EVALUATION OF GREEN PROPELLANTS FOR AN ICBM POST-BOOST PROPULSION SYSTEM.pdf

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Propellant SYSTEM concept weight HTP fuel gas PBPS PROPULSION SYSTEM scenario solid Propellant OEC engine oxidizer Specific impulse
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Propellant SYSTEM concept weight HTP fuel gas PBPS PROPULSION SYSTEM scenario solid Propellant OEC engine oxidizer Specific impulse
UNCLASSIFIED
        UNCLASSIFIED
1
AN EVALUATION OF GREEN PROPELLANTS FOR AN ICBM POST-BOOST PROPULSION SYSTEM
Brian J. German*, E. Caleb Branscome*, Andrew P. Frits*, Nicholas C. Yiakas*, and Dr. Dimitri N. Mavris†
Aerospace Systems Design Laboratory
School of Aerospace Engineering
Georgia Institute of Technology
Atlanta, GA 30332-0150
ABSTRACT  *†
Propellant toxicity is a major concern in storing,
maintaining, and transporting strategic missiles.  Many
low toxicity “green” propellants have been developed
which hold the potential of increasing the safety and
lowering the operation and support costs of liquid-
fuelled strategic missile propulsion systems.  This study
evaluates several green propellants for use in a notional
next-generation post-boost propulsion system (PBPS).
The mission and physical dimensions for this PBPS
were defined by the requirements of the current
Minuteman III propulsion system rocket engine
(PSRE).  Possible propellants were initially screened in
terms of toxicity, performance, and technical feasibility
for the PBPS application with a multi-attribute ranking
method based on an overall evaluation criterion (OEC).
Promising propellants were identified, and candidate
PBPS concepts were developed and sized for each of
these propellants.  These concepts were evaluated in
terms of weight, cost, and technical risk to determine
which concepts, and hence propellants, show the most
promise for the application.  Probabilistic techniques
were employed to explore the effects of uncertainty in
the propellant performance and structural weight
estimates.  The results indicate that high-test peroxide
(HTP) combined with either an ethanol-based nontoxic
hypergolic miscible fuel (NHMF) or competitive
impulse non-carcinogenic hypergol (CINCH) is a very
viable propellant solution.
NOMENCLATURE
ACS
Attitude Control System
CDF
Cumulative Distribution Function
CINCH Competitive Impulse Non-Carcinogenic Hypergol
DACS
Divert Attitude Control System
HTP
High Test Peroxide (High Concentration H2O2)
Isp
Specific Impulse
HTPB
Hydroxy-Terminated Polybutadiene
MM
Minuteman
MM III
Minuteman III
MMH
Monomethyl Hydrazine
NIOSH National Institute for Occupational Safety and
Health
                                                          
