downstream of the rocket nozzle exit without affect-
ing the pressure in the rocket chamber or within the nozzle.
This last statement must be qualified if the boundary layer
thickness (which may be subsonic) is a significant factor of
the throat diameter.  It must also be recognized that supersonic
flow represents a larger than normally encountered kinetic en-
ergy.  This energy is manifested when one attempts to decelerate
supersonic flow in a duct.  For adiabatic flow without mass ad-
dition, a compression practically equal to the rocket chamber
pressure is required for deceleration to normally encountered
speeds.

STATIC TEST FIRINGS OF EXPERIMENTAL SYSTEMS
     Experimental rockets are tested extensively  in the evalua-
tion process of research, experimental development, and occa-
sionally, engineering development of weapon systems.  One phase
of the testing is the static firing, or firing while the rocket
motor is held immobile.  Overall performance of the rocket motor
can be determined by monitoring the thrust, chamber pressure,
and other variables.  Combustion temperature is usually so high
that it is not monitored directly.  Nozzle performance at high
altitudes is determined in special facilities capable of attain-
ing low pressures.   Because of the potential for uncontrolled
combustion (explosion),  the testing is conducted by remote con-
trol from a protected barricade or enclosure.

                               8

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COMPOSITION OF SOLID ROCKET EXHAUST
     The AFRPL tests solid rockets which use three different
types of propellants.  For purposes of design of air pollution
control equipment the relatively high energy formulations of
these propellants are used:

     1.  Conventional Composite
         a. Ammonium perchlorate with binder
         b. 21% by mass aluminum powder
     2.  Composite with Fluoride Additive
         a. Ammonium perchlorate with binder
         b. Fluoride additive
         c. 18% by mass aluminum powder
     3.  Double Base
         a. Nitrocellulose-nitroglycerine
         b. 19.5% by mass aluminum powder
     The chemical compositions of these three main types of high
energy propellants are given in Table 1.   The listed composi-
tions are based on thermochemical computer program predictions
supplied by AFRPL.  The table shows only those components which
are a significant mole percent of the exhaust.   Trace amounts of
various compoundsj elements, and radicals are also predicted.

FLOW CONDITIONS OF SOLID ROCKET EXHAUST
     Of the three types of propellants listed in Table 1, the
conventional composite type produces the most halogen acid gas.
This gas, as will be discussed in detail in Section 3, is
the one of primary concern.  Because the composite propellant
produces the most pollution of the three types,  the flow condi-
tions for it will be discussed in detail in the  following para-
graphs .
     The AFRPL has supplied a thermochemical computer program
prediction for the composite propellant rocket based on a

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          TABLE 1.  COMPOSITION OF SOLID ROCKET EXHAUST
Propellant
Component
CO
C02
HC1
HF
H2
H20(gas)
N2
A120a ,.„
Type
M.W.
28
44
36.5
20
2
18
28
102
Composite
Mole 1
19.4
1.4
17.1
0
29.4
12.6
8.7
11.4
Fluoride
Mole 1
28.9
1.2
3.1
1.9
26.0
5.8
23-8
9.3
Double B
Mole %
33.0
0.5
0
0
28.9
2.5
25.1
10.0
    (solid)
Compositions are based on thermochemical predictions at
approximately 2,400°K and 101 kPa (14.7 lbf/in2).
                              10

-------
chamber stagnation pressure of 6.89 MPa (1,000 lbf/in2) and a
nozzle area expansion ratio of 10.  A summary of the prediction
is given in Table 2.  Also included in the table are calcula-
tions of the molecular weight based on the composition listed
in Table 1, and the exit stagnation (total) conditions based on
a specific heat ratio of 1.15.  The nozzle exit stagnation pres
sure and temperature are not usually the same as the chamber
stagnation conditions because chemical reactions occur between
the chamber and the nozzle exit.  These exit conditions are
needed to extrapolate to other nozzle expansion ratios.
Volume Flow Rate
     Volume flow rate can be calculated from the ideal gas law:

                         m  £  R  T
                     Q =
                              pe
where   Q = volume flow rate, m3/s
       mT = propellant mass flow rate, kg/s
       f  = mass fraction of gases in exhaust
       R  = gas constant for exhaust gases, J/kg-°K
        O
       Te = exit temperature, °K
       p  = exit pressure, N/m2 (Pa)
For a rocket using the composite propellant with a specific
burning rate or mass flow rate a wide variety of volume flow
rates are possible.  This is because the exit temperature and
pressure can vary widely, depending on the design of the nozzle.
The nozzle is usually expanded to provide atmospheric exit pres-
sures; but sometimes in the interests of economy it is under-
expanded.  The stagnation pressure, which also affects the exit
conditions, can also vary widely.  Table 3 presents an estimate
of the extremes in ratio of the exit temperature to pressure,
which directly affect the volume flow rate.
     Based on these extremes in the factors affecting volume flow
rate, Figure 1 has been drawn.  This  figure presents the  actual
                               11

-------
  10,000
w
H
   1,000
     500
o
H

CJ
100
      50
      10
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                    Figure 1.    Rocket exhaust  flow rates.
                            12

-------
    TABLE 2.  FLOW CONDITIONS OF SOLID ROCKET EXHAUST -
              COMPOSITE PROPELLANT
                                   Metric
                   English
Assumptions:
   Chamber pressure
   Nozzle Expansion ratio

Predictions:
   Exit pressure
   Exit temperature
   Specific impulse
   Ratio of specific heats
   Chamber stagnation
         temperature
  6.89 MPa       1,000 lbf/in2
            10
   104 kPa          15 lbf/in2
 2,400°K         4,320 °R
 2,590 m/s         264 s
             1.15
 3,730°K
6,700°R
Based on composition listed in Table 1:
   Molecular weight of
       total exhaust
   Molecular weight of
       exhaust gases
   Gas constant of total
       exhaust
   Gas constant of exhaust
       gases
            29.2

