U.S. patent application number 11/215475 was filed with the patent office on 2006-12-07 for propellant composition and methods of preparation and use thereof.
This patent application is currently assigned to Estes-Cox Corporation. Invention is credited to Edwin Duane Brown, Scott James Dixon, Barry Richard Tunick.
Application Number | 20060272754 11/215475 |
Document ID | / |
Family ID | 32297163 |
Filed Date | 2006-12-07 |
United States Patent
Application |
20060272754 |
Kind Code |
A1 |
Dixon; Scott James ; et
al. |
December 7, 2006 |
Propellant composition and methods of preparation and use
thereof
Abstract
The invention relates to propellant compositions comprising a
solid inorganic perchlorate oxidizing agent, a nitrogen-containing
fuel, and a burn rate catalyst. Such compositions may be used as a
propellant material, (e.g., in rocketry), a pyrotechnic material,
an explosive material, a light generating material, a heat
generating material, or a sound generating material.
Inventors: |
Dixon; Scott James;
(Colorado Springs, CO) ; Tunick; Barry Richard;
(Colorado Springs, CO) ; Brown; Edwin Duane;
(Rockvale, CO) |
Correspondence
Address: |
LOWRIE, LANDO & ANASTASI
RIVERFRONT OFFICE
ONE MAIN STREET, ELEVENTH FLOOR
CAMBRIDGE
MA
02142
US
|
Assignee: |
Estes-Cox Corporation
Penrose
CO
|
Family ID: |
32297163 |
Appl. No.: |
11/215475 |
Filed: |
August 30, 2005 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
10295308 |
Nov 14, 2002 |
|
|
|
11215475 |
Aug 30, 2005 |
|
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Current U.S.
Class: |
149/19.4 |
Current CPC
Class: |
C06B 45/10 20130101;
C06B 23/007 20130101 |
Class at
Publication: |
149/019.4 |
International
Class: |
C06B 45/10 20060101
C06B045/10 |
Claims
1. A rocket engine comprising: a case having a tensile strength of
less than about 5,000 KPa; and a composite propellant grain
disposed in the case, the composite propellant grain comprising an
oxidizer, a burn rate catalyst, and a binder comprising a vinyl
resin, and a polyfunctional ether.
2. The rocket engine of claim 1, wherein the burn rate catalyst is
acicularly-shaped.
3. The rocket engine of claim 2, wherein the burn rate catalyst
comprises iron oxide.
4. The rocket engine of claim 1, wherein the oxidizer comprises
potassium perchlorate.
5. The rocket engine of claim 1, wherein the polyfunctional ether
comprises a triglycidyl ether.
6. The rocket engine of claim 5, wherein the triglycidyl ether
comprises a polyepoxide resin.
7. The rocket engine of claim 6, wherein the binder further
comprises an amide.
8. The rocket engine of claim 7, wherein the amide comprises
dicyandiamide.
9. The rocket engine of claim 1, wherein the vinyl resin comprises
a hydroxyl-functional polymer selected from the group consisting of
vinylchloride, vinylacetate, hydroxyalkylacetate, and mixtures
thereof.
10. The rocket engine of claim 1, wherein the case has a bursting
strength of less than about 1,000 KPa.
11. The rocket engine of claim 1, wherein the case comprises paper
or paperboard.
12. The rocket engine of claim 1, wherein the case is comprised of
a material that is free of metals or metal alloys.
13. The rocket engine of claim 1, wherein the composite propellant
grain comprises: an oxidizer consisting essentially of potassium
perchlorate; a burn rate catalyst consisting essentially of
acicularly-shaped iron oxide nanoparticles; and a binder consisting
essentially of a vinyl resin and a triglycidyl ether cross-linked
with dicyandiamide.
14. A method of preparing a pressable composite propellant
formulation comprising: mixing a vinyl resin, a vegetable oil, and
a solvent to provide a binder mixture; adding a burn rate catalyst
to the binder mixture; adding an amide to the binder mixture; and
reducing the solvent from the binder mixture to about 7 wt % to
provide a mixed formulation.
15. The method of claim 14, wherein the vegetable oil comprises
epoxidized castor oil.
16. The method of claim 14, wherein the vinyl resin comprises a
hydroxyl-functional polymer selected from the group consisting of
vinylchloride, vinylacetate, hydroxyalkylacetate, and mixtures
thereof.
17. The method of claim 14, wherein the burn rate catalyst
comprises acicularly-shaped iron oxide nanoparticles.
18. The method of claim 14, wherein the amide comprises
dicyandiamide.
19. The method of claim 14, wherein the solvent comprises an
ester.
20. The method of claim 19, wherein the solvent comprises ethyl
acetate.
21. The method of claim 14, further comprising an act of drying the
burn rate catalyst prior to performing the act of adding the burn
rate catalyst to the binder mixture.
22. The method of claim 14, further comprising an act of extruding
the mixed formulation to form pellets thereof.
23. The method of claim 22, further comprising an act of drying the
pellets to reduce the amount of solvent present therein.
24. The method of claim 23, further comprising an act of encasing
at least one pellet in a case comprising a case material having a
tensile strength of less than about 2,000 KPa.
25. The method of claim 24, wherein the case material has a
specific gravity of less than about 2.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation-in-part application of
and claims the benefit under 35 U.S.C. .sctn. 120 to pending U.S.
patent application Ser. No. 10/295,308, entitled COMPOSITE
PROPELLANT COMPOSITIONS, filed Nov. 14, 2002, which is incorporated
herein by reference in its entirety.
BACKGROUND OF THE INVENTION
[0002] 1. Field of the Invention
[0003] This invention relates to propellant compositions as well as
to methods of preparation and use of such compositions and, in
particular, to composite propellant compositions suitable for solid
rocket motor and gas generating applications.
[0004] 2. Discussion of Related Art
[0005] Composite propellant compositions typically contain separate
fuel and oxidizer components that are intimately mixed. Black
powder, sometimes referred to as gunpowder is the oldest composite
propellant composition. Black powder is made with charcoal, sulfur,
and potassium nitrate (saltpeter). In black powder, potassium
nitrate functions as the oxidizer, while sulfur and charcoal (i.e.,
carbon) are the fuel components. Even though black powder has a
relatively low specific impulse, it has been used in rocketry and
pyrotechnics for centuries.
[0006] Black powder rocket engines (rocket motors) are usually
manufactured by pressing black powder into multi-layer paper
casings under high pressure. The rear section of the engine is
fitted with a nozzle through which exhaust gases escape and which
is typically made from heat-resistant materials. In a typical
rocketry application, an intimate mixture of 75% potassium nitrate,
15% charcoal, and 10% sulfur (by weight) is tightly packed into a
case or casing, usually a multi-layered paper tube. An electrical
igniter is used to ignite the rocket engine. Because the black
powder is tightly packed to a uniform density, it can burn evenly
and produces thrust as the hot expanding gases escape the rocket
engine through the nozzle. Because black powder is granular and
pressable, rocket engine production can be easily automated by
means of multiple feed, hydraulic pressing machinery.
[0007] Although black powder is relatively inexpensive and readily
available, it has a relatively low specific impulse. Other solid
composite propellants have been developed using ammonium
perchlorate, a fuel, and a binder. These composite propellants,
sometimes referred to as AP composites or castable composites, can
be formulated to produce greater energy and superior physical
properties; however, these improvements are achieved at the expense
of processing simplicity. By their nature, castable propellants are
high viscosity liquids that can be cast (poured) into large
diameter motor cases with relative ease; however, casting becomes
more difficult as case dimensions decrease. Castable compositions
also begin to cure the moment they are mixed; therefore, viscosity
is time dependent and continues to increase until the composition
can no longer be processed. In consequence, the casting process
generates a substantial quantity of "waste" material.
[0008] The majority of castable propellant compositions are based
on plural component, curable polymer systems consisting of a polyol
(resin) and a diisocyanate (curative). Diisocyanates exhibit high
rates of reaction with water; therefore, the first step in the
manufacturing process is to reduce the water content of all
propellant ingredients to near anhydrous levels, which in practice
is less than 0.02%. Once the raw materials are dried, with the
exception of the curative, all liquid ingredients are charged into
a vertical vacuum mixer and mixed under vacuum for a prescribed
period of time to remove dissolved gases and obtain a uniform blend
of ingredients. This is expensive and time-consuming.