* 
Graduate Research Assistant
† Director, Aerospace Systems Design Laboratory, Boeing Prof.
Copyright © 2000 by Dimitri Mavris.  Published by the Defense
Technical Information Center, with permission.  Presented at 2000
Missile Sciences Conference, Monterey, CA.
DISTRUBUTION STATEMENT: Approved for public release;
distribution is unlimited
NHMF
Ethanol Nontoxic Hypergolic Miscible Fuel
NTO
Nitrogen Tetroxide
OEC
Overall Evaluation Criterion
PBPS
Post Boost Propulsion System
PBV
Post Boost Vehicle
PE
Polyethylene
PSRE
Propulsion System Rocket Engine
REL
Recommended Exposure Limits
INTRODUCTION
As evidenced by the Minuteman III and Peacekeeper
systems, storable liquid rocket propellants have been
historically preferred for land-based ICBM post-boost
propulsion.  Typically consisting of a hydrazine-based
fuel and a nitrogen-based oxidizer, they have been
selected for many ICBM upper stage and spacecraft
applications because they offer very high performance
for 
storable 
(non-cryogenic and non-degrading)
propellants.  The specific impulse for many of these
formulations approaches 300 seconds.  As hypergolic
liquids (propellants that ignite automatically upon
mixing) these formulations also allow for the precise
thrust metering/impulse cycling necessary for the
stringent angular positioning and range maneuvering
requirements of the post-boost application.
These propellants have one major detriment, however:
toxicity.  Any significant exposure to either the liquids
or vapors can be extremely harmful or fatal.  A leak or
spill could be devastating both in loss of life and in
environmental damage.  The threat of a spill incident
mandates the need for costly safety systems and
procedures and causes concern in transporting the
propellants.  Additionally, planned future reductions in
exposure limits from the Occupational Safety and
Health Administration (OSHA) and the National
Institute for Occupational Safety and Health (NIOSH)
will further constrain the handling of the propellants.
Based on these concerns, many new lower toxicity
propellants have been undergoing development.  These
propellants 
include 
fuels 
such as dimethyl-2-
azidoethylamine, known by the 3M trade name
Competitive Impulse, Non-Carcinogenic Hypergol
(CINCH)1, and a doped ethanol Nontoxic Hypergolic
Miscible Fuel (NHMF)2,3 that may offer comparable
performance 
to monomethyl hydrazine 
(MMH).
Renewed interest in high concentration hydrogen
peroxide, also called high-test peroxide (HTP), as a less
hazardous oxidizer has also spurred significant
UNCLASSIFIED
        UNCLASSIFIED
2
research4.  These, in addition to inherently less
dangerous solid, hybrid, and gel formulations, are often
classified as “green” propellants because of the lower
environmental and personnel dangers that they pose
compared to other high-performance propellants.
Although less toxic, these propellants do not have a
proven record for ICBM application.  Research and
testing must prove them to be capable of hypergolic or
near instantaneous ignition, to provide high levels of
specific impulse, and to be storable without significant
degradation for a service life of as much as 30 years.
Hypergolicity is a challenge because many of the
proposed formulations require soluble catalysts to
provide the capability.  Storage life is a concern for
HTP because it degrades significantly over time at
normal storage temperatures.
This study evaluates the potential of the various green
propellants for use in a notional next-generation Post-
Boost Propulsion System (PBPS).  The authors assert
that the most effective method of implementing this
evaluation is through discerning the system level effects
of the propellant selection.  This evaluation is
accomplished by the following process:
• 
Identify propellant formulations that have the most
potential for the application
• Formulate and size a PBPS concept for each
propellant, 
incorporating proposed propulsion
technologies that may be required to realize system
performance
• Evaluate the concepts in terms of key metrics such
as weight, cost, and technical risk
To identify green propellants with the most potential for
the application, an Overall Evaluation Criterion (OEC)
method was used.  This method allowed the ranking of
propellant candidates based on qualitative and
quantitative attributes such as performance, system
complexity, and toxicity.  After comparing propellant
combinations through the OEC method, the most
plausible candidates from each propellant “family” (e.g.
liquid, solid, etc.) were selected for further analysis.  A
PBPS concept was sized for each candidate propellant
based on extrapolations from current systems.  These
concepts were then compared based on considerations
such as their total weight, cost, and possible technology
development issues.  These results were used to draw
conclusions on the applicability of the various green
propellants for a future PBPS system. This process is
summarized in Figure 1.
The current Minuteman III Propulsion System Rocket
Engine 
(PSRE) uses monomethyl-hydrazine and
nitrogen 
tetroxide, 
two highly 
toxic hypergolic
propellants that provide good specific impulse as well
as the rapid restart capabilities required for the precise
positioning of multiple reentry vehicles.  For the
purposes of this study, the MM III PSRE has been
taken as the archetype for the mission definition and
physical dimensions of the candidate concepts.
Propellant
Combinations
Overall
Evaluation
Criterion
Candidate Green
Propellants
Concept
Definitions
Sized
Concepts
Most Plausible
Green Propellants
Largest
OEC Value
Lowest
System
Weight,
Cost,
and Risk
Figure 1: Overview of Propellant Evaluation Method
PROPELLANT IDENTIFICATION
As a first step in identifying candidate green
propellants, extensive data available in the literature
were collected for a wide variety of propellant
formulations.  The data consisted of metrics that are
valuable for evaluating propellant applicability in the
areas of environmental and/or personnel hazard (i.e.
“greenness”), 
performance, 
and 
PBPS 
design
integration.   The following quantitative metrics were
compiled for each propellant formulation:
• Mixture ratio
• Densities of oxidizer and fuel
• Viscosities of oxidizer and fuel
• Vapor pressures of oxidizer and fuel
• Frozen equilibrium specific 
impulse for an
expansion from 300 psia to 14.7 psia
• National Institute for Occupational Safety and
Health (NIOSH) recommended exposure limits
(REL) for oxidizer and fuel
The NIOSH recommended exposure limits were used to
quantify propellant toxicity because these are the most
stringent government guidelines5.
In addition to these quantitative metrics, qualitative
rankings were established for certain characteristics that
are either not easily represented as continuous numeric
values or not widely available other 
than as
generalizations.  These rankings were specified as
integers ranging from one to five with 
three
corresponding to the metric ranking of the current MM
III PSRE, one corresponding to significantly worse than
the existing system, and five indicating significantly
better than the baseline.  These metrics include the
following:
• Storability of oxidizer and fuel
• Ease of inerting oxidizer and fuel
UNCLASSIFIED
        UNCLASSIFIED
3
• 
Ignition delay
• Complexity of implementing the propellant in a
propulsion system
Data was obtained 
for 68 distinct propellant
formulations (i.e. combinations of oxidizer and fuel at a
certain mixture ratio) that have been used in production
rocket engines or have been studied extensively in
laboratory experiments.  These formulations include
liquid (including mono- and bi-propellants), solid,
hybrid, and gel propellants.  In addition to green
propellants, traditional propellant formulations were
included for comparative purposes. 
 Cryogenic
propellants were eliminated from consideration because
of 
their 
incompatibility with 
the Minuteman
infrastructure and their very low storage lives, and
gaseous propellants were omitted because their very
low density make packaging impractical. 
 The
propellant data was culled from several sources
including technical reports and papers, the proceedings
of 
the Second International Hydrogen Peroxide
Propulsion Conference, the NIOSH internet web site,
and the Weight Engineers Handbook.  These sources
are listed as references 6-13.6,7,8,9,10,11,12,13
In order to identify the propellants that represent the
most suitable blend of performance and “greenness”, an
overall evaluation criterion method was used.  The
OEC is a single numerical metric that can be used to
compare alternatives.  The OEC is formulated such that
higher values represent more highly preferred solutions.
The core of the OEC method is its defining equation.
The equation consists of a sum of several terms, each of
which represents a particular characteristic upon which
the alternative is evaluated14.  Each term is formed as a
quotient comprised of the value of the respective
characteristic for a particular alternative and the value
for a chosen baseline alternative.  Because higher
values of the OEC are intended to represent more
favorable options, the numerator and denominator of
this quotient are interchanged as necessary to obtain the
desired direction of optimality.  For instance, if
increasing a certain metric is desirable, the value for the
alternative is placed in the numerator with the baseline
value in the denominator.  This formulation will yield a
value of the quotient greater than 1.0 for an alternative
with a higher value of the metric than the baseline.  The
ratio is inverted for metrics that are more desirable
when minimized.  The general form of the OEC
equation is shown in Figure 2.
OEC = α
Metric
Baseline Metric
 
 
 
 
 
 
 
 + β Baseline Metric
Metric
 
 
 
 
 
 + γ
Metric
Baseline Metric
 
 
 
 
 