            19.8

285 J/kg-°K      53.0 Ibf-ft/lbm-°R

412 J/kg-°K      76.6 Ibf-ft/lbm-°R
Based on specific heat ratio of 1.15:
   Exit stagnation (total)
       pressure
   Exit stagnation (total)
       temperature
7.23 MPa
4,170°K
1,050 lbf/in:
7,510°R
                              13

-------
TABLE 3.  ESTIMATED EXTREMES OF THE RATIO OF NOZZLE EXIT
          TEMPERATURE TO PRESSURE
Stagnation
Pressure
kPa (lbf/in2)

1
6
7
,890
,200
(1,
(2,
Stagnation
000)
500)
Exit Exit
Pressure Temperature
kPa (lbf/in2) °K(°R)
103
138
temperature
(15)
(20)
is
2
2
assumed
,400
,150
to
(4,
(3,
be 4
Ratio
°K/kPa f°R/ (lbf/in2)
320)
780)
,170°K.
23.3 (288)
15.6 (189)

 volume  flow  rate  for  the  composite propellant  as  a  function  of
 the  thrust level.  Also shown  is  the  total mass flow rate  from
 the  rocket based  on the specific  impulse of  2,590 m/s.

 Enthalpy  of  the Rocket Exhaust
      The  ideal gas enthalpy and enthalpy of  formation at 298°K
 for  rocket exhaust gases  can be found in Van Wylen  § Sonntag (1973)
 and  Section  3  of  Perry $  Chilton  (1973).  Table 4 presents some
 useful  data  from  these references.  For the  composite propellant
 listed  in Table 1 the static enthalpy at 2,400°K, relative to
 298°K,  is 96.0 kJ/gmol (22.9 kcal/gmol).  The  specific impulse
 is roughly equal  to the exit velocity of the gas.  A velocity
 of 2,590 m/s is equivalent to a dynamic enthalpy of 97.9 kJ/gmol
 (23.4 kcal/gmol)  for a molecular weight of 29.2 g/gmol.  The
 total enthalpy is thus almost equally divided  between the heat
 and  the kinetic energy of the rocket exhaust.  The total en-
 thalpy  relative to 298°K  (static) is the sum,  194 kJ/gmol  (46.4
 kcal/gmol).  Figure 2 shows the total enthalpy rate (or power)
 of rocket exhausts for rockets of various thrusts based on the
 composite propellant.
 Propellants Other than Composite Type
     The properties of the exhaust flow of rockets containing
propellants other than the high energy composite type are not
 significantly different.   For precise estimates of the volume
and energy flow rates one must follow the procedure used in  the
preceding paragraphs.
                               14

-------
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   5,000
   2,000







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c Impulse = 2,
ar Wt = 29.3
225,000 Ibf
239 kcal/s
3.41 x 106 BTU
r i 11 in












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        0.01    0.02
                    0.05    0.1    0.2

                           Thrust, MN
0.5
                        Figure  2.   Rocket  exhaust  power.
                                  15

-------
TABLE 4.   ENTHALPIES OF ROCKET EXHAUST COMPONENTS
 Enthalpy of Formation at 298°K
Component
CO
C02
H20
H20(l)
HC1
HF
Al203(s)
Ideal Gas Enthalpy
Component
CO
C02
H20
HC1
HF
Al203(s)
N2
H2
02
kJ/gmol
- 110.5
- 393.5
- 241.8
- 308.3
92.3
- 268.6
-1,610.
at 2,400°K,
kJ/gmol
71.35
115.8
93.60
69.*
69.*
272.8
70.65
66.91
74.49
kcal/gmol
- 26.417
- 94.054
- 57.798
- 73.678
- 22.060
- 64.20
-384.84
Relative to 298°K
kcal/gmol
17.052
27.674
22.372
16.5*
16.5*
65.210
16.886
15.993
17.804
  *Estimated
                       16

-------
ALUMINUM OXIDE PARTICLES
     While the total amount of the solid aluminum oxide par-
ticles produced was presented in Table 1, their size distribu-
tion was not.  Data on size distributions of aluminum oxide as
well as the adsorptive characteristics related to hydrogen
chloride and hydrogen fluoride will be discussed in this sub-
section.
Size Distribution
     Various techniques have been used to acquire data on
aluminum oxide particle size from solid rocket firings.   Prior
to 1970 most techniques were biased to the larger sizes be-
cause they were based on the settling of particles in dishes
downstream of the rocket nozzle.  The collected particles
were then analyzed by electron or optical microscopy.  A sum-
mary of these data is given in Radke et al. (1967).  These
data have been correlated by Philco-Ford into the following
equation based on throat size:

           d   = 1.439 + 6.2403xlO-1D - 3.0535 x 10"3 D2      (5)
            Jr o
                 - 4.278xlO'4D3 + 8.6194 x 10'6D4
                 - 4.634 x 10"8D5
where  d   = particle mean mass diameter, ym
        IT o
         D = throat diameter, inches (cm/2.54)
The standard deviation of the polynomial is 0.9 ym and the re-
lation is not reliable over a throat diameter of 226 cm  (89.1
in) or below a throat diameter of 2.4 cm (0.945 in).  This
correlation demonstrates that small rockets produce smaller
average size particles than do larger rockets.
     The standard deviations of the particle size  distributions
are not well characterized.  However, Radke et al.  (1967) pre-
sent data for a rocket with a throat diameter of 31 cm (12 in)
which has a geometric standard deviation of about 1.7 about
a geometric mass mean diameter of 6.9 ym.

                               17

-------
     In recent years emphasis has been placed on submicron
particles.  Nadler  (1976) reports that recent investigators
have found that solid rocket motors produce a bimodal distri-
bution of alumina.  One distribution peaks in the submicron
region and the other somewhere between 1 and 20 urn.  Since
most of the mass would be represented by the upper mode, the
above equation (5) probably still represents the mass distribu-
tion.  If number or surface area distributions were important
the smaller mode would be important.  Strand and Varsi (1974)
and others working for the National Aeronautics and Space Ad-
ministration (NASA) and Eisel et al. (1974) working for the
Navy have been taking fine particulate data from solid rocket
exhausts.
Adsorption of HC1/HF
     Nadler (1976) has performed experiments to determine the
amound of HC1 adsorbed on alumina from rocket exhausts.   He
found that a small amount (6.7 percent by weight) of the total
HC1 produced could be adsorbed on the A1203 in a typical rocket
exhaust.   Similar small amounts of HF adsorption are also
likely.