[0009] To achieve useful levels of energy and density, castable
propellants are formulated in such a manner as to obtain high
solids loadings (82% to 88%), which is defined as the weight
percent solid ingredients to liquid ingredients. At these solids
loadings, great care must be taken to preserve the fluidity,
therefore, the castability of the mixture. This is accomplished by
grinding solid ingredients (usually the oxidizer) to a series of
predetermined and carefully controlled particle size distributions,
thereby decreasing the void volume between solid particles. In
doing so, the volume ratio of liquid to solids is maximized, which
reduces the end-of-mix viscosity to manageable levels.
[0010] Once the solid ingredients are properly ground and sized,
they are charged into the mixer in increments to avoid a
potentially dangerous condition referred to as "dry mix." Following
the final ingredient addition, the mixture is mixed under vacuum
for a specified period of time, usually between four and ten hours.
Upon completion of the mix cycle, the curative is added and the
entire composition mixed under vacuum for an additional hour. At
this point, the propellant can be cast directly into motor cases or
liners.
[0011] Due to the relative ease by which viscous liquids can entrap
air bubbles, castable propellants must be introduced into
individual motor cases under vacuum, or by bottom feeding by means
of a casting bayonet. This is a complicated, capital equipment
intense procedure that does not lend itself to high rate
automation. In addition, the long cure time at ambient temperature
(two to four weeks) results in an excessive amount of material and
floor space allocated to "work in progress." Beyond the financial
disadvantage, the material being held as "work in progress"
normally cannot be subjected to quality assurance testing until the
final cure takes place.
[0012] Rocket engines using castable composites are typically
assembled by machining the grain(s) to achieve the desired grain
geometry, including the core, fitting the composite grain(s) into a
special casing; inserting a nozzle through which exhaust gases
escape (typically made from thermoset plastic); and adding a
bulkhead closure that may contain a delay element and a cap to
strongly secure the assembly in the casing and maintain the
internal pressure necessary for operation. Manufacture and assembly
of known castable composite rocket engines is therefore labor
intensive, equipment limited, and difficult to automate.
BRIEF SUMMARY OF THE INVENTION
[0013] The present invention overcomes the deficiencies noted above
for black powder and castable composite propellant compositions,
and provides compositions with a functionally desirable specific
impulse, a high combustion temperature, and improved manufacturing
characteristics.
[0014] In one aspect, the invention relates to propellant
compositions comprising a solid inorganic oxidizing agent; a
nitrogen-containing fuel; and a burn rate catalyst.
[0015] The invention also relates to rocket engines (also referred
to as rocket motors) packed with a composition comprising a solid
inorganic perchlorate oxidizing agent, a nitrogen-containing fuel,
and a burn rate catalyst.
[0016] In accordance with yet another aspect, the invention relates
to high burn rate, high combustion temperature propellant
compositions comprising a solid inorganic perchlorate oxidizing
agent, a nitrogen-containing fuel consisting essentially of
dicyandiamide; a burn rate catalyst which is preferably an oxide of
copper, chromium, cobalt, manganese, iron, vanadium, or a mixture
thereof; and a binder. The burn rate catalyst of these compositions
may be particulate.
[0017] In accordance with still another aspect, the invention
relates to an energetic composition comprising a solid inorganic
oxidizing agent selected from the group consisting of potassium
perchlorate and ammonium perchlorate; a nitrogen-containing fuel;
and a burn rate catalyst, wherein said burn rate catalyst is
nanoparticulate.
[0018] In accordance with one or more embodiments, the invention
relates to a rocket engine comprising a case having a tensile
strength of less than about 5,000 KPa and a composite propellant
grain disposed in the case. The composite propellant grain can
comprise an oxidizer, a burn rate catalyst, and a binder comprising
a vinyl resin, and a polyfunctional ether.
[0019] In accordance with one or more embodiments, the invention
relates to a method of preparing a pressable composite propellant
formulation. The method comprises acts of mixing a vinyl resin, a
vegetable oil, and a solvent to provide a binder mixture; adding a
burn rate catalyst to the binder mixture; adding an amide to the
binder mixture; and reducing the solvent from the binder mixture to
about 7 wt % to provide a mixed formulation.
[0020] Described further herein are similar propellant
compositions, as well as specifics regarding the relative amounts
of the various components. In particular, the identities and
relative amounts of oxidizing agent, nitrogen-containing fuel, burn
rate catalyst, and binder in the compositions of the invention are
disclosed. Also disclosed are methods for making and using the
propellant compositions of the invention.
BRIEF DESCRIPTION OF THE DRAWINGS
[0021] The accompanying drawings are not intended to be drawn to
scale. In the drawings, each identical or nearly identical
component that is illustrated in various figures is represented by
a like numeral. For purposes of clarity, not every component may be
labeled in every drawing. In the drawings:
[0022] FIG. 1 is a graph showing the specific impulse of propellant
compositions in accordance with one or more embodiments of the
invention relative to black powder and typical castable composite
formulations;
[0023] FIG. 2 is a graph showing the density impulse of propellant
compositions in accordance with one or more embodiments of the
invention relative to black powder and typical castable composite
formulations;
[0024] FIG. 3 is a chart illustrating the preparation of propellant
compositions in accordance with one or more embodiments of the
invention;
[0025] FIG. 4 is a graph showing the baseline ballistic performance
at various chamber pressures (psia) of the composite propellant in
accordance with one or more embodiments of the invention wherein
I.sub.d refers to the Density Impulse, I.sub.sp refers to the
specific impulse (lb/sec), T refers to the combustion temperature
(.degree. F.), and MW refers to the exhaust gas molecular
weight;
[0026] FIG. 5 is a graph showing the ballistic performance at
various chamber pressures (psia) of black powder wherein Id refers
to the Density Impulse, I.sub.sp refers to the specific impulse
(lb/sec), T refers to the combustion temperature (.degree. F.), and
MW refers to the exhaust gas molecular weight;
[0027] FIG. 6 is a graph showing the specific impulse performance
of the composite propellant in accordance with one or more
embodiments of the invention relative to the amount of potassium
perchlorate oxidizer (KP) and fuel (DICY);
[0028] FIG. 7 is a graph showing the specific impulse performance,
at various chamber pressures, of the composite propellant
formulation in accordance with one or more embodiments of the
invention at various binder concentrations;
[0029] FIG. 8 is a graph showing the burn rate at various chamber
pressures of the composite propellant of the invention at various
levels of SICOTRANS.TM. L 2715 D iron oxide;
[0030] FIG. 9 is a graph showing the ballistic performance at
various chamber pressures (psia) of the composite propellant in
accordance with one or more embodiments of the invention wherein
I.sub.d refers to the Density Impulse, I.sub.sp refers to the
specific impulse (lb/sec), T refers to the combustion temperature
(.degree. F.), and MW refers to the exhaust gas molecular
weight;
[0031] FIG. 10 is a graph showing the ballistic performance at
various chamber pressures (psia) of the composite propellant in
accordance with one or more embodiments of the invention wherein
I.sub.d refers to the Density Impulse, I.sub.sp refers to the
specific impulse (lb/sec), T refers to the combustion temperature
(.degree. F.), and MW refers to the exhaust gas molecular weight;
and
[0032] FIG. 11 is a graph showing the ballistic performance at
various chamber pressures (psia) of the composite propellant in
accordance with one or more embodiments of the invention wherein
I.sub.d refers to the Density Impulse, I.sub.sp refers to the
specific impulse (lb/sec), T refers to the combustion temperature
(.degree. F.), and MW refers to the exhaust gas molecular
weight.
DEFINITIONS
[0033] As used herein, the term "plurality" refers to two or more
items or components.
[0034] The terms "comprising," "including," "carrying," "having,"
"containing," and "involving," whether in the written description
or the claims and the like, are open-ended terms, i.e., to mean
"including but not limited to." Thus, the use of such terms is
meant to encompass the items listed thereafter, and equivalents
thereof, as well as additional items. Only the transitional phrases
"consisting of" and "consisting essentially of" are closed or
semi-closed transitional phrases, respectively, with respect to the
claims.
[0035] As used herein, the term "propellant composition" should be
understood to encompass use of the claimed compositions in
rocketry, pyrotechnics, military weapons and ammunition, as well as
other applications, as propellant materials, as pyrotechnic
materials, as explosive materials, as light generating materials,
as heat generating materials, or as sound generating materials. As
also used herein, the term "energetic composition" is intended to
include the "propellant composition" defined above.
DETAILED DESCRIPTION OF THE INVENTION
[0036] This invention is not limited in its application to the
details of construction, compositions, and/or arrangements of
components or acts set forth in the following description. The
invention is capable of being practiced and/or of being carried out
in various ways or embodiments beyond those exemplarily presented
herein and in the accompanying figures.