 
 
 
Weights indicate importance of metric in decision
Maximize Metric
         Minimize Metric
             Maximize Metric
Figure 2: General Form of the OEC Equation
As shown in the figure, the terms of the OEC equation
are preceded by scaling factors.  These factors are used
to weight the importance of each term relative to the
OEC metric.  The weights are chosen as fractions
whose sum is 1.0.  By choosing such a format, the value
of the OEC for the baseline alternative is 1.0.
Alternatives with an OEC value greater than that of the
baseline are more desirable, while alternatives with
values less than the baseline are less favorable.
For the purposes of propellant selection, an OEC
equation was developed that is a function of the
previously identified propellant metrics.  The OEC is
shown in Equation 1 below.
Equation 1: Propellant Selection OEC
+






+






+






=
BL
BL
BL
Greenness
Greenness
Isp
Isp
 Weight
Specific
 Weight
Specific
OEC
γ
β
α




+




Delay
Ignition 
Delay
Ignition 
Complexity
 
Sys.
Complexity
 
Sys.
BL
BL
ε
δ
The specific weight and greenness terms shown in the
equation are defined as functions of multiple metrics.
The specific weight is the inverse of the weighted
average of the oxidizer and fuel densities.  As such, it is
a function of mixture ratio that indicates the volumetric
efficiency of the propellant combination.  Equation 2
shows the expression for the greenness metric.
Formulated as a general indicator of the propellant
environmental and personnel hazard, it is intended to
capture not only the direct toxicity of the propellant but
also the ease of containing and inerting any propellant
spills.  For this reason, it was specified as a function of
the REL to indicate toxicity directly and metrics such as
vapor pressure (VP) and ease of inerting to indicate the
hazard associated with a propellant leak.  The
logarithms of the REL and vapor pressure metrics that
appear within the equation are used to compress the
orders of magnitude variations that can occur in the
parameters for different propellants into a linear form
suitable for application in an OEC.
UNCLASSIFIED
        UNCLASSIFIED
4
Equation 2: Toxicity OEC Term




+




+




=




)
REL
Fuel
*
100
log(
)
REL
Fuel
*
100
log(
)
REL
.
Ox
*
100
log(
)
REL
.
Ox
*
100
log(
6
Greenness
Greenness
BL
BL
BL
γ
γ