LIQUID PROPELLANTS
     Although the main emphasis of this report is the solid
rocket, the liquid propellant rocket should be mentioned.  The
exhaust products vary widely because the formulations vary.
The products may be very similar to those of solid rockets,
except for the aluminum oxide particulates.  The primary ex-
haust products are shown in Table 5 for two typical liquid
propellants.   The hydrogen fluoride is produced in higher per-
centages  by the F2- H2 liquid rocket than are either hydrogen
fluoride  or hydrogen cloride in most solid rockets.
                              18

-------
TABLE 5.  COMPOSITION OF LIQUID PROPELLANT ROCKET EXHAUST
Propellant Type:
Component          M.W.
CO                  28
C02                 44
HF                  20
H2                   2
H20  (gas)           18
N,                  28
Mole
77.9
22.1
N20i»- AZ - 50
   Mole %
    3.6
    3.8

   22.6
   39.9
   30.1
 Reference:   Garrett,  et  al.  (1972)
                               19

-------
                          SECTION 3
                EMISSION REDUCTION OBJECTIVES

     As of mid-1977, because of the intermittent and unusual
nature of experimental rocket testing, the local air pollution
control districts (APCD) had not specified any emission limits.
There had previously been some activity in a local APCD to set
some emission limits for rocket testing.   These limits were 400
parts per million by volume of hydrogen fluoride (HF) and SOO
parts per million of hydrogen chloride (HC1).   None of the other
exhaust products listed in Tables 1 and 5 were of concern.  Be-
cause the scrubber design required an emission limit objective
these limits on HF and HC1 emissions have been used.

REQUIRED EFFICIENCY
     The required removal efficiency of HF and HC1 based on the
composition of the rocket exhaust gas given in Table 1 is pre-
sented in the following table (6):
TABLE 6. REQUIRED REMOVAL EFFICIENCIES
Propellant Type
Composite
Fluoride Composite

Double Base
HC1
Vol.%
19.3
3.4

0
HF
Vol.%
0
2.1

0
Required
Efficiency,!
99.6 (HC1)
97.7 (HC1)
98.1 (HF)
0
      Volume  percent  as  given in Table 6 is different than mole
percent  in Table  1  since  the aluminum oxide particles do not
occupy a significant volume.   The  required efficiency was cal-
culated  from the  following equation:

                                20

-------
                                       x 100
where:
       nR = required efficiency, %
      £EL = fraction (or %) by volume emission limitation
       £  = fraction (or %") by volume of pollutant in
            effluent gas

     Particulate removal was not required under present rules.
However, because of the gas-liquid contacting required to re-
move the HC1 and HF gases a high percentage of particulates
will also be removed.  Calculations of predicted particle re-
moval efficiency, which will be presented later in this report,
show that for many rockets, removal of enough HC1 to meet emis-
sion limits will also result in removal of 99.9% by mass of the
particulates.

CONSTRAINTS ON CONTROL METHOD
     Since removal of large volumes of gas is required, it is
likely that an absorber type control device will be used.  Gas
absorbers usually use water as the absorbing medium, which means
that a supply of water will be required.  The use of water is
thought not to be a constraint because the use will be only
intermittent.
     Other possible physical constraints to be considered are
topographical.  Some test stands are situated on buttes and
hillsides which require special design considerations.
                                21

-------
                           SECTION 4
                  AFRPL PILOT SCRUBBER TESTS

 INTRODUCTION
     Part of the AFRPL program to reduce emissions resulting
 from experimental rocket testing activities was the evaluation
 of  a pilot scale scrubber sized for a 22 kN (5,000 Ibf) thrust
 rocket.  These tests were conducted by AFRPL test personnel
 under the direction of the AFRPL project engineers, Mr. J.
 Hewes and Mr. L. Sedillo.  The AFRPL made a limited but signi-
 ficant attempt to evaluate the pilot scrubber, however, incon-
 clusive results were obtained because of sampling difficulties.
 This effort represented a significant commitment of resources.
 For reasons that will be presented later the project was termina-
 ted. The testing program has been reported by Sedillo  (1978).
     The testing program consisted of nineteen liquid propellant
 and four solid propellant tests.   The initial test was conducted
 with liquid fluorine and gaseous hydrogen. An explosion of gases
 resulted in extensive damage to the demister section of the scrub-
 ber.  After extensive modifications the pilot scrubber was tested
 eighteen times using nitrogen tetroxide and aerozine 50,a 50-50
mixture of unsymmetrical dimethylhydrazine and hydrazine.  Many
problems during this period plagued this project.  This liquid
rocket test was used to establish what modifications and improve-
ments were required to satisfactorily test solid rocket motors.
     The main concern of this section is the analyses of the four
solid motor tests of 10 second durations conducted by the AFRPL.

 SOLID ROCKET EXHAUST  CONDITIONS
      The propellant used in the tests was a composite type
 containing 14% aluminum by weight.  The predicted exhaust
 composition is shown in Table 7 .
                                22

-------
TABLET. COMPOSIT
(Compos
Component
CO
C02
HC1
H2
H20 (g)
N2
Al203(s)
ION OF TEST
ite, 14% Al
M.W.
28
44
36.5
2
18
28
102
ROCKET EXHAUST
by weight)
Mole %
20.3
3.6
16.8
23.5
20.3
8.5
7.0
           Total  Molecular  Weight  =  27.0
           Gas  Molecular  Weight  =  21.4
       Composition is  based on AFRPL supplied  thermochemical
       prediction at  2,100°K and 101 kPa  (14.7 lbf/in2).
Other predicted flow conditions are listed in Table 8 .
TABLE 8 . PREDICTED AFRPL