[0037] Energetic compositions have differing properties which are
useful for specific purposes or applications. For example, some
energetic compositions may be used for destruction, others for
pushing or propelling, and still others for generating light or
sound, and so on. Most, if not all, energetic compositions contain
at least an oxidizer and a fuel, which may sometimes be in the same
molecule. A binder may also be included to bind or keep the
oxidizer and fuel together, as well as to prevent physical
degradation and to increase mechanical strength. To specify or
modulate the rate of oxidization of the fuel, and/or the compounded
propellant's sensitivity to pressure or temperature changes, a
burn-rate catalyst may also be added.
[0038] The individual components of the present invention were
selected based on several factors including, inter alia, cost,
availability, ease of use, and safety of handling. The propellant
compositions of the invention, also referred to as VULCANITE.TM.
EB-75 herein, are a significant improvement over both black powder
and known castable composites, especially in model rocketry
applications, because they can provide improved ballistic
performance, improved volumetric and mass efficiency, a beneficial
reduction of combustion products toxicity, and improvement in
production efficiency, as described in more detail below. Indeed,
propellant compositions in accordance with the present invention
facilitate higher levels of automation; therefore, they allow
greater production rates with a significant decrease in
work-in-process inventories.
[0039] When the selected finely ground chemical components are
mixed or compounded into an intimate mixture, the resulting
propellant compositions of the invention may be granulated,
pelletized, pressed or even "molded" via pressing or co-casting
with other binders or propellants. This attribute makes the
propellant compositions uniquely versatile for mass production with
automated pressing equipment or other manufacturing methods deemed
desirable. Because they can be processed in dry powder or granular
form, the propellant compositions of the present invention are
pressable, and are not subject to the viscosity, air entrapment, or
pot-life problems inherent to castable propellants. As such, rocket
engine production is easily automated by means of single or
multiple feed, hydraulic pressing machinery.
[0040] Particularly regarding its use in manufacturing of rocket
engines, the propellant compositions of the invention can be used
as granules or pellets which can then be pressed into the engine
casing, or they can be pressed into molds to form "grains," the
mass of solid propellant used in rocket engines or other
applications. Grains can then be inserted into engine casings or
sold separately, allowing the user to load the casing on an "as
needed" basis.
[0041] Grains or pellets of the compositions of the invention have
improved ballistic performance as discussed below, and will find
utility in a variety of other applications and devices, such as,
but not limited to, ignition transfer pellets used in piccolo
tubes, pyrotechnics, artillery shell igniters, hand signals,
maritime smokes and signals, expelling charges and various other
applications.
[0042] The propellant compositions of the invention may be used as
a high burn rate (at least about 5/10 inch per second at a chamber
pressure about 150 pounds per square inch absolute (psia))
propellant for rocket engines, including, for example, in model
rocket engines, because it allows easily producible grain
geometries and operation at reduced chamber pressures. The solid
nature of the propellant compositions of the invention allows
grains of various geometries to be produced using automated
procedures such as molding and pressing. Because the propellant
composition can be tightly packed or pressed to a uniform density,
it burns evenly. Pressed grains are typically of higher uniform
quality because consolidation under pressure prevents formation of
air bubbles or cavities, as can occur in liquid mixtures. Such
cavities are undesirable because they result in burn rate
variability. Depending on the grain geometry selected, the
operating chamber pressure can be modified, as is known in the
art.
[0043] The propellant compositions of the invention also have a
high combustion temperature (as high as about 4,000.degree. F. and
even greater). An advantageous consequence of this high combustion
temperature is that the composition is more efficient because the
higher temperature imparts greater energy to the propulsive gases.
As a result, less composition is required to produce the same total
impulse as a given weight of black powder, as shown in Table 2,
where the Isp of the propellant composition of the present
invention can be about 1.73 times that of black powder and can
produce a combustion temperature nearly 1.5 times that of black
powder. Generally, the combustion temperature is at least about
3,500.degree. F., and in some cases, the combustion temperature is
greater than about 4,000.degree. F. Less propellant mass is
required to produce the same total impulse as a given weight of
black powder; therefore, it provides an increase in volumetric
efficiency at comparable densities.
[0044] In addition, the propellant compositions of the invention
have improved ballistic properties. Compositions of the present
invention have a specific impulse of at least about 100 to about
120 seconds at about 100 psia (pounds per square inch absolute),
and thus are more energetic than black powder, which typically has
a specific impulse in a range of 80 to 101 at a pressure of about
100 psia. In selected embodiments, the propellant compositions of
the invention have a specific impulse of at least about 120% that
of black powder. See Table 2, below.
[0045] Specific impulse, Isp, may be calculated according to the
following formula: Isp = 1 g .times. ( 2 .times. .gamma. .gamma. -
1 ) .times. ( RT c M ) .function. [ 1 - ( P e P c ) ( .gamma. - 1 )
.gamma. ] ##EQU1## where [0046] .gamma. is the ratio of the
specific heats of the combustion gases (Cp/Cv), [0047] R is the
universal gas constant, [0048] T.sub.c is the absolute combustion
temperature, [0049] g is the gravitational acceleration constant,
[0050] M is the average molecular weight of the exhaust gases,
[0051] P.sub.c is the combustion chamber pressure, and [0052]
P.sub.e is the exhaust gas pressure at the nozzle exit.
[0053] To maximize specific impulse, the molecular weight of the
exhaust gas and the relative amounts of solid combustion byproducts
should be minimized, and the combustion temperature and ratio of
chamber pressure to exit pressure should be maximized. In rocketry
applications, especially in model rocketry cases, the ratio of
chamber pressure to exit pressure is typically determined by the
design of the rocket engine itself in conjunction with the
ballistic properties of the propellant.
[0054] Another benefit of the propellant composition of the
invention is the ease of obtaining reproducible performance. The
chemicals used are consistent in composition and purity and are
consistent in the amount of energy delivered per unit of material.
Additionally, the components of the compositions are less subject
to change from the absorption of moisture than, for example, the
components of black powder. These features represent a further
advantage over conventional composite propellant compositions and
black powder.
[0055] In comparison, black powder performance (burn rate, energy
produced, burn temperature, etc.) can vary significantly from batch
to batch. Primarily, this is because black powder is made with
charcoal. The properties of charcoal, which is manufactured from
wood, can vary greatly depending on the species of wood, where the
wood was harvested, climate conditions during the life of the wood,
and the temperature and manner in which the wood is converted into
charcoal. Furthermore, charcoal tends to absorb moisture from the
atmosphere. The use of black powder in the manufacturing of rocket
engines requires extensive batch testing of the powder prior to
production so that its ballistic characteristics are known. The
compositions of the present invention require less extensive
pre-manufacturing testing because the chemical components do not
vary. Consequently, the compositions of the present invention do
not exhibit the same batch to batch variability.
[0056] The compositions of the invention produce exhaust gases
(combustion gases) of low average molecular weight, less than about
45, and preferably about 39 to about 40. By comparison, the average
molecular weight of exhaust gases from black powder are higher,
about 48, and exhaust gases from a typical castable composite have
an average molecular weight of about 23. Lower average molecular
weight combustion products are easier to accelerate to higher
velocities, making the compositions of the present invention more
efficient propellants than black powder. Additionally, in some
rocketry applications, the invention provides less solid residue,
which can alter the performance of a rocket engine. Further,
reducing the solid residue also is desirable because it produces
less build-up on launch equipment.
[0057] The very low percentage of non-expandable solid by-products
in the exhaust stream, preferably less than about 5 mole %,
increases efficiency. By comparison, black powder engines typically
have 15.96 mole % solid by-products in the exhaust stream. Castable
compositions typically have about 1 mole % solid by-products. This
makes the propellant compositions of the present invention more
efficient than black powder in creating energy, with less
undesirable residue. Furthermore, fewer particulates in the hot
exhaust stream reduce the risk of fire to the surroundings in a
variety of applications.