+




+




+




BL
BL
BL
BL
Fuel
Inert
Fuel
Inert
Ox.
Inert
Ox.
Inert
)
VP
Fuel
log(
)
VP
(Fuel
log
)
VP
Ox.
log(
)
VP
Ox.
log(
This OEC method for propellant selection is not unique
and does not necessarily assure the best selection.  It is
only one of several decision methods for ranking
alternatives based on subjective and objective data.  It
also suffers from some shortcomings including the
presumption that all attributes are independent.  The
OEC method was chosen, however, because it was
tractable and provided good visibility and traceablity
for the propellant pre-screening process.
After developing the propellant evaluation OEC
equation, a series of weighting scenarios was developed
against which the propellants should be judged.  Each
scenario is comprised of a set of weighting scalars,
whose sum is 1.0, that indicate the relative importance
of each metric to the OEC.  Multiple scenarios were
used to prevent a biasing of the results caused by a
single particular choice of weightings.  This method
ensures that no propellants are eliminated from
consideration because of poor performance on only one
scenario.  Propellants were judged on their suitability
across the entire spectrum of scenarios studied.  Table I
shows the scenarios developed for propellant screening.
As shown in the table, the six scenarios range
successively in emphasis from performance to toxicity.
Table I: Propellant Evaluation Scenarios
Scenario 1
Scenario 3
Scenario 2
Scenario 4
Scenario 5
Scenario 6
Performance
Scenario
Toxicity
Scenario
Scenario
Specific 
Weight,
Isp,
Toxicity,
System 
Complexity,
Ignition 
Delay,
Performance (1)
0.3
0.5
0
0.1
0.1
2
0.3
0.4
0.1
0.1
0.1
3
0.3
0.3
0.2
0.1
0.1
4
0.25
0.25
0.3
0.1
0.1
5
0.25
0.25
0.4
0.1
0
Toxicity (6)
0.2
0.2
0.5
0.1
0
α
β
δ
γ
ε
Using the weightings within each scenario, the OEC
was calculated for each propellant within the database
using the current MM III PSRE propellants, NTO and
MMH, as the baseline.  From the results, the propellant
within each family with the highest average OEC value
across the scenarios was noted.  An example of this
“average” OEC performance for some of 
the
propellants examined is shown in Figure 3.  In this
excerpt from the OEC results, 
it 
is clear the
HTP/ethanol NHMF is, on average, superior to the
other propellants listed for the spectrum of weighting
scenarios considered.  The HTP/CINCH formulation
closely follows.  It is interesting to note that the
HTP/NHMF is superior to the NTO/MMH that is used
in the current PSRE in every case except for the purely
performance weighted scenario in which it trails only
slightly.  The CINCH formulation outperforms the
NTO/MMH even for the performance scenario.
0
0.5
1
1.5
2
2.5
3
Nitrogen
Tetroxide/MMH
98%HTP/
Ethanol
90% HTP/
Hydrazine
Whole Fuming 
Nitric Acid/JP-4
HTP/CINCH
OEC
Performance Scenario
Scenario 2
Scenario 3
Scenario 4
Scenario 5
Toxicity Scenario
Figure 3: Example of Average OEC Across the Spectrum
of Scenarios
The propellants within each family of similar concepts
(e.g., liquid, solid, etc.) with the highest average OEC
performance were selected as candidates for which a
sized PBPS concept would be developed.  These
propellants are the following:
• Ethanol Nontoxic Hypergolic Miscible Fuel
(NHMF) / High Test Peroxide (HTP) liquid
• Competitive Impulse Non-Carcinogenic Hypergol
(CINCH) / HTP liquid
• Monomethyl hydrazine (MMH) / nitrogen tetroxide
(NTO) loaded gel
• Polyethylene (PE) / HTP hybrid
• Aluminum / hydroxy-terminated polybutadiene
(HTPB) / ammonium perchlorate solid
Figure 4 shows the weighted values of each OEC term
for the selected liquid, hybrid, and gel propellants
evaluated for the performance and toxicity scenarios.
Because desirability relative 
to 
the baseline 
is
comprised of OEC values greater than 1.0, it is clear
that all of the selected propellants exhibit significantly
lower toxicity and, at worst, only marginally decreased
performance relative to the MMH/NTO liquid used in
the existing MM III PSRE.  These results indicate that
the selected green propellants should prove good
candidates for the PBPS application.
UNCLASSIFIED
        UNCLASSIFIED
5
0
0.5
1
1.5
2
2.5
OEC
Ignition Delay
System Complexity
Greenness
Specific Impulse
Specific Weight
98% HTP / Ethanol NHMF
Liquid
Gelled NTO/MMH
98% HTP / Polyethylene
Hybrid
Performance Scenario
Toxicity Scenario 
Performance Scenario
Toxicity Scenario
Performance Scenario
Toxicity Scenario
Performance Scenario
Toxicity Scenario
98% HTP / CINCH
Figure 4: OEC Component Breakdown for Selected Propellants
Current PSRE
1
2
3
4
5
Propellant
MMH / NTO
liquid
HTP / alcohol 
liquid
HTP / CINCH 
liquid
HTP / 
polyethylene
NTO / MMH
 gel
HTPB / ammonium 
perchlorate solid
Thrust metering 
system
valving
valving
valving
valving
valving
pintle/hot gas bypass 
system
Thrust 
vectoring
gimballed 
engine
gimballed 
engine
gimballed 
engine
gimballed 
engine
gimballed 
engine
gimballed engine
System type
ducted primary 
propellant 
thrusters
ducted primary 
propellant 
thrusters
ducted primary 
propellant 
thrusters
ducted primary 
propellant 
thrusters
ducted primary 
propellant 
thrusters
ducted hot gas 
thrusters/DACS
Thrust metering 
system
valving
valving
valving
valving
valving
hot gas bypass 
system
Feed system
pressurized 
cold gas
pressurized 
cold gas
pressurized 
cold gas
pressurized 
cold gas
pressurized 
cold gas
no feed system
Pressure tank 
shape
cylindrical
spherical
spherical
spherical
spherical
no tank
Fuel tank shape
cylindrical
spherical
spherical
cylindrical
spherical
cylindrical thrust 
chamber
Oxidizer tank 
shape
cylindrical
spherical
spherical
spherical
spherical
integral to fuel 
tank/thrust chamber
Post Boost Propulsion System Concepts
Propellant 
Storage 
Assemblies
Primary 
Propulsion 
System
Attitude 
Control 
System
Propellant 
Expulsion and 
Distribution 
Systems
Figure 5: Matrix of Candidate PBPS Concepts
CANDIDATE CONCEPT DEFINITION
In order to develop a PBPS concept for each propellant,
several alternative subsystem design and technology
concepts were explored.  The subsystem categories
include the primary propulsion system, the attitude
control 
system, 
the 
propellant 
expulsion 
and
distribution systems, and the propellant storage
assemblies.  Structural support alternatives were not
initially explored during the development of PBPS
concepts. 
 This decision 
to 
isolate 
structural
technologies was made in order to allow comparison of
the concepts only through the implications of the
propellant selection on the subsystem and components.
For 
the 
purposes 
of 
developing 
conceptual
configurations, 
modified 
versions 
of 
the
aluminum/magnesium monocoque structure from the
existing MM III PSRE were assumed.