Thrust
Specific Impulse
Mass Flow Rate
Volume Flow Rate, actual
Temperature, static
TEST ROCKET FLOW
Metric
22 kN
2,570 m/s
8.56 kg/s
45 m3/s
2,100°K
CONDITIONS
English
5,000 Ibf
262 s
18.9 Ibm/s
95,000 ACFM
3,780°R
SCRUBBER DESIGN
     The AFRPL pilot scrubber was a gas-atomized type, designed
by ARO, Inc. of Arnold Air Force Station, Tennessee and reported
in Garrett ,et al. (1972).  A sketch of the scrubbing system is
shown in Figure 3.  The scrubbing duct was steel, 0.91 m (3 ft)
in diameter and 9.1 m (30 ft) long.
                               23

-------
                                                             ENTRAINMENT SEPARATOR
to
            ROCKET
                              SCRUBBER DUCT
u
SPRAY
:HAMBEP
	






BAFl
                                 Figure  3.   AFRPL 22kN (5,000  Ibf) pilot  scrubber

-------
     The mist eliminator section was made of fiberglass rein-
forced plastic and packed with 2.54 cm  (1 inch) polypropylene
TellerettesQy to a depth of about 0.61 m (2 ft).  It was rec-
tangular and inclined 70° from the horizontal.  The mist elimi-
nator was designed by the manufacturer  (The Ceilcote Company,
Berea, Ohio) for 45.5 m3/s (120,000 CFM) and a total pressure
drop of 0.75 kPa (3 in W.C.).
     The outlet of the scrubber system was designed to facili-
tate sampling and was 2.13 m  (7 ft) in diameter.
     The liquid was supplied by a nitrogen gas pressurized
spherical stainless steel tank with a 75.7 m3 (20,000 gal)
capacity.  The liquid was a solution of potassium hydroxide
(KOH).  The effluent liquid was drained into an evaporation
pond.
     The inlet to the scrubber contained a cylindrical diffuser
with a converging conical inlet to capture the rocket exhaust
gases.  The diffuser was intended to aid in collecting the
lead and tail-off propellant gases by the aspirating action
of the liquid spray nozzles.  The  lead  gases were mainly those
emitted during ignition of liquid propellants for which this
pilot scrubber was also designed.  Another function of the
diffuser was to minimize blowback at tailoff.   It was doubtful
however that the diffuser had any effect on the scrubber oper-
ation other than to protect the rocket nozzle from splash of
the injected scrubber liquor at startup and tailoff.
     The diffuser did not reduce the flow to subsonic since
the back pressure in the scrubber was not nearly great enough
to cause a shock in the diffuser.  The velocity at the diffu-
ser exit was primarily governed by the area ratio between the
diffuser to the rocket nozzle throat.
     The scrubber liquid was injected through a number of
radially situated jet nozzles.  The nozzles were placed so
that the liquid jets were directed inward to penetrate the
rocket exhaust.  Some of the jet nozzles were inclined down-
stream to provide some pumping action during startup and
                               25

-------
 tailoff.  The tips of the nozzles were outside the expected
 exhaust plume.  The arrangement is shown in Garrett's report.

 DESIGN MODIFICATION - LIQUID INJECTION SYSTEM
      The only major design modification was a new liquid in-
 jection system.  Since the supersonic exhaust flowing from the
 diffuser was expanded only to about atmospheric pressure, the
 plume retained its diameter for many diameters downstream.
 This column of supersonic, high temperature gas would be ex-
 pected to melt any ordinary type of water injection nozzle
 located within it.  Injection of water at the flow center-
 line was prerequisite to obtaining the maximum transfer of
 energy and momentum from the exhaust gases  to the water. An
 injector that would not melt was one protected by a massive
 (high heat capacity)  angle iron bar that was also cooled by
 jets of water from the injector.
      The new injector design had three components:
   1. A stainless steel pipe with holes drilled in the side
      to act as nozzles.   The holes  would be  placed so  as to
      inject water  perpendicular  to  the  flow  from  the  diffuser
      and inject  water  toward the angle  iron  bar shield to
      help  cool it.
   2.  An angle  iron bar to  shield the  tube from the hot  gases
      flowing from  the  diffuser.
   3.  A  steel clamp  to  hold  the shield on the  tubing.   A  clamp
      is  used because the  shield  may have to  be  replaced  after
      one or more motor firings.
Figure  4 is  a  sketch of the  injector.   Three  of these  injectors
were  used  in place  of  the old jet nozzles.   Each  spanned the
full  diameter  of the spray  chamber  and  intersected the center-
line  of the  rocket  exhaust.  They were  rotated  60°  apart to
fully cover  the rocket exhaust.
      In addition to providing a  better  means  of injecting water
into the center of  the exhaust plume, the new  design caused
much greater mixing of the plume, the injected water,  and the
stagnant air.
                              26

-------
         HOLE IN SIDE OF PIPE
                   U-BOLT
                                    PLATE
 ANGLE IRON
                             PIPE
Figure  4.  Injector,
             27

-------
 Heat Transfer  Analysis  of the  Angle Iron Bars
      The  heat  transfer  to the  angles  which protect the water
 injection pipes  in the  new injector design was calculated.  The
 analysis  consisted of determining  the flow condition, estimat-
 ing the heat transfer coefficients, and  performing calculations
 based on  an  idealized model.
      The  flow  condition on the upstream  side  was assumed to be
 the condition  that occurred just behind  a normal shock in the
 rocket exhaust.   The  shock would be present because the super-
 sonic rocket exhaust  must slow down to stagnation at the tip
 of the angle iron.  The flow on the downstream side contained
 mixed liquid-vapor due  to boiling  on  the  surface.
      The  convective heat  transfer  coefficient  on the upstream
 side was  estimated from analysis of forced convection over
 exterior  surfaces  (Holman,  1972, Chapter  6):
                    0.246 k£
               hj =	  (Re,,)0-588  [Pr>33               (7)
where "f" denotes that the property is evaluated at the arith
metic mean of the free stream and surface temperatures.   For
this case:
               T£ = 2,500°K

and the following estimates are made:
Conductivity,  k£ = 4x10"" cal/s-cm-°C (0.1 BTU/hr-ft-°F)
Viscosity,     y£ = 6xlO"5kg/m-s
Prandtl No.,   Pr£ =0.8