[0058] The propellant compositions of the invention can also have a
high burn rate coefficient as compared to typical castable
compositions. The burn rate, r.sub.b, may be expressed as
r.sub.b=a(P.sub.c).sup.n,
[0059] wherein a is the burn rate coefficient, P.sub.c is the
chamber pressure, and n is the burn rate exponent. As shown in
Table 1, below, in comparison with typical castable compositions,
the propellant compositions of the invention combine a reasonably
low burn rate exponent with a relatively high burn rate
coefficient, as obtained from Crawford Bomb and Micro Motor
firings, as is known in the art. This combination of relatively low
burn rate exponent, less than about 0.5, with relatively high burn
rate coefficient, greater than about 0.1, is both unique and
beneficial. It is beneficial because, inter alia, it allows for
higher mass flow rates at low chamber pressures in a range of about
50 to about 200 psia, pressures at which known castable
compositions do not function reliably. TABLE-US-00001 TABLE 1
Comparative Burn Rates Burn Rate (above Burn Rate Burn Rate about
100 psia) Exponent Coefficient Black Powder r =
0.5974P.sub.c.sup.0.0789 0.05-0.11 0.50-0.70 Typical Castable r =
0.0482P.sub.c.sup.0.3607 0.32-0.70 0.02-0.06 Composition VULCANITE
.TM. EB-75 r = 0.1033P.sub.c.sup.0.3613 0.36-0.50 > about 0.10
Propellant Composition
[0060] The low burn rate exponent reduces variation in performance
based on pressure. A low burn rate exponent greatly reduces the
pressure sensitivity of the propellant, thereby allowing larger
burning surface area changes without extreme pressure increases.
This is beneficial for use with the materials typically used in the
construction of rocket engines, including, for example, model
rocket engines, where the inner diameter of the casing may vary
from casing to casing and even within individual casings. The lower
sensitivity can be especially significant at low combustion
pressures because any variation at such pressures can attenuate any
undesirable variations in burn rates. Thus, in embodiments of the
invention wherein low-strength casings are utilized, burn rate
exponent sensitivity can lead to unpredictable performance
characteristics. The burn rate exponent of compositions of the
invention may be less than about 0.5.
[0061] In accordance with some embodiments, the invention relates
to propellant compositions comprising a solid inorganic perchlorate
oxidizing agent; a nitrogen-containing fuel; and a burn rate
catalyst. The burn rate catalyst can be an oxide of copper,
chromium, cobalt, manganese, iron, vanadium, or a mixture thereof.
The burn rate catalyst is preferably a high surface area
particulate. Other embodiments of the invention pertain to rocket
engines packed with propellant compositions comprising a solid
inorganic perchlorate oxidizing agent, a nitrogen-containing fuel,
and a burn rate catalyst. In accordance with some embodiments, the
invention relates to high burn rate, high combustion temperature
propellant compositions comprising a solid inorganic perchlorate
oxidizing agent, a nitrogen-containing fuel consisting essentially
of dicyandiamide, a burn rate catalyst, and a combustible binder.
An advantageous example of the propellant compositions of the
invention is a high burn rate, high combustion temperature
composition comprising about 64 to about 72 wt % of a solid
inorganic perchlorate; about 15 to about 23 wt % of a
nitrogen-containing fuel; about 0.5 to about 10 wt % of an oxide of
one or more of copper, chromium, cobalt, manganese, iron, vanadium;
and about 0.75 to about 12 wt % of a combustible binder.
[0062] The burn rate catalyst can be instrumental in lowering the
burn rate exponent of the propellant compositions of the invention.
The catalyst is preferably a high surface area particulate, and is
typically irregularly-shaped. The burn rate catalyst typically has
a large surface area, which can increase the catalytic activity
thereof. The burn rate catalyst can comprise one or more inorganic
oxides of a metal. In accordance with some embodiments of the
invention, the average size of the catalyst is relatively small and
may thus be characterized as nanoparticulate. Thus, for example,
the metal oxides can be utilized in the compositions of the
invention as nanoparticulates wherein at least one dimension
thereof is in the nanoscale domain. The nanoscale dimension can be
a smallest dimension or a largest dimension of the catalytic
particle. Further embodiments of the invention contemplate the use
of a mixture of catalysts having a variety of shapes. For example,
some embodiments of the invention can include a burn rate catalyst
having an acicular shape in combination with one or more burn rate
catalysts having a globular shapes.
[0063] The propellant composition of the present invention can
comprise one or more thermally conductive species. In accordance
with some aspects pertinent to one or more compositional
embodiments of the invention, a particulate ingredient thereof can
provide or at least facilitate heat transfer during the combustion.
For example, the thermally conductive species can have a high shape
aspect ratio such that the species can be characterized as being
acicular. Such habits can facilitate thermal conductivity through
the propellant composition body, e.g. the grain, because a long
dimension thereof can thermally expose a region of the propellant
composition to a higher temperature when the bulk propellant
composition has a lower thermal conductivity, relative to the
thermally conductive material. Thus, the burn rate catalyst, in
accordance with some embodiments of the invention, can facilitate
desirable propellant characteristics by chemically and thermally
modifying the behavior thereof.
[0064] The surface area of the catalyst is preferably greater than
about 50 m.sup.2/g. The preferred catalysts are metal oxides that
can modify, e.g., accelerate, the reduction/oxidation (redox)
reactions associated with propellant combustion. An example of a
burn rate catalyst suitable in the compositions of the invention
include those commercially available as SICOTRANS.RTM. Red L 2715 D
iron oxide, available from BASF Corporation, Florham Park, N.J. In
some compositions of the invention, the burn rate catalyst may
constitute about 0.1 wt % to about 15 wt % of the composition and,
in some cases, may be about 0.25 wt % to about 10 wt %, based on
100 wt % of the propellant weight.
[0065] A binder may also be used in the propellant compositions of
the invention. The binder can comprise at least one resin material.
The binder can further comprise at least one material that
comprises at least one functional group that can react with one or
more components of the binder. For example, the binder can comprise
a compound having a functional group that reacts with the polymeric
precursor species and/or the at least one modifier. Indeed, the
binder materials can be selected to provide or facilitate the
transformation of propellant composition properties from a viscous
material, capable of being disposed into a case, to a dimensionally
stable solid propellant grain. The binder materials can be selected
to provide any further desirable propellant grain properties. For
example, the binder precursor materials can be selected to provide
a propellant grain having high burn rates at low chamber pressures.
Moreover, the species comprising the propellant compositions of the
invention can also be selected to provide desirable energetically
stable compositions. The binder of one or more propellant
compositions of the invention can provide a matrix that can serve
as carrier for the oxidizer and/or a fuel. In accordance with some
embodiments of the invention, the binder matrix can also provide or
serve as a fuel, or even as a fuel supplement, during combustion.
Typically, the binder matrix can also facilitate processing of the
composite propellant composition. Indeed, in some embodiments of
the invention, the binder matrix serves as a carrier for at least
one oxidizer and/or at least one burn rate catalyst such that the
composite can be pressed into a case or casing and/or form
propellant grains therein. In accordance with some aspects
pertinent to one or more embodiments of the invention, the binder
matrix can facilitate production of propellant grains by providing
compositions with desirable rheological properties during
preparation and processing operations thereof and further provide
propellant grains that have desirable mechanical properties during,
for example, storage, transport, and/or uses thereof. Thus, one or
more propellant compositions of the invention can have desirable
rheological properties during, for example, mixing and/or pressing,
and can further have desirable, stable mechanical properties during
propulsive service. For example, one or more components of the
binder matrix can reduce the viscosity of the mixture to facilitate
mixing.
[0066] The binder may prevent the propellant composition granules
from being degraded, e.g., broken down into potentially dangerous
dust during the manufacturing processes and/or during transport and
storage thereof. The binder may also increase the mechanical
strength of the composition, which, in some embodiments of the
invention, can be effected by chemical cross-linking. For example,
one or more ingredients or species comprising the binder can react
to form a cross-linked network, thereby imparting greater
mechanical strength to the propellant composition structure. The
binder may include one or more modifiers or be modified by an
ingredient to provide desirable processing characteristics. For
example, the binder may comprise one or more plasticizing
modifiers. The preferred binders are understood to assist, rather
than interfere with, the propellant functionality. Various
ingredients can be utilized in the binder matrix. For example, the
binder matrix can comprise a polymeric material or a precursor
thereof that forms or reacts to provide a dimensionally stable
propellant grain. Preferred binders and/or components thereof are
combustible and produce low molecular weight by-products, and are
typically combustible organic polymers. Unlike inorganic binders
such as silica, the preferred binders do not increase the quantity
of non-expandable solid combustion products. In some cases, the
binder can comprise one or more modifiers that can modify a
Theological property of a mixture thereof, e.g., reduce a viscosity
of the precursor mixture. Thus, in accordance with some aspects of
the invention, the modifier can serve, among other things, as a
plasticizer of the propellant composition to provide desired
rheological properties.
[0067] In accordance with one or more embodiments of the invention,
the binder can comprise at least one resin, at least one modifier,
and/or at least one reactive agent. One or more components
comprising the binder can include chemically reactive species. In
some cases, one or more components of the binder can include
functional groups, which can react with other functional groups
pendant on or a part of other components of the propellant
composition. Non-limiting examples of compounds, such as resins,
which can be utilized in the propellant compositions of the
invention, include those comprising one or more vinyl functional
groups and/or those comprising one or more hydroxyl functional
groups. Non-limiting examples of compounds, such as modifiers,
which can be utilized in the propellant compositions of the
invention, include functionalized species, e.g. having mono or
polyfunctional groups like polyfunctional ethers. Specific examples
of such polyfunctional ethers include di- and trigycidyl ethers or
epoxidized organic compounds.