The principal design decisions in the development of a
PBPS configuration are those defining the primary
propulsion system.  The primary propulsion system
design decisions were classified as selection of the
propellant formulation, the impulse cycling method,
and the thrust vectoring technique.  Because concepts
were developed for each propellant formulation, the
UNCLASSIFIED
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6
decisions on impulse cycling and vectoring must
correspond to the particular propellant.
Based on these subsystem alternatives, a notional PBPS
subsystem configuration was developed for each of the
five selected propellants.  The five configurations are
outlined in Figure 5.  These configurations were formed
by considering the subsystem options that were
expected to yield the minimum weight and minimum
technological risk solutions.  For instance, all of the
configurations employing liquid propellants contain
spherical propellant storage tanks.  Trade studies
indicated that these tanks provided the minimum weight
solutions and could be packaged within the dimensional
envelope of the current MM III PSRE.  Also, the liquid
systems were all designed with cold gas expulsion
systems.  Cold gas systems were believed to provide
lower weight (relative to a pump) and more accurate
expulsion control (relative to a hot gas system).  For
thrust vectoring, all designs incorporate mechanically
gimbaled engines.  These systems were viewed to be
potentially lighter and lower risk than fluidic control,
jet tabs, or jet vanes.
The systems incorporating solid propellant components
have some significant differences from the liquid
propellant concepts.  For instance, both the hybrid
HTP/PE and the solid ammonium perchlorate/HTPB
system use cylindrical fuel storage assemblies that
serve as the combustion chambers.  The cylindrical
design allows for integration of a solid grain conducive
for even burning.  In the pure solid system, the oxidizer
consists of grains commingled with the fuel, while in
the hybrid system, the oxidizer is stored as a liquid in
separate spherical tanks.  These concepts also have
significantly different attitude control systems (ACS)
configurations than their liquid counterparts.  The
hybrid system uses the oxidizer as a monopropellant for
the attitude control engines.  The pure solid system
employs ducted hot gas from the solid propellant gas
generators and a divert attitude control system (DACS)
to provide a multi-tiered approach to attain precise
vernier positioning.
HTP/NHMF, HTP/CINCH LIQUID CONCEPTS
The schematic for the HTP/NHMF concept is shown in
Figure 6.  The HTP/CINCH concept is identical except
for the change in fuel (and the corresponding tank
sizing).  One of the major design challenges that must
be overcome in developing a feasible HTP system is the
mitigation of the oxidizer’s tendency to degrade.  High
purity HTP is required to maintain acceptable levels of
performance; however, HTP slowly decomposes to
water and oxygen.  This decomposition is a strong
function of temperature.  At a typical silo temperature
of 70°F, the HTP decomposition rate is between 1%
and 2% per year15.  With a 1%/year rate, HTP at an
initial concentration of 98% degrades to 72.5%
concentration within the 30 year storage life expected
for the next generation PBPS.  If it is maintained at a
temperature 
near 
25°F, 
however, 
the 
yearly
decomposition rate is reduced to approximately 0.01%.
This rate of decomposition means that an initial
concentration of 98% HTP would decompose to 94.5%
during the 30 year storage life.  This level of
degradation is manageable from a propellant sizing
standpoint.
One possible solution to hold the hydrogen peroxide
temperature near 25°F during nominal silo storage is to
attach solid-state cooling devices, called thermo-electric
coolers (TECs) to the oxidizer tanks.  These devices are
commercially available, inexpensive, lightweight, and
have demonstrated 30 years of continuous operation in
space-based applications15.  Due to their low weight
and cost, the TECs could be made multiply redundant.
ACS engines
Mechanically gimballed 
axial nozzle
Oxidizer
Manifold
Insulated
H2O2
Tanks
Gas
Storage
Assembly
Fuel Manifold
NHMF
(Ethanol) Tank
Thermoelectric
devices and
heat sinks
Vent valve
(1 of 2)
Figure 6: HTP/NHMF Candidate Schematic
NTO/MMH LOADED GEL CONCEPT
The NTO/MMH gelled propellant concept consists of a
single spherical oxidizer storage assembly, a spherical
fuel storage assembly, and a spherical gas storage
assembly in addition to the axial and attitude engines.
One of the key drivers of the gelled propellant design is
the high system pressure that is required. Because the
gels are thixotropic (they become liquid under
pressure), high pressures are needed before the gels will
flow through the propellant ducting.  In order to handle
the system pressures, both the gas and propellant
storage assemblies and the propellant distribution
system must be strengthened, adding significant weight
to the structure.
UNCLASSIFIED
        UNCLASSIFIED
7
Another unique challenge of the gelled propellant
concept is design of the engine propellant valves.
Although loading the gelled fuel with aluminum
particulates increases the propellant performance, the
particulates have a tendency to clog valves and
injection orifices.  Because of the technological risk
associated with a large loading fraction, the concept for
this study employs only the minimal loading level
required to negate the performance degradation due to
the presence of the gellant material.  This low level of
loading should preclude the need for complex valve
assemblies, but some changes over a liquid system may
prove necessary.  If future studies indicate that valve
clogging will be an issue, a valve design employing a
translating centerbody with a wiping capability could
be incorporated.
HTP/POLYETHYLENE HYBRID
The primary assemblies comprising the configuration of
the HTP/polyethylene hybrid concept are three HTP
oxidizer storage assemblies, a gas storage assembly,
and a fuel storage/combustion chamber.  The necessity
for three oxidizer storage assemblies is driven by the
high 
value 
of 
the 
optimum mixture 
ratio.
Correspondingly, the fuel storage/combustion chamber
is comparatively small.
The primary advantage of a hybrid propulsion concept
is 
that 
it 
incorporates many of 
the favorable
characteristics of both liquid and solid propellant
systems.  As with liquid systems, hybrid engines allow
impulse cycling. 
 This cycling is achieved by
incorporating a throttleable valve that controls the
oxidizer injection into the combustion chamber.  The
hybrid system gains two critical advantages over liquid