The cross section diameter was:
                d = 0.0718 m

so, the Reynolds No.,  Re,£ = 323,000
                              28

-------
Substituting,  the  heat  transfer coefficient is:
                hi  =  0.022  cal/s-cm2-°C (160 BTU/hr-ft2-°F)
     On  the  downstream  side  the flow  of water contains mixed
liquid and vapor.  According to Figure 10-14 in Kreith (1965)
the most conservative heat transfer coefficient occurs at steam
qualities between  50 and 100 percent,  the coefficient at 100
percent  being  little different from the minimum.
The heat transfer  coefficient is then  (Kreith, 1965,  Chapter 8):
                                        0.2  /   \- I
                                           N
                    \_0.2  /   \-0.67           ,Q ,
= 0.023 C  p V  fRe, ,.\     /Pr,-\              (y J
 Here,           T£ = 500°K
                C  = 0.46 cal/g-°C
               Prf = 0.95
                 d = 0.65 cm
                y£ = 1.71 x 10"5 kg/m-s
                pV = 494 g/cm2-s
 So,           Re    = 1.90 x JO6
Thus,      h2 = 0.30 cal/s-cm2-°C  (2,200 BTU/hr-ft2-°F)
                               29

-------
The conductivity of steel at elevated temperatures is about:

            k = 0.083 cal/s-cm-°C (20 BTU/hr-£t-°F)

So that the heat transfer coefficient through the angle is:

                    hA - i                                 (9)

                         0.085  cal/s-cm2-°C  (t in cm)
                           t
                         £0  BTU/hr-ft2-°F  (t in cm)
                          t
The problem can be treated as a steady thermal resistance prob-
lem (Kreith, 1965, Chapter 1), with radiation neglected.  The
heat transfer per unit area is:
                         = hA

where the subscript "s" denotes the surface condition, and
          q = heat flow rate, cal/s
          A = area, cm2

The results were that, considering only heat transfer, the
thinner the better in this size range.  The choice of the
0.64 cm (0.25 in) thickness was made as being convenient to
use and stronger than the thinner angle.  The average tempera-
ture of the 0.64 cm (0.25 in) angle^was predicted to be 814°K
(1,005°F)  which means that, based on Sutton (1963), the strength
of 302 stainless steel was 71 percent of its room temperature
strength.   The stainless steel would be preferred even though
a slightly higher surface temperature would result due to the
lower conductivity for stainless steel than was assumed in
this analysis.  The mild steel angle should hold up at the
                              30

-------
reduced  strength  because  it  is  backed  up  by  the  pipe.   Because
the heat  transfer is  initially  unsteady the  time taken  to  reach
steady state  provides  a safety  factor.  About  eight  seconds
should be the time required  for temperature  buildup.

INSTRUMENTATION
     The  scrubber was  instrumented  to  determine  its  performance
characteristics.   Pressure,  temperature,  and flow rates were
monitored and 16  mm movie  films taken.

Static Pressure  (Refer to  Figure 5)
     1.   Rocket Chamber       - transducer
     2.   Diffuser Entrance    - oil-filled manometer
     3.   Spray Chamber       - oil-filled manometer
     4.   Scrubbing Duct       - oil-filled manometer
     5.   Entrainment  Separator- oil-filled manometer
The levels of the oil-filled manometers were recorded on 16 mm
movie film.

Dynamic and Total  Pressure
     A rake across  the flow  field was  located at  the downstream
end of the scrubbing duct to determine the velocity profile.
On one side of the  flow centerline there were ten impact (total)
pressure  tubes.   On the other side there were three pairs of
pitot-type tubes.   The pressure probes were  all  connected to
oil-filled manometers, which were filmed.

Liquid Pressure
     Transducers  were  used to monitor  the pressure of the
liquid lines  into  the  spray chamber.

Liquid Flow Rate
     The  flow  rate of  the liquid into  the spray  chamber was
monitored by  in-line flow meters.
                               31

-------
O  Manometer  probe
^  Total  and  dynamic  pressure  rake
A  Thermocouples
A  Thermocouples  added  for Test  4
 o  .         o
 Figure  5.  Location  of manometer probes  and thermocouples.

-------
Temperatures  (Refer  to  Figures  5  and  6)
     1.  Scrubber Duct          - thermocouples
     2.  Entrainment Separator  - thermocouples
     3.  At Exit                - thermocouples
It should be noted that thrust was not measured.

GAS AND  PARTICULATE  SAMPLING
     The sampling system was located at the system exit.  The
gas sampling was done with evacuated 1-liter bottles.  The
probes were arranged on a support frame as shown in Figure 7 .
The sampled gas was  analyzed by AFRPL using a mass spectrometer.
     The system for  collecting particles was designed by A.P.T.
and consisted of an  array of eighteen probes, precutter jars,
plastic  filter holders containing 47 mm Nuclepore R  filters at
the system exit, and polyethylene tubing connected through a
PVC pipe manifold to a Roots AF 22 blower.  The blower was
capable  of pulling about 0.47 l/s (1 CFM) through each of the
0.8 ym pore size filters.  The  suction through the filters was
begun and stopped by switching  the blower on and off at times
corresponding to start and shut-down of the rocket motor.  The
probes were located  just within the scrubber exit section.  A
0-1 atm  differential pressure transducer was placed on the in-
let connection to the blower to monitor the blower inlet pres-
sure (vacuum).  The  filters were returned to A.P.T. for gravi-
metric and optical sizing analysis.