[0068] The propellant compositions of the invention may be 0.5 wt %
to about 15 wt %, or more particularly 0.75 wt % to about 12 wt %
binder. The binder may be comprised of at least one organic polymer
alone or in combination with a plasticizer such as dioctyl adipate,
dioctyl sebacate, hydrocarbon ester tackifier, and combinations
thereof.
[0069] Materials of the binder can be selected to have a high
density, which can provide desirable ballistic performance. The
selected binder materials can also provide mechanically stable
grains, especially when loaded into paper motor cases. For example,
the binder materials can provide grains that can be readily pressed
into cellulose-based cases without crumbling to excessive dust. In
some cases, the binder materials can effect moldability of the
composite propellant mixture in the case. Such materials can also
accommodate automated pressing operations, e.g., by hydraulic
pressing. Examples of preferred binder materials include, but are
not limited to, those comprising alkyl acetate, vinyl acetate,
and/or vinyl chloride resins. Further, the binder materials can
include one or more plasticizing agents, tackifying agents, bonding
agents, and/or wetting agents. Such agents can also be reactive or
have one or more reactive functional groups. As discussed herein,
the functional group can react with itself and/or one or more other
components in the composition. The desirable binder materials can
provide combustion products having a low average molecular weight
and high combustion temperatures, especially with respect to black
powder. For example, some amides have a typical combustion
temperature of greater than about 4438.degree. F., which is
relatively higher than the combustion temperature of black powder
(about 3426.degree. F.). Further, the binder materials can be
selected to provide gaseous combustion products, especially with
the disclosed oxidizing agents discussed herein. The binder
materials can also be selected based on toxicity such that
preferred candidates are considered to be non-toxic, as defined or
regulated by, for example, governmental entities.
[0070] Non-limiting examples of components of binders effective in
the practice of the present invention include polyvinylchloride,
polyvinylacetate, polyvinylalcohol and/or copolymers thereof such
as the solution vinyl resins commercially available as UCAR.TM.
VAGH, VAGD, VAGC, and VROH resin, each of which is available from
The Dow Chemical Corporation, Midland, Mich., GEON-121 (available
from B. F. Goodrich Corp.), poly (2-ethyl-2-oxazoline), and epoxy
or acrylate resin, epoxidized trimethylolpropane, trimethylol
ethane triglycidyl ether, such as HELOXY.TM. Modifier-44 from Shell
Chemical Company or Resolution Performance Products, Houston, Tex.,
epoxidized soybean oil, and combinations or mixtures thereof.
Non-limiting examples of agents include flexibilizers or modifiers
such as those commercially available as HELOXY.TM. esters, e.g.,
HELOXY Modifier 505 castor oil polyglycidyl ether, from Resolution
Performance Products, Houston, Tex.
[0071] Non-limiting examples of compounds that can react with one
or more of the resin and the modifier include those that are
nitrogen-containing species, thus serving, at least partially, as a
fuel source. The fuel component of the propellant composition
typically reacts with the oxidizer to produce the propulsive
exhaust gas. In some cases, the fuel species can react with a
functional group of a component of the propellant composition and
provide the propellant composition with desirable mechanical, e.g.,
rheological properties, and also react during propulsive service
with the oxidizer. The nitrogen-containing fuel is generally a
cyano-, amide-, or amine-containing material, or a mixture thereof.
Some examples include acrylonitrile, amino tetrazole,
aminoguanidinium bitetrazole, ammonium dicyanamide,
bistriaminoguanidiniumdecacarborane, bis(trinitroethyl)nitramine,
calcium bitetrazole, dicyandiamide (or cyanoguanidine),
nitroaminoguanidine, triaminoguanidine, and
triaminoguanidinedicyanamide. The fuel component of the propellant
compositions of the invention may be about 10-30 wt %, and more
particularly, about 15-23 wt %.
[0072] An example of a composition according to the invention
consists of about 62-72 wt % potassium perchlorate, about 15-30 wt
% dicyandiamide, about 2-10 wt % iron oxide, and about 3-12 wt % of
a VROH base stock comprising a plasticizer and VROH mixed in a
solvent and subsequently evaporated. The VROH base stock mixture
comprises about 3:2 VROH to plasticizer (e.g., dioctyl sebacate).
More particularly the composition may consist of about 68 wt %
potassium perchlorate, about 19 wt % dicyandiamide, about 5 wt %
iron oxide, and about 7.5 wt % of a VROH base stock comprising a
mixture of about 58.8:41.2 VROH to dioctyl sebacate.
[0073] The oxidizer typically facilitates combustion, such as by
oxidation, of one or more of the ingredients of the propellant
composition. The oxidizer can comprise an oxygen-containing
compound. Typically, the oxidizer can comprise an ammonium, alkali,
or alkali metal salt of an oxygen-containing compound. In
accordance with one or more embodiments of the invention, the
oxidizer facilitates stable combustion or deflagration at low
pressures, typically at less about 200 psia. The oxidizing compound
can include those considered as suitable for high pressure
conditions and/or associated with high flame temperatures or high
burn rates. The oxidizer can also be selected based on material
density. Typically, oxidizing agents having a high density are
suitable in some formulations of the invention. Further, the
selected oxidizer is preferably non-hygroscopic to facilitate
storage and ease of use. For example, in accordance with one or
more embodiments of the invention, the oxidizer can comprise an
ammonium or alkali perchlorate. However, in some cases, the
oxidizer is selected to provide compositions having high burn rate
exponents and poor low-temperature ballistic properties. Although
such properties are typically undesirable in certain applications
of the invention, they can be addressed by the utilization of other
ingredients, such as the burn rate catalyst. The solid inorganic
oxidizing agent may be potassium perchlorate or ammonium
perchlorate. In some embodiments, the oxidizing agent is about 60
to about 75 wt %, or more particularly, about 64 to about 72 wt %
of the composition.
[0074] The propellant compositions of the invention can be prepared
by any suitable technique. FIG. 3 presents a flow chart exemplarily
illustrating acts that may be utilized to prepare propellant
compositions in accordance with one or more embodiments of the
invention. Thus, any suitable equipment may be utilized to prepare
the grain forming compositions of the invention. Further, the acts
of propellant composition preparation may be performed in any
suitable order or sequence. Indeed, in accordance with one or more
embodiments of the invention, the propellant compositions of the
invention can be prepared by utilizing pre-batching techniques
which typically involves preparing an aggregate of two or more
components of the composition. Further, where desired, or in some
cases, necessary, preparatory procedures can be performed to render
any ingredients suitable or active in the propellant
composition.
[0075] For example, a pre-batch or pre-blend can be prepared by
mixing one or more vinyl resins and one or more solvents, and,
optionally, one or more modifiers. Pretreatment or preparatory
procedures, such as drying, can be performed prior to or after the
pre-batch preparation. The one or more solvents can be any suitable
solvent that is selected to dissolve the resin. Examples of
suitable solvents include those that readily evaporate such as
esters, including, but not limited to, ethyl acetate. Any suitable
amount of solvent can be used. Typically the amount of solvent
utilized provides desirable pre-blend mixture rheological
properties. For example, the amount of ethyl acetate can be about
equal to the amount, e.g., the weight, of vinyl resin. In
accordance with some embodiments of the invention, the amount of
solvent utilized can be selected to provide a target pre-batch
viscosity while resulting a minimal amount, e.g., about 7 wt %, in
the aggregate propellant composition or mixed formulation.
[0076] The pre-batch mixture can then be charged into a mixing
apparatus wherein additional components or ingredients can be added
and mixed. For example, dried burn rate catalyst can be mixed with
the pre-batch in a vertical mixing apparatus until the burn rate
catalyst, e.g., acicularly-shaped, nanoparticulate iron oxide, has
been wetted by the pre-batch mixture. Drying of the catalyst can
performed at, for example, a temperature of about 350.degree. F.
for about three hours under a vacuum of about 29 inches of mercury
to achieve a moisture content close to about 0%. Examples of
suitable mixing apparatus include those commercially available from
Charles Ross & Son Company, Hauppauge, N.Y. Mixing can be
performed under vacuum. Any suitable vacuum level can be utilized.
For example, a vacuum of about 29 inches of mercury can be applied
during this mixing operation.