rocket systems.  These are the increased density and
decreased environmental hazard of the fuel.  Because of
the increased density, more fuel mass can be packaged
in a smaller volume.  This characteristic is especially
advantageous for a volume-constrained upper stage
application such as a PBPS.  The solid fuel has a very
low environmental and personnel hazard both because
it cannot leak and flow and because it has negligible
toxicity.
In order to facilitate thrust chamber packaging, a
circumferential swirl injection combustion chamber
design was adopted.  This design concept, similar to the
Surrey Vortex Flow “Pancake” Engine16, allows a low
aspect ratio construction because oxidizer is injected at
ports 
located around the circumference of 
the
combustion chamber rather than in one.  The injectors
are oriented such that the liquid stream has a significant
circumferential component.  This orientation causes the
oxidizer to “swirl” around the combustion chamber
toward the center in a similar fashion as water drains
from a sink.  As combustion occurs, the hot gas is
expelled through the centrally located nozzle throat.
The general configuration of this concept is shown in
Figure 7.
Chamber
Nozzle
Fuel
Grain
Injection Velocity
Vectors
Oxidizer 
Injectors
Figure 7: Circumferential Swirl Injection Thrust
Chamber
This “swirl” effect has additional advantages.  The
circumferential component of the oxidizer injection
velocity serves to improve propellant mixing and to
reduce the chamber size necessary to achieve a
residence time greater than the ignition delay.  The
swirling also results in a layer of cool oxidizer on the
periphery of the combustion chamber.  This layer
shields the chamber walls from high heat loads.
Another unique feature of the HTP/PE concept is the
design of the attitude control engines.  Hybrid systems
are often characterized by poor ignition delay times, so
the development of an effective ACS engine may
involve significant technical risk.  An ACS thruster
concept that uses HTP as a monopropellant may prove
more risk averse.  The hydrogen peroxide fuel can be
used as a monopropellant by catalyzing the HTP such
that it exothermically decomposes into water vapor and
oxygen gas17.  The resulting hot gaseous products can
be expelled through a nozzle to produce the required
thrust.  Though the reaction time of this catalytic HTP
decomposition would not be low enough to ensure an
adequate ignition delay and minimum impulse for the
ACS engines, an integral accumulator reservoir within
the ACS engine has been incorporated for storing
catalyzed decomposition products. The presence of this
tank within the thrusters allows the HTP to be catalyzed
in a quasi-steady manner such that hot gas is always
available for rapid engine firings.  This thruster
configuration is depicted in Figure 8.
UNCLASSIFIED
        UNCLASSIFIED
8
Accumulator Reservoir
Hot Gas Valve
Propellant Valve
Catalyst Bed
Nozzle
Figure 8: HTP Monopropellant ACS Engine
Although the HTP monopropellant thruster concept
should overcome the issue of ignition delay for the
ACS engines, ignition delay could remain a problem for
the axial engine.  Many current hybrid rocket engines
have ignition delays of several seconds.  Such delays
are unacceptable for the PBPS application.  This
concern is amplified by the fact that hybrid motors have
not been demonstrated for upper stage application.  An
additional concern for the hybrid concept is in
packaging the combustion chamber.  The low aspect
ratio of the chamber allows it to fit within the outer
PBPS geometric envelope, but sizing studies have
indicated that a purely cylindrical chamber may intrude
into the volume occupied by the guidance set gyro
platform of the current MM III.  This intrusion would
necessitate a more complicated combustion chamber
geometry.  All of these concerns indicate that a hybrid
concept would entail significant developmental risk.
SOLID PROPELLANT CONCEPT
Figure 9 depicts a schematic of the aluminum/
HTPB/ammonium perchlorate solid propellant concept.
The concept consists of three primary solid propellant
gas generators, three vernier positioning gas generators,
and a final positioning divert attitude control system.
This 3-tiered propulsion scheme is the key attribute of
the solid propellant concept.
Valves
ACS Engine
Primary Gas Generators
(with thrust termination
capability)
Vernier Positioning Gas
Generators
(with thrust termination
capability)
DACS
Primary Propulsion / ACS
Radial Vent Nozzles
Mechanically Gimballed
Axial Nozzle
Valve
Hot Gas Manifold
Figure 9: Solid Propellant PBPS Concept
To attain acceptable levels of pointing accuracy and
precise downrange and crossrange velocity increments,
a PBPS design must allow high fidelity thrust control
for both the axial and the attitude control engines.  This
thrust control is characterized by a small minimum
impulse capability, low ignition delay, and impulse
cycling capability.  A high level of thrust control is
difficult to achieve in solid propellant systems because
the engines fire at a designed burn rate until all of the
propellant is expended.  Although some control may be
afforded by designing the propellant grain structure to
vary thrust over a certain profile during the burn, the
exact burn profile requirement for a given PBPS
mission cannot be known a priori and certainly cannot
be generalized to a single grain design.  This difficulty
arises from the flexibility that the PBPS must afford: it
must be capable of correcting any positioning errors
resulting from the boost phase of the mission, and it
must allow independent targeting of three RVs over a
wide range of flight paths.  To allow this flexibility
within a solid propellant system, a multi-tiered
“tunneling” approach was devised to attain accurate
positioning.  The system envisioned for this study
consists of three discrete levels of propulsive capability
in which each subsequent tier provides more precise
thrust control that can correct for the positioning errors
imposed by the previous propulsive level.
The first propulsive tier is provided by three primary
gas generators.  These solid propellant thrust chambers
are fired to produce major velocity increments for bulk
downrange and crossrange maneuvering.  One primary
chamber is provided for each RV.  The chambers are
connected to the attitude control and axial nozzles via
high-temperature hot gas ducting.  During a burn, the
effluent products from the chamber provide both axial
and attitude control 
thrust. 
 Should the thrust
requirements at a given point during the burn be less
than the capability of the chamber, the excess hot gas is
expelled through a series of radially-oriented vent
nozzles.  These nozzles are placed symmetrically