MOMENTUM REDUCTION EXPERIMENTS
     The original test plan called for eight 30-second dura-
tion tests, six of which would  test momentum reducer plates.
These plates were made but no tests were made.  The principle
behind the use of the plates was described in the Phase I re-
port (Calvert and Stalberg, 1975).  The purpose was to achieve
a velocity reduction without having to expand the duct cross
section  or increase  the amount of liquid required.
                              33

-------
          ENTRAINMENT SEPARATOR SECTION
     UPSTREAM
  OF PACKING
DOWNSTREAM
OF PACKING
                                                      2>4>5
Figure 6.   Sketch of location of thermocouples
           in entrainment separator section.
                     34

-------
/ 7
O
8
O
9
0
13
\ o
                              DIA. = 2.1m (7 ft)
Figure 7 .   Location of filter holders in the
            entrainment separator exit.
                        35

-------
The plates would act as scoops to collect drops, thereby re-
ducing their forward momentum, and then reinject the drops into
the stream to absorb extra exhaust momentum.  The experiments
were planned to test this design and a flat, inclined, design
to determine if the reinjection was significant relative to
the fluid drag of the plates.
DISCUSSION OF TESTS

Test No.  1 - October 8, 1975
     On October 8, 1975 a 10 second duration solid rocket of
22,000 newton (5,000 Ibf) nominal thrust was fired into the
AFRPL pilot scrubber.  The rocket performed as expected, but
the cameras, set up to record manometer readings and to pro-
vide a visual record of what happened at the scrubber exit,
did not function.  Thus, no pressures were recorded that
would indicate what the conditions were inside the scrubber.
The most  unexpected event was the melting and smoldering of
the sampling apparatus set up at the exit of the scrubber.
The polyethylene lines to the gas sampling bottles apparently
melted through and let air into them.  The plastic filter
holders remained intact, although some were severely damaged.
The amount of flow through them was not known since some of
their polyethylene tubing also melted.

Test Data-
     The  recorded data are shown in Figures 8-11 .  The scrub-
bing duct wall temperatures were less than 100°C as predicted.
The scrubbing liquid flow rate was about 80 H/s which is about
9.3 times the rocket mass flow rate.

Post Test Inspection-
     The appearance of the apparatus after the test indicated
that a flame or other high temperature producing phenomenon
had been present.   The effects can be seen in the post-test
photo of the exit in Figure 12. The cameras, set up to provide
a  pictorial record of the test, did not operate, so it was not
certain that a flame was present.
                                36

-------
                                   9   10  11  12  13  14  15
01234567

                        TIME,  s

    Figure  8.   Test 1,  rocket chamber pressure.
                                                                       100
                                                                                                              1  -  Side, Mid-Duct    ::
                                                                                                              2  -  Side, End  of Duct :
01   2   3   4    5    6    7    8    9   10   11   12   13  14  15

                         TIME,  s

   Figure  9.   Test 1,  scrubber duct  wall  temperature.

-------
§
U
CO
        01   2   3   4   5   67    8    9   10  11  12  13  14
 rt
 D,
<
>
                                  TIME,  s


              Figure 10. Test 1, scrubbing liquid flow rate
       0   12   3   4   5   6   7   8   9  10  11  12  13  14  15

                                 TIME, s


            Figure 11. Test 1, sampling pump vacuum pressure.
                               38

-------
Figure 12.   Photograph of AFRPL scrubber exit after test 1
                               39

-------
     Heat  effects were physically evident in the upper  samplers.
Numbers  1,  2,  3, and  4 in Figure 7 showed the greatest  damage.
On  these samplers the following were observed:
  1. The polyethylene tubing had melted through.
  2. The polycarbonate plastic filter holders showed bubbling,
     melting,  and charring.
  3. The PVC reducing bushings were charred on one side.
  4. The red rubber hose connections showed signs of melting
     on  the outside and were sticky to the touch.
  5. The gummed paper labels were charred.
  6. The nylon tubing fittings were only very slightly  charred.
     To understand the causes of the damage to the apparatus
the following  data and observations were compiled:
  1. The temperature of the gas should have been no higher
     than  95°C since that was the temperature measured  up-
     stream in the mixing duct, and equilibrium theory  pre-
     dicted a  temperature of less than 100°C.
  2. The exhaust of the solid rocket tested contains about
     20 percent carbon monoxide and 2 percent free hydrogen
     (by weight).
  3. High  temperature effects were not observed for liquid
     rockets tested previously.  The liquid rockets tested
     previously produce very little CO and H2.
  4. The pattern of severest temperature effect was at  the
     top and along the outside of the scrubber outlet.
  5. Nylon, polycarbonate, and polyethylene have maximum
     service temperatures in the range 120-150°C.
  _. 	 	r	   filters, made of polycarbonate plastic
     and with a maximum service temperature of 140°C, did
     not appear to be affected at all.
  7. The polypropylene Tellerettes,  used as packing in the
     demister, showed no high temperature effects.
     The findings led to the conclusion that the most probable
occurrence was combustion of carbon monoxid^e and hydrogen either
during or just after the rocket firing.
                               40

-------
Weight Gain of Filters -
     The filters were weighed on a Cahn Model 4100 Electro-
balance before and after the test.  Although Nuclepore filters
are non-hygroscopic they were kept in a desiccant jar and also
baked at 80°C for a few hours prior to final weighing.  The
mean weight gain was 0.8 milligrams and the standard deviation
was 0.3 mg.  The largest gain (1.6 mg) was found on filters
4 and 10, and the least gain (0.5 mg) on filter 18 (see Figure 5)
     The pore size of the Nuclepore filters was 0.8 ym diameter
which means 1001 retention efficiency for particles larger than
0.8 ym diameter and somewhat less efficiency for smaller par-
ticles.
     Since some of the polyethylene tubes melted through,  pre-
sumably while the blower was in operation, the particulate mass
loading of the gas could only be estimated.  If burning of the
connecting tubing occurred after the end of the sampling period
the nominal flow of 0.47 £/s (1 CFM) for ten seconds  would have
passed through the filters and the mean particulate loading
would have been about 5 mg/m3 (actual) .   Converting to loading
in terms of dry normal volume it would have been 40 mg/DNm3,
since the temperature was about 92°C and the moisture content
was about 83% by volume.  This represents a particulate col-
lection efficiency of about 99.99% since the rocket produces
a particulate (AlzOs) loading of about 0.4 kg/DNm3.
Particle Size Analysis -
     Several of the filters were examined using a Nikon model
S-kt optical microscope.   The filters were placed on slides
and left uncovered during the examinations.  Although illu-
mination by transmitted light was available, the best illu-
mination occurred using reflected light.  Photo-micrographs,
using a Polaroid attachment, were made of selected particles.
     Most of the counting was done at 1500X with immersion
oil directly on the Nuclepore filter.  The sizing was accom-
plished by comparison with a Porton graticule incorporated
                               41