[0077] The oxidizer can then be added. One or more preparatory
procedures can likewise be performed on the oxidizer. For example,
potassium perchlorate can be ground to have a desired particulate
size, e.g., about 10 .mu.m, in a hammer mill. Further the oxidizer
can also be dried in substantially the same fashion as described
with respect to the burn rate catalyst.
[0078] One or more binder material-reactive components can be
added. For example, dicyandiamide can be added. As with the
previously-mixed ingredients, preparatory procedures such as
grinding and/or drying may be performed prior to addition thereof
in the mixing apparatus.
[0079] Mixing to incorporate any of the propellant ingredients can
be performed for any suitable period and under any suitable vacuum
condition. Moreover, mixing can be performed at any suitable
temperature but is typically performed at about room
temperature.
[0080] Following addition of all ingredients, mixing can be
maintained under an applied vacuum, e.g., about 29 inches of
mercury, until the mixture provides desired rheological properties.
For example, vacuum mixing can be performed until about 7 wt % of
solvent remains.
[0081] The mixed composition can also be further processed to
remove any entrapped gas. The mixed composition can also be
rendered in any desired shaped. For example, the composition can be
extruded to remove trapped gas and then form pellets. The extruded
pieces can have any suitable or desired size. For example, the
pieces can be extruded through a die having an aperture diameter in
range of about 0.06 to about 0.08 inch. Further, the extruded
pieces can have any desired aspect ratio. For example, the length
to diameter can be about 2:1 to about 3:1.
[0082] Further processing can optionally be performed on the
extruded pieces prior to disposing into a case, which is further
described below. For example, the extruded pieces can be shaped or
formed to be spherical in a spheronizer. In some cases, further
drying can be performed in, for example, a fluidized bed to remove
any residual solvent. Any suitable fluidizing gas can be used at
any suitable temperature that facilitates solvent evaporation. For
example, an inert gas at a temperature of about 120.degree. F. can
facilitate solvent evaporation.
[0083] Further modifications directed to the propellant formulation
of the invention include the addition of agents that provide or
modify the other physical or performance properties. For example,
one or more glazing or anti-caking agents may be utilized to reduce
the likelihood of clumping or caking of any of the ingredients.
Examples of such agents include, but are not limited to, stearic
acid coatings. Further, agents that preserve one or more of the
propellant formulation ingredients may be utilized. For example, a
sterically-hindered amine anti-oxidant may be utilized to
facilitate stable long term storage of the vinyl resin, the
functionalized modifier, a mixture thereof, and/or the resultant
composite propellant. Other agents such as one or more
sterically-hindered phenol compounds, which may also be utilized
with amide-based sequestering or chelating agents, may facilitate
processing by at least partially inhibiting any reaction between,
for example, the functionalized fuel and the functionalized
modifier components. Such agents would thus allow for longer
work-in-process periods. Further, anti-static agents may be
utilized to reduce the electrostatic sensitivity and accommodate
air conveyor-based transport systems and techniques.
[0084] The present invention also relates to a method of
manufacturing a rocket engine in which propellant compositions of
the invention are loaded into a rocket engine chamber. Such a
method can be used to make one engine at a time using a
hand-operated pressing machine or used with automated equipment
capable of making thousands of engines per day.
[0085] Processed in the dry powder or granular form, the pressable
propellant composition of the present invention is typically not
subject to the viscosity, air entrapment, or pot-life problems
inherent to castable propellants. As such, pressable propellant
rocket engine production is easily automated by means of multiple
feed, hydraulic pressing. In comparison to castable motor
production, this method produces little waste material, can be
quality inspected on a near real time basis and results in a
minimum of product held as "work in progress." Further, the
composite propellant compositions in accordance with some
embodiments of the invention advantageously undergo a B-stage cure,
which can be at ambient conditions, which in turn can
advantageously increase process equipment availability. After
pressing, a first batch of propellant grains can be removed from
the pressing tools and processing a next batch can be initiated.
Because the propellant compositions are pressable and generally
retain a shape or configuration after pressing, the first batch
need not be fully cured before removal from the pressing equipment.
The availability and processing capacity is thus increased,
especially with respect to castable compositions.
[0086] Because the propellant composition of the present invention
contains little or no moisture sensitive raw materials and is post
processed in dry powder or granular form; it is typically not
subject to the same processing constraints as castable
compositions. Processing is straightforward and very flexible with
respect to suitable mixing equipment.
[0087] In another embodiment of the invention, rocket engines are
manufactured by pressing clay or other suitable heat resistant
material into a casing to form the rocket engine nozzle.
Alternatively, the nozzle may be pre-formed and then inserted into
the casing. The casing is preferably a multi-layered paper or
cardboard in any suitable configuration, such as corrugated or
laminated or combinations thereof; may be made from plastic,
fiberglass, a filament-wound glass/epoxy composite, paper phenolic,
plastic phenolic, and aluminum or its alloys, as is known in the
art.
[0088] The case material is selected based on several factors
including, but not limited to, cost, availability, and mechanical
properties. As noted, typical cases can comprise paper,
cellulose-based materials, or non-metallic materials. Thus, the
specific gravity of the material comprising the case can be less
than about 2. The case can further comprise a composite or assembly
of various types of materials to provide desired mechanical
properties. In accordance with some embodiments of the invention,
the case is constructed and arranged based on a chamber operating
pressure of less than about 300 psia, and in some cases, based on a
chamber pressure of about 100 psia. Thus, in some cases, the case
have a burst strength of less than about 2,000 KPa, and/or a
tensile strength of less than or equal to about 5,000 KPa. However,
where it is advantageous to do so, the burst strength can be about
1,000 KPa, or less. For example, where a reduced cost can be
realized without a sacrifice or an increase in the likelihood of
failure, the case can be constructed and arranged utilizing a
material that has a burst strength of about 1,000 KPa, or less.
[0089] The propellant composition powder (for producing thrust) is
fed into the casing in incremental amounts sufficient to achieve a
uniform pressed density, and compressed at high pressure (about
10,000 psia or higher) to form a single propellant grain with
uniform density. Alternatively, the grain may be pre-formed outside
of the casing, and then inserted into the casing. The required
amount of propellant composition needed to produce the desired
total impulse is calculated by dividing the desired total impulse
by the specific impulse of the propellant compound, as is known in
the art. For example, a type "C" engine has a desired total impulse
of 10.0N-S therefore, 7.5 grams of the propellant composition is
used.
[0090] Delay powder is then fed into the engine casing and pressed
to achieve the desired time delay prior to igniting the ejection
charge. Alternatively, it can be formed into a unit and inserted in
the casing after forming. The amount of delay powder used may be
calculated based on, inter alia, the burn rate of the delay powder,
the size, e.g., diameter of the engine and the desired time
requirement.
[0091] Ejection powder, typically black powder, is then put into
the engine casing. The purpose of the ejection powder is to provide
gas to deploy the recovery mechanism.
[0092] Clay or similar heat resistant material is then inserted
into the casing and pressed at a low pressure sufficient to retain
the ejection powder in the casing yet allow the release of the
ejection gases in order to activate the recovery mechanism.
[0093] This invention is further illustrated by the following
examples, which should not be construed as limiting.
EXAMPLES
[0094] The function and advantages of these and other embodiments
of the invention can be further understood from the examples below,
which illustrate the benefits and/or advantages of the one or more
systems and techniques of the invention but do not exemplify the
full scope of the invention. In the examples, the performance of
the various propellant formulations was characterized utilizing,
inter alia, a widely available program, PROPEP. Because the
associated JANAF material database thereof did not include VROH
resin and HELOXY.TM. 505, thermochemically similar species, GEON
121 and dioctyl adipate, were respectively substitituted
therefor.
Example 1
Preparing a Propellant Composition of the Invention (VULCANITE.TM.