around the PBPS circumference and opened in unison
such that the net radial thrust is negligible.
Three vernier positioning gas generators comprise the
second tier of the propulsion system.  These chambers,
similar to those employed by the Trident SLBM18, are
fired to allow positioning during “coast” phases
between major downrange or crossrange maneuvers.
Although connected to the same hot gas manifold as the
primary gas generators so that both axial and ACS
impulse can be provided, the vernier positioning
chambers are sized primarily for the ACS duty cycle.
The final level of propulsive control is afforded by the
divert attitude control system.  This system consists of
UNCLASSIFIED
        UNCLASSIFIED
9
multiple banks of small solid propellant impulse
charges that are oriented in a pattern similar to that of
the 10-engine primary ACS of the current MM PSRE.
These charges are used to provide precise positioning of
the post-boost vehicle (PBV) immediately prior to RV
release.  Because the impulse of these charges is small
and can be accurately predicted, multiple charges can
be fired immediately before each release to damp out
angular positioning errors and to allow PBV back-away
from the RV without applying any significant jolt to the
RV being released.
Although this solid propellant concept offers significant
advantages in propellant “greenness”, it has several
disadvantages.  One issue is weight.  The hot gas
ducting and valves needed for this concept must be
fabricated from materials capable of withstanding high
temperatures and pressures.  For this reason, the hot gas
distribution system will be heavy.  The use of multiple
combustion chambers also amplifies structural weight
significantly.  In addition to high weight, the concept is
also very aggressive from a technology standpoint.
Although some elements of the concept have been
demonstrated in the Trident SLBM system, the
technological 
risk associated with 
the hot gas
distribution system and the DACS is significant.  Also,
a solid propellant system has not been demonstrated for
the stringent land-based ICBM accuracy requirements.
Another disadvantage of the solid propellant system is
that the propellant is a dangerous ordnance item.
CANDIDATE CONCEPT SIZING
Candidate concepts were sized by using a rocket
equation approach.  Each sized candidate concept was
provided with enough propellant mass to maintain the
same velocity budget as the current Minuteman III
system (including boost stages and the PSRE).  PBPS
empty weight was estimated by scaling the current
PSRE sub-system weights based on propellant mass
and pressure.  Figure 10 presents the methodology
formulated used for sizing the candidate concepts.
Assumptions for the concept sizing are listed in Table
II.  Additional details on the sizing process are
presented in reference 15.
Select a propellant for the
concept
Determine the theoretical
Isp
Calculate the required
propellant mass from the Rocket
Equation
Determine the required
components and draw
a schematic
Estimate the component and
total system masses
Has the structural mass
estimate changed?
Yes
Develop a packaging
 concept
No
Draw an accurate
 concept layout
Can the system
 be packaged into the
required volume?
No
Yes
Converged Design
Figure 10: Concept Development Process Flowchart
Table II: Table of Sizing Assumptions and Characteristics
Sizing Characteristic
Sizing Assumptions
Chamber pressure for liquid
concepts
Fixed for all concepts at 125 psia
Chamber pressure for solid,
hybrid, and gelled concepts
Fixed for all concepts at 600 psia
Expansion ratio/exit pressure
Fixed to estimated baseline expansion
ratio of 26 for all except hybrid, which
was sized for fixed length nozzle with
correspondingly higher expansion ratio
Structural weight
Fixed for all concepts
Propellant storage assemblies
Scaled according to required chamber
pressures, fuel volumes, mixture ratios
and densities
Gas storage assemblies
Scaled according to required chamber
pressure
Fittings (pressure transducers,
filters, squibs, etc.)
Weight fixed, but quantity adjusted
according to concept
Examples of concept layouts are given in Figure 11.
Converged weight breakdowns for the systems are
given in Table III.
Main Engine Nozzle
Oxidizer Tank
Fuel Tank
GSA Tank
Oxidizer Tank
Guidance Intrusion
Vernier Gas Generator
(1 of 3)
Guidance Intrusion
Primary Gas Generator
(1 of 3)
Main Engine Nozzle
Figure 11: Layout of HTP/NHMF Concept (left) and Solid Propellant Concept (right)
UNCLASSIFIED
        UNCLASSIFIED
10
Table III: Converged Mass Breakdowns for the Candidate Systems
Current
MMIII PSRE
(Estimated) ‡
HTP/
NHMF
HTP/
CINCH
HTP/PE
NTO/MMH
 Gel
Solid
System Average Isp (sec)
280.2
269.8
273.9
274.1
284.7
264.7
Subsystem
Mass
(lbm)
Mass
(lbm)
Mass
(lbm)
Mass
(lbm)
Mass
(lbm)
Mass
(lbm)
Gas Storage Assembly
55.0
55.0
55.0
26.0
60.0
--
Internal Structure
27.0
27.0
27.0
27.0
27.0
27.0
Rocket Engine Assemblies / Thrust Chamber Assemblies
69.0
69.0
69.0
82.0
69.0
72.0
Pyrotechnic Command Cable
7.0
7.0
7.0
7.0
7.0
7.0
Thruster Actuation and Gimble Command Cable
14.0
14.0
14.0
14.0
14.0
14.0
Thermal Insulation Blanket Assembly
15.0
15.0
15.0
15.0
15.0
15.0
External Shell/Interstage Assembly
73.0
73.0
73.0
73.0
73.0
73.0
Propellant Distribution System
19.5
19.5
19.5
10.8
19.5
--
(Propellant Gas Combustion and Distribution System)
---
--
--
--
--
216.6
Oxidizer Propellant Storage Assembly
26.6
13.2
13.2
19.1
19.8
--
Fuel Propellant Storage Assembly / Solid Fuel Grain Casing
26.6
6.9
6.9
26.0
19.8
--
Fasteners / Miscellaneous
9.0
9.0
9.0
9.0
9.0
9.0
Oxidizer Coolant System
---
10
10
--
--
--
Total Structural Mass
341.7
318.6
318.6
295.9
333.1
433.6
Oxidizer Mass (Primary Gas Generator Propellant)
160.0
171.4
151.2
269.8
147.7
400.0
Fuel Mass (Secondary Gas Generator Propellant)
100.0
45.2
58.2
22.3
92.3
115.0
(DACS Propellant)
---
--
--
--
--
40.0
Total Propellant Mass
260.0
216.6
209.4
292.1
240.0
555.0
Total Mass
601.7
535.2
528.0
588.0
573.0
988.6
( ) indicates a subsystem that is only present in the solid propellant concept
‡ references 15,19, and 20
UNCERTAINTY ANALYSIS
A refined version of the HTP/NHMF concept with
additional weight savings achieved through modern
structural technologies was chosen as an example to
evaluate the effects of uncertainty in the sizing process
on total PBPS weight.  A probabilistic study was
conducted to capture uncertainty in both the structural
weight and the propellant specific impulse, as these
parameters are the primary drivers on system sizing to
meet the fixed performance requirements.  The specific
goals of the design study were as follows:
• Define reasonable bounds for uncertainty in the
specific impulse and structural weight