-------
 in  one  of  the  oculars  of  the binocular  eyepiece.   The  slide
 was  moved  randomly  about  the center of  the  filter  until  about
 200  particles  were  counted.  About 50 particles could  be  seen
 at  each location.   The smallest discernible particle corres-
 ponded  to  circle number 2 on the graticule, which  was  cali-
 brated  at  0.32 ym diameter.  Most of the large particles
 (greater than  1 ym) observed were agglomerates.  These ag-
 glomerates  appeared to be made up of spheres about 0.6 ym to
 0.8  ym  diameter fused  together.  The agglomeration appeared
 to  have occurred while the particles were very hot, and not
 on  the  filter.  The results of two counts are given in Fig-
 ure  13.  AS  can be seen, the count mean  diameter was around
 1 ym.
      Since  immersion oil was used for viewing at 1,500 power,
 comparisons of refractive indices between the particles and
 the  oil could  be made.  The particles had a higher refractive
 index than  the immersion oil (1.515).   This determination
 eliminates  the possibility of potassium chloride (r.i.=1.49) "
 and  strengthens the probability that the particles are alumi-
 num  oxide  (r.i. 1.70-1.77).

 Comparison  of Particle Size Distribution Reported  in
 the  Literature -
     Recent investigations of rocket exhaust alumina have been
 reviewed briefly by Nadler (1976) as reported in Section 2.
 He found that the size distribution may be bimodal, with one
 peak in  the submicron region and the other from about 1-20 ym.
 Small motors produce a smaller average particle size and a
 greater percentage  of the gamma (y) crystalline form of alum-
 ina, while larger motors produce a greater percentage of alpha
 (a)  alumina.  The rocket tested in this program had a throat
diameter of 5 cm,  which, according to data compiled by Radke,
Delaney and Smith (1967) should produce a particle distribu-
tion with a mass mean diameter of about 2.7 ym.  This size
should correspond to the distribution peak in the  large size
region,  although in rockets this small the peaks may not be
far  apart.
                               H" "

-------
     The data presented in Figure 13 for Test  1  show  a  geometric
count mean diameter of about 1 urn and a "a "  of about  2.35.
Converting to mass size distribution yields the value  of  9 ym
for the geometric mass mean diameter.  This high result for
the geometric mass mean diameter may be due to  "d  " and  "a "
                                                 o         6
values which are probably high because particles smaller  than
0.3 ym could not be seen so that bias toward  counting  the
larger particles existed.
     It should also be remembered that the data shown  here are
for a sample taken after the scrubber, which  should alter the
size distribution considerably.  The mean diameter should be
lowered by the scrubber which is more efficient on the larger
sizes.  The scrubber could have, however, caused agglomera-
tion.

Test No.  2  - February  27, 1976
     A second 10-second duration test in the AFRPL pilot scale
scrubber was made on February 27, 1976.   The  instrumentation
and cameras performed successfully.   The scrubber seemed to
work except that, again, there was heat generated somewhere in
the entrainment separator section.  This heat generation was
evident from thermocouple data and on the films.  The films
showed the polyethylene tubing, hung at the scrubber exit,
actually burning in flames.   Thermocouple data just upstream
of the entrainment separator packing (Tellerettes)  showed  tem-
peratures below 100°C.   Another unexpected occurrence was  the
absence of a steam condensation plume out the scrubber exit.
The only steam condensation cloud appeared on top of the en-
trainment separator section around the pressure relief hatches.
The gas coming out the  scrubber exit was very clear with no
visible plume.   Gas samples  were not taken.  The liquid (3% KOH)
flow rate was about 100 l/s.

Pressure and Temperature Data-
     The pressure transducer and thermocouple data are shown
in Figures 14-17. One side of the scrubbing duct was over-
                                43

-------

"-
5 4~
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2

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 5   10    20   30  40 50  60  70  80    90   95    98  99



        CUMULATIVE PERCENT BY COUNT




Figure 13. Test 1, size distribution data.





                  44

-------

to
a:
a:
w
09
               2   3   4   S   6    7   8   9  10  11  12  13  14  15


                                  TIME, s


                Figure  14. Test 2, rocket chamber pressure.
<
3
                                                                                                                    1    Right  Side,  Mid-Ductr
                                                                                                                    2    Left  Side, Mid-Duct
                2    3    4    5   6   7   8   9  10  11  12  13  14   15


                                   TIME,  s


             Figure  15.  Test  2,  scrubber  duct wall  temperature.

-------
  400
             UPSTREAM OF DEMISTER PACKING
             (AVG.  OF 5)
                   T [SEE  FIGURE 6
      01234567   8   9   10   11   12  13  14   15
Figure 16.  Test 2, gas temperature in entrainment separator from
            thermocouple arrays.
     01   23   4   5   6   7    8    9   10  11  12  13   14   15

                                TIME, s


           Figure 17,  Test 2, sampling pump vacuum pressure.
                              46

-------
heated, but this was due to improper arrangement of the liquid
injection pipes, which allowed deflection to one side.  This
was later corrected.  Arrays of thermocouples were placed in
the entrainment separator section to quantify the effects of
afterburning.  From Figure 16 it is apparent that afterburning
started about 4-5 seconds into the test firing and occurred
only downstream of the demister packing.  These high tempera-
tures occurred mainly in the gas stream since the walls and
packing were not significantly affected.

Velocity Profile at End of Scrubbing Duct-
     One of the cameras filmed the manometer board to show the
total, impact pressures across the scrubber at the end of the
straight mixing chamber section.  The pressure data are plot-
ted in Figure 18, and show a fairly uniform velocity except
at ±10% from the center.  The maximum peak velocity is esti-
mated to be about 200 m/s.