EB-75 Composition A)
[0095] The mixer is charged with 100% of the VROH binder (Dow
Chemical) in solvent solution form. Under mixing action in a
planetary mixer, the prescribed amount of mono-modal particle size
potassium perchlorate below is added to the VROH solution until a
uniform, paste like consistency is obtained. Other types of
suitable mixers are known in the art, such as whip mixers, twin
screw mixers and Mueller mixers. The dicyandiamide and iron oxide
are then added to the mixer in respective order and processed until
a uniform, paste-like consistency is achieved. In one-kilogram
batches such as described here, this is a matter of minutes. The
entire mixing process is achieved easily within a very short period
of time, usually in 15 minutes or less. [0096] 0.92% VROH
hydroxyl-modified vinyl copolymer (The Dow Chemical Company,
Midland Mich.) [0097] 64.22% potassium perchlorate (Service
Chemical Inc., Hatfield, Pa.) [0098] 27.52% dicyandiamide (Air
Products and Chemicals, Inc., Allentown, Pa.) [0099] 7.34% iron
oxide nanoparticles (SICOTRANS.RTM. Red L 2715 D)
[0100] Once a uniform mixture is obtained, again, a matter of
minutes, the material is removed from the mixer and excess solvent
removed by evaporation If evaporation is accomplished manually,
this will take several hours. If a dryer, such as a hot air
recirculating tunnel is used, this will take minutes. Upon reaching
a predetermined solvent content whereby the mixture is a pliable
mass capable of being screened without being so fluid as to stick
together, or too dry to be screened, the material is screened to a
desired particle size. For this example, the particle size was an
8-mesh market grade; the particle size can be varied according to
the application. The particles are then allowed to dry completely
over a period of hours, if accomplished manually, prior to
packaging.
Example 2
Preparing Another Propellant Composition of the Invention
(VULCANITE.TM. EB-75 Composition B)
[0101] Following the steps described above, a propellant
composition B was manufactured using 68.04% potassium perchlorate,
19.22% dicyandiamide (.gtoreq.7.mu.), 5.66% iron oxide
nanoparticles, and 7.08% DOS modified VROH. The DOS modified VROH
is prepared by mixing 59.52% methyl ethyl ketone, 23.81% VROH, and
16.67% dioctyl sebacate under agitation until dissolved.
Example 3
Engine Manufacture Using a VULCANITE.TM. EB-75 Propellant
Composition of the Invention
[0102] Clay (about 2.5 grams) is fed and pressed into the
multi-layered paper casing for a rocket engine to form the rocket
engine nozzle. The propellant composition powder is then fed and
pressed into the engine casing in increments sufficient to achieve
a uniform pressed density, at a high pressure of about 10,000 psia,
to form a single propellant grain of uniform density. The amount of
powder used depends upon the type of engine being manufactured and
is determined by calculating the amount needed to produce the
desired total impulse. In this example, a type "C" engine was made,
the total impulse for which is no more than 10.00 N-S. Therefore,
about 7.5 grams of propellant composition was used.
[0103] Delay powder is then fed into the engine casing and pressed
to achieve the desired time delay prior to igniting the ejection
charge. In this example, a time delay of less than 8 seconds was
desired, and 1.5 grams delay powder was added, based on internal
casing diameter and burn rate calculations. Delay powder used in
this example was PYRODEX.RTM. HF-20.TM. from Hodgdon Powder
Company, Inc., Shawnee Mission, Kans., but may be selected from
other slow burning compositions (about 0.05 inches per second) as
is known in the art.
[0104] Ejection powder (black powder) is then put into the engine
casing to create gas to deploy the recovery mechanism. The amount
of ejection powder used is determined by engine casing size and
estimated size of rocket to be flown, and ranges from about 0.5
grams for a "C" type engine to about 1.2 grams for larger
engines.
[0105] Clay, about 0.5 grams for a "C" engine to about 1.2 grams
for larger engines is then inserted into the casing and pressed at
a low pressure to retain the ejection powder in the casing but
still allow the release of the ejection gases in order to activate
the recovery mechanism.
[0106] During and after manufacture, sample engines are
periodically tested to ensure that they function as expected and
otherwise continue to meet engine performance specifications. After
"aging" for at least ten days, engines may be retested to ensure
that they continue to meet specifications.
Example 4
Comparison of Specific Impulse (I.sub.sp) of Black Powder and
Compositions of the Invention
[0107] The combustion characteristics of black powder and a
propellant composition of Example 6 were compared based on
theoretical values. TABLE-US-00002 TABLE 2 VULCANITE .TM. EB-75
Black Powder Composition C Combustion temperature 2221 4313
(.degree. F.) Exhaust gas pressure at nozzle exit 12.1 14.7 (psia)
Combustion chamber pressure 100 50 (psia) Molecular weight of
exhaust 48.341 39.159 mixture Total Exhaust solids 15.96 2.61 (mole
%) Cp/Cv 1.1318 1.2059 Delivered I.sub.sp 75-80 148.6 (lb.
seconds/lb.)
[0108] As can be seen from the table, the propellant composition C
of the invention (Table 3, infra) showed a .DELTA. I.sub.sp=80.13%
increase in specific impulse (I.sub.sp) over black powder. This was
accompanied by a .DELTA. solids of -83.65% reduction in total
exhaust solids, and a .DELTA.T 94.19% increase in combustion
chamber temperature.
Example 5
Combustion and Safety Characteristics
[0109] The granular propellant composition VULCANITE.TM. EB-75
Composition B of Example 2, consisting of 68.04% potassium
perchlorate, 19.22% dicyandiamide (.gtoreq.7.mu.), 5.66% iron oxide
nanoparticles, and 7.08% DOS modified VROH (prepared from 59.52%
methyl ethyl ketone, 23.81% VROH, and 16.67% dioctyl sebacate) was
found to have the following characteristics: [0110] Auto-ignition
temperature: greater than about 500.degree. F. [0111] r=0.1033
P.sub.c.sup.0.3613. [0112] Failed to explode when subjected to DOT
Impact Test. [0113] Only 0.067% weight loss per 200 g at 75.degree.
C. after 48 hours.
[0114] The flame temperature of the composition was calculated to
be 4250.degree. F., the molecular weight of the exhaust mixture was
38.5, and C.sub.p/C.sub.v was approximately 1.2.
Example 6
Calculated Thermodynamic Data for Various Compositions of the
Invention
[0115] Using a Propellant Evaluation Program (PROPEP) widely used
and available via the World Wide Web from, inter alia, The Gas
Dynamics Lab, thermodynamic equilibrium data for various propellant
compositions of the invention was calculated. The results appear in
Table 3, below. Table 4 lists the composition of the propellant
formulations evaluated in Table 3. TABLE-US-00003 TABLE 3 VULCANITE
.TM. EB-75 Composition Composition C Composition D E Combustion
Temperature 4313 4091 4365 (.degree. F.) Exhaust gas pressure at
14.7 14.7 14.7 nozzle exit (psia) Combustion chamber 50 50 50
pressure (psia) Avg. molecular weight of 39.159 28.90 33.466
exhaust products C.sub.p/C.sub.v 1.2059 1.2383 1.2275 I.sub.sp
148.6 166.5 160.5 (lb. seconds/lb.)
[0116] TABLE-US-00004 TABLE 4 VULCANITE .TM. EB-75 Composition C
Composition D Composition E Potassium 68.04% 56.04% 60.04%
perchlorate Nitrogen- 19.22% 33.58% 29.58% containing dicyandiamide
triaminoguanidine triaminoguanidine species dicyanamide iron oxide
5.66% 5.66% 5.66% GEON-121 4.16% 2.78% 2.78% dioctyl adipate 2.92%
1.94% 1.94%
Example 7
Comparison of Specific Impulse and Density Impulse
[0117] The graphs in FIGS. 1 and 2 plot the specific impulse and
density impulse of Propellant Compositions C to E, above, relative
to black powder and a typical castable propellant. As can be seen,
there is a significant increase in mass and volumetric efficiency
of the compositions of the present invention over black powder.
Example 8
Propellant Composition
[0118] A propellant composition, indicated as "Baseline
Formulation," in accordance with one or more embodiments of the
invention was prepared from a base stock mixture or pre-blend
comprising the ingredients listed in Table 5 by mixing the
ingredients listed. The pre-blend mixture was then charged in a
vertical mixer. To the pre-blend mixture, a burn rate catalyst, an
oxidizing agent, and a fuel source were added in the amounts listed
in Table 6. Preparation of the propellant composition was performed
in substantial accordance with the flow chart presented as FIG.
3.