• Bound the uncertainty in the total PBPS weight
that might result from variation in the specific
impulse and structural weight from their presumed
values
• Ensure that the specific impulse and structural
weight uncertainty does not result in an inability to
create a sized design to meet the mission
requirements
The first step in the study involved establishing bounds
for the maximum expected uncertainty for the specific
impulse and structural weight estimates.  Because of a
high 
confidence 
in 
the 
predicted 
propellant
performance, a small uncertainty bound was placed on
the nominal Isp predictions.  This uncertainty bound was
set at ±2.5% of the calculated propellant Isp.
The structural weight estimates, however, are more
uncertain.  This uncertainty is partially a function of the
lower fidelity and predictive capability of the structural
weight estimation method used in preliminary sizing.
Another factor in the uncertianty is in the estimates of
the fixed weights (e.g. valves, ducting, etc.).  Because
this uncertainty is signficantly greater, bounds of ±20%
of the nominal structural weight estimate was chosen.
Since the actual structural weight and Isp values are
most likely close to the nominal value, and because the
likelihoods of errors being either high or low are likely
equal, the parameters were defined as normally
distributed random variables for the purposes of this
probailisitic design study.  In order to generate the
mean and standard deviation necessary to define these
variables, the nomial predictions were set as the means,
and the standard deviations were chosen such that the
predicted bounds in the uncertainties were set to
enclose ± 3 standard deviations.  An illustration of the
random variable definition for Isp is shown in Figure 12.
UNCLASSIFIED
        UNCLASSIFIED
11
Isp (sec)
263.1
-2.5%
-3σ
269.8
Nominal
276.6
+2.5%
 +3σ
Likelihood
Figure 12: Specific Impulse Random Variable Definition
After the specific impulse and structural mass were
defined as probability distributions, a Monte Carlo
analysis was coupled with the PBPS sizing algorithm.
In this analysis, the Monte Carlo simulation selected
10,000 discrete values of structural weight and specific
impulse according to their respective probability
distributions.  The simulation then generated a sized
PBPS concept for each of the 10,000 cases.  The
resulting PBPS total mass outputs from the sizing
algorithm were collected to produce a cumulative
distribution function (CDF).  This CDF indicates the
likelihood that the total mass is less than a specified
value.  The CDF for PBPS total mass is shown in
Figure 13.
0.00
0.10
0.20
0.30
0.40
0.50
0.60
0.70
0.80
0.90
1.00
400
420
440
460
480
500
520
540
560
PBPS Total Mass (lbm)
Probability
75%  probability < 510 lbm
99.79%  probability < 550 lbm
Figure 13: PBPS Total Mass Cumulative Distribution
Function
The probabilistic analysis shows that likelihood of the
PBPS weight being less than 550 lbm for the presumed
uncertainty in structural mass and specific impulse is
nearly 100%.  This result indicates that the HTP/NHMF
PBPS concept is certain to weigh less than the current
MM III post boost system given the assumptions for
structural weight and Isp made in the study.  The results
also indicate that a HTP/NHMF PBPS can successfully
be sized in the entire presumed range of uncertainties,
i.e., no combination of structural weight and Isp result in
an unreasonable fuel mass requirement.  Similar results
were obtained for the HTP/CINCH concept (i.e. the
weight was less than that of the current PSRE).  As
should be expected from the deterministic results
shown in Table III, the other concepts showed
significant probability of weights greater than that of
the PSRE.
CANDIDATE CONCEPT EVALUATION
From a weight 
standpoint, 
the HTP/NHMF,
HTP/CINCH, and NTO/MMH gel concepts were all
predicted to be comparable or superior to the baseline
PSRE.  The CINCH and NHMF were especially
promising because of the relatively high propellant
specific impulse values and the lower-weight spherical
propellant tanks.  As weight is often directly related to
cost, these systems may also have lower acquisition
costs.
Weight and cost are not the only discriminating factors,
however; technical risk is also a key consideration.  All
of the concepts presented have some degree of
developmental risk.  The HTP concepts would need
technology development and validation of the oxidizer
cooling system and the fuel catalysts, and the gel
concept would need to demonstrate a feasible
propellant distribution and valving system.  The most
risky concepts, however, are the hybrid and solid
propellant concepts.  The hybrid has two very difficult
technology challenges: proving that ignition delay can
be reduced to acceptable levels and demonstrating an
adequate HTP monopropellant thruster.  The solid
propellant concept must be shown capable of producing
the precise positioning control with a multi-tiered
attitude control propulsion system and must also use
aggressive materials technologies to keep the weight of
the system reasonable.
Based on these considerations, the HTP/NHMF and
HTP/CINCH concepts appear to offer the most promise
for the PBPS application.  This study has indicated that
the systems can be produced at a weight (and possibly
cost) lower than that of the current MM III PSRE, even
considering uncertainty in the sizing presumptions.  The
weight may be further reduced if other structural
technologies such as an isogrid shell and composite
interior structure are implemented.  In addition to their
weight advantage, the systems also appear to be
feasible in terms of technical risk, especially as
contrasted to the solid and hybrid propellant concepts
presented.
CONCLUSIONS
From this study, it is apparent that several “green”
propellants are adequate for the PBPS application.  The
most promising green propellant formulations appear to
UNCLASSIFIED
        UNCLASSIFIED
12
be HTP/CINCH and HTP/NHMF.  The propellants’
performance, which approaches that of the NTO/MMH,
results in sized PBPS concepts that have equivalent
performance and weigh less than current Minuteman III
PSRE.  These propellant formulations also have a
reasonable level of technical risk, mostly residing in the
development of soluble fuel catalysts that are required
for hypergolic 
ignition with HTP and in 
the
demonstration of a low-weight oxidizer cooling system.
Based on these results, it is clear that in addition to their
inherent advantages in storage, personnel safety, and
handling, green propellants offer 
the necessary
performance required for land-based strategic ICBM
post-boost propulsion systems.
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