Particulate Data-
     Nuclepore filter (0.8 ym pore diameter) samples of the
exhaust were taken.  For this test aluminum tubing was used
instead of polyethylene tubing.   The mean loading was 1.66 mg
with a standard deviation of 0.78 mg.  It was estimated that
about 0.51 m3 of gas went through the eighteen filters so  that
the loading was approximately 33 mg/m3 at a temperature some-
where around 300°C.  These loading data are very approximate
because four of the filters had small visible holes due to
melting, and the temperature was not known with any accuracy.
Particle size distribution determinations were not made.

Chemical Analyses -
     The entrainment separator packing and sump liquid were
analyzed by AFRPL and the results were:
  1.  The surface of the Tellerettes was examined by scanning
     electron microscopy.  A small quantity of particulate

                               47

-------
     1.0
     0.5
 rt  -0.5
0,
 H
    -1.0
         0
0.2     0.4     0.6     0.8
RADIAL FRACTION FROM CENTER
1.0
     Figure  18.  Test  2, velocity profile  in  scrubber
                        48

-------
     matter was observed.  Aluminum  (Al or A1203)  and  silicon
     (sand) were the principal constituents.  Traces of  iron
     were also present.
  2. The liquid phase of the demister sump sample  contained
     high levels of potassium and lesser quantities of chlor-
     ide and aluminum.  The solids were primarily  aluminum,
     probably aluminum oxide.

Performance of Liquid Injectors -
     The liquid injector pipe/angle  iron system was photographed
after the test.  One of the photographs is shown as Figure 19.
Note that while the first angle iron burned off, the pipe was
still intact.  The second and third  angle irons were not
burned seriously.  It was decided to provide more  cooling to
the angle iron of the first injector by drilling more holes
in the pipe around the center of the duct.

Test No. 5 - November 5, 1976
     A third 10-second duration test was conducted with water
sprays on the entrainment separator packing.   The  sprays were
an attempt to dilute and cool the gas to prevent afterburning
downstream of the packing.  A total  flow rate of about 4 &/s
(66 GPM) of fresh water was used for the sprays.    Neither gas
samples nor particle samples were taken.  Again 113 &/s  (1800
GPM) of a 3% KOH solution was used.

Pressure and Temperature Data-
     The pressure transducer and thermocouple data are shown
in Figures 20-22.  Temperatures within the scrubber duct and
upstream of the entrainment separator packing were less than
100°C as expected.   Downstream of the packing the  thermocouple
data showed that some exit locations reached 650°C (1,200°F)
during the firing.   This was the first time the thermocouple
scale used was that high.   Previously, the readings had  gone
off the scale.  Even though the thermocouple registered  so
                                49

-------
AIR FORCE FLIGHT TEST CENTER
EDUARCS UK FORCE BASE,  CA
fiOMRECORD MATERIAL
OFFICIAL USAF PHOTOGRAPH
BY:  MR. RICHARD VAN-ETTEN
             Figure  19.    Photograph  of  liquid  injectors after
                              test  2.
                                            50

-------
t_n
        m
        e
        •P
        
-------

     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

-------
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

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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

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     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

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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

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                        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

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                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

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"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

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   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

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          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

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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

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                                                                        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

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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

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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

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   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

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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

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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

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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

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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

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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

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    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

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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

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  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

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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

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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

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     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

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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

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    Liquid Injection
    at Inlet to Duct
Figure 44.   Sketch of gas-atomized scrubber section.
                           122

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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).

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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

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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

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     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.

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                         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.

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                                                                 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.

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  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

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           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

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         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

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 Retaining
 Wall
                                            Supply
                                            Aqueduct

Figure 60.   Typical cross-section of a non-conventional
            absorber configuration.
                             153

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     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

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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

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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

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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

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 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

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     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

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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

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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

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Figure 61.    Sketch showing arrangement of tube bank
             separator.
                            163

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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

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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

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                        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

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           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.

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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

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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

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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

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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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                        REFERENCES
Abel, M.D. and Sutay, R.J., Verification Study of Air Pollu-
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AFB, CA, to be published.

Aviation Week and Space Technology, USAF to Ground-Test Toxic
Propellants, April 24, 1967, pp. 115-120.

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.
GARB Bulletin, 7(3):7, March 1976.

Calvert, S.  Source Control by Liquid Scrubbing, 2nd Ed.  In:
Air Pollution, A.C. Stem, ed.  1968.

Calvert, S., J. Goldshmid, D. Leith, and D. Mehta.  Scrubber
Handbook.  EPA Contract CPA-70-95,  NTIS/PB213016, 1972.

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
Scrubbers - Final Report. EPA-650/2-74-119-6, NTIS/PB248-050,
1975.

Chaput, L.S.  Federal Standards of Performance for New Stationary
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Chemical Propulsion Information Agency, JANNAF Handbook - Rocket
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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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                            180

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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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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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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
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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.
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Hill, P.G. and C.R. Peterson, Mechanics and Thermodynamics
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Holman, J.P.  Heat Transfer, 3rd ed.  McGraw-Hill, 1972.

Kemen, R.J.  Unpublished results of tests conducted in 1973 and
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Kempner, S.K., E.N. Seiler, and D.H. Bowman.  Performance of
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Kohl, A.L.,and F.C. Riesenfeld.  Gas Purification.  McGraw-Hill,
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Kreith, F.  Principles of Heat Transfer, 2nd ed.   International
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Lerman, S.  The Phototoxicity of Missile Exhaust Products: Short
Term  Exposure of Plants to HC1, HF, and A1203.  AMRL-TR-75-102,
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                             181

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Magill, P.L., F.R. Holden, C. Ackley, and F.G. Sawyer.  Air
Pollution Handbook.  McGraw-Hill, 1956.

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Perry, R.H., and C.H. Chilton.  Chemical Engineer's Handbook,
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Peters, M.S., and K.D. Timmerhaus.   Plant Design and Economics
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Popper, H. ed.   Modern Cost-Engineering Techniques.  McGraw-Hill,
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                            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

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     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

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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-3HAL
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
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


-------
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

-------
                         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

-------
                                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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