[0119] The ballistic performance, at various chamber pressures, of
the composite propellant composition is presented in FIG. 4. The
ballistic performance was simulated by utilizing PROPEP code (June
1998 version), substituting GEON-121 for VROH, and dioctyl adipate
for HELOXY.TM. 505 modifier. TABLE-US-00005 TABLE 5 Pre-blend
Mixture Concentration Ingredient (wt %) Resin 23.81 (VROH) Modifier
16.67 (HELOXY .TM. 505 Modifier) Solvent 59.52 (Methyl Ethyl
Ketone)
[0120] TABLE-US-00006 TABLE 6 Baseline Formulation Concentration
Ingredient (wt %) Potassium perchlorate 68.04 (particle size of
about 10 .mu.m) Dicyandiamide 19.22 Iron Oxide 5.66 (SICOTRANS
.RTM. Red L 2715 D) Base Stock Mixture 17.49
[0121] TABLE-US-00007 TABLE 7 Baseline Formulation Chamber Results
Exhaust Results Temperature 2696 2097 (K) Pressure 100 12.1 (psia)
Specific Heat (molar) Gas 11.31 11.23 Total 11.31 11.32 Molecular
Weight 38.7 39.4 (grams/mole)
[0122] TABLE-US-00008 TABLE 8 Black Powder Chamber Results Exhaust
Results Temperature 2132 1826 (K) Pressure 100 12.1 (psia) Specific
Heat (molar) Gas 10.94 10.69 Total 17.58 17.71 Molecular Weight
59.2 59.8 (grams/mole)
[0123] Table 7 lists the expected performance of the propellant
formulation at a chamber pressure of about 100 psia and Table 8
lists the corresponding expected performance of black powder as
derived utilizing the PROPEP simulation software. The expected
ballistic performance of black powder at various chamber pressures
is presented in FIG. 5. FIGS. 4 and 5, show the respective Density
Impulse (I.sub.d), Specific Impulse (I.sub.sp in lb/sec),
Combustion Temperature (T in .degree. F.), and Gas Molecular Weight
(MW) relative to chamber pressure (psia).
[0124] The burn rate of the propellant composition of an embodiment
of the invention was characterized to be
r=0.108P.sub.c.sup.0.3613.
[0125] At a chamber operating pressure of about 100 psia, the
predicted molecular weight of the combustion mixture of the
propellant formulation is about 38.7 grams per mole, which compares
favorably relative to the predicted molecular weight of the
combustion product of black powder of about 59.2 grams per
mole.
[0126] Thus, compared to black powder, the propellant formulation
of the present invention can provide a reduction in exhaust gas
molecular weight of at least 20.3%. The resultant specific impulse
was measured to be about 138.7 lbs/lbs-m/sec, which is about 47%
greater than the specific impulse of black powder.
Example 9
Variation of Oxidizer to Fuel Ratio
[0127] Several modifications of the propellant formulation as
substantially disclosed in Example 8 were evaluated. In particular,
the effect on the specific impulse of varying the ratio of
potassium perchlorate to dicyandiamide was evaluated utilizing the
widely available propellant evaluation program, PROPEP code.
[0128] The predicted specific impulse at a chamber pressure of
about 100 psia is presented in FIG. 6, which shows the propellant
formulation of the invention can provide a specific impulse of at
least 145 lb/sec for potassium perchlorate/dicyandiamide weight
ratios from about 64:23.2 to about 71:15 and that a peak specific
impulse of about 148 lb/sec can be achieved at a potassium
perchlorate/dicyandiamide weight ratio of about 68/19.2.
[0129] Notably, the peak ratio is not stoichiometrically balanced;
but can be considered to be underoxidized. This condition, it is
believed, leads to reducing, if not minimizing, the average
molecular weight of the exhaust gases. That is, the reducing
atmosphere in the chamber can provide species that are not
converted to the oxidized state. These conditions facilitate
utilization of paper casing when exposed to the combustion gases at
a temperature of about 4438.degree. F. Specifically, the exhaust of
the propellant formulation of the invention is expected to comprise
about 0.000386 moles of free oxygen at a chamber pressure of about
100 psia, which is considerably less than the corresponding oxygen
concentration of about 0.349 moles of black powder exhaust. Indeed,
if free oxygen were present, oxidizable casings, such as paper,
would degrade, e.g. burn, to failure. In the absence of oxygen,
however, paper casings would pyrolyze to form a layer of insulating
char that protects unexposed casing material.
Example 10
Effect of Binder Content
[0130] Further modifications of the propellant formulation as
substantially disclosed in Example 8 were evaluated. In particular,
the amount of vinyl resin (VROH) and modifier (HELOXY.TM. 505
modifier) was evaluated.
[0131] Increasing the amount of binder would improve pressability
of the propellant composition into propellant grains as well as
reduce the sensitivity thereof to friction, impact, electrostatic
discharge, blasting cap; a reduction in specific impulse would
likely result, as illustrated in FIG. 7. In particular, FIG. 7
illustrates predicted specific impulse (in lb/sec) relative to the
chamber pressure with about 5 wt % binder content (indicated as
reference 5), with about 7 wt % binder content (indicated as
reference 7), and with about 10 wt % binder content (indicated as
reference 10).
Example 11
Effect of Catalyst Content
[0132] Further modifications directed to the catalyst content of
the propellant formulation as substantially disclosed in Example 8
were evaluated.
[0133] Although, it is believed, that specific impulse and density
is typically determined by the combination of the oxidizer (e.g.,
potassium perchlorate), fuel (e.g., dicyandiamide), and the binder
(e.g., vinyl resin/epoxidized castor oil), the burn rate catalyst
may dictate other characteristics and stability of the sustained
combustion at low, less than about 500 psia, or even atmospheric
pressures. In particular, the selected burn rate catalyst of the
invention can reduce or modify the high burn rate exponent
(typically about 0.7) associated with potassium perchlorate while
improving the burn rate coefficient.
[0134] For example, the effect of the amount of SICOTRANS.RTM. Red
L 2715 D is presented in FIG. 8. In particular, the burn rate at
about 2 wt % (indicated as reference 2), at about 4 wt % (indicated
as reference 4), at about 6 wt % (indicated as reference 6), and at
about 8 wt % (indicated as reference 8) of SICOTRANS.RTM. Red L
2715 D iron oxide catalyst relative to chamber pressure is
shown.
Example 12
Effect of Oxidizer and/or Fuel Component
[0135] Further modifications directed to the catalyst content of
the propellant formulation as substantially disclosed in Example 8
were evaluated.
[0136] In particular, the propellant formulation as substantially
disclosed in Example 8 was modified to replace a portion of the
potassium perchlorate with guanidine nitrate resulting in about
50.04 wt % potassium perchlorate and about 22 wt % guanidine
nitrate. The specific impulse at a chamber pressure of about 100
psia is predicted to be about 145 lb/sec. The ballistic performance
relative to chamber pressure is presented in FIG. 9.
[0137] The propellant formulation as substantially disclosed in
Example 8 was also modified to replace dicyandiamide with
triaminoguanidine (about 31.58 wt %). The specific impulse at a
chamber pressure of about 100 psia is predicted to be about 166
lb/sec, which may be advantageous in highly energetic applications
such as, but not limited to, inflators, guillotine cutters, power
cartridges, bursting charges, expelling charges, pin pullers or
pushers, bellows motors, signal rockets, and explosive actuated
switches or valves. The ballistic performance relative to chamber
pressure is presented in FIG. 10.
[0138] The propellant formulation as substantially disclosed in
Example 8 was also modified to replace dicyandiamide with a
comparable amount of hexamethylene tetramine (about 14.22 wt %),
which may be advantageous in highly energetic applications. The
balance of the formulation comprised of about 73.04 wt % potassium
perchlorate, about 5.66 wt % iron oxide catalyst as SICOTRANS.RTM.
Red L 2715 D, about 4.16 wt % VROH resin, and about 2.92 wt %
functionalized modifier as HELOXY.TM. 505. The specific impulse at
a chamber pressure of about 100 psia is predicted to be about 153
lb/sec. The ballistic performance relative to chamber pressure is
presented in FIG. 11.
INCORPORATION BY REFERENCE
[0139] The entire contents of all patents, published patent
applications, and other references cited herein are hereby
expressly incorporated herein in their entireties by reference.
Pending U.S. patent application Ser. No. 10/295,308, entitled
COMPOSITE PROPELLANT COMPOSITIONS, filed Nov. 14, 2002, is
incorporated herein by reference in its entirety.
[0140] Having now described some illustrative embodiments of the
invention, it should be apparent to those skilled in the art that
the foregoing is merely illustrative and not limiting, having been
presented by way of example only. Numerous modifications and other
embodiments are within the scope of one of ordinary skill in the
art and are contemplated as falling within the scope of the
invention. In particular, although many of the examples presented
herein involve specific combinations of method acts or system
elements, it should be understood that those acts and those
elements may be combined in other ways to accomplish the same
objectives.
[0141] Further, acts, elements, and features discussed only in
connection with one embodiment are not intended to be excluded from
a similar role in other embodiments.
[0142] It is to be appreciated that various alterations,
modifications, and improvements can readily occur to those skilled
in the art and that such alterations, modifications, and
improvements are intended to be part of the disclosure and within
the spirit and scope of the invention.
[0143] Moreover, it should also be appreciated that the invention
is directed to each feature, system, subsystem, or technique
described herein and any combination of two or more features,
systems, subsystems, or techniques described herein and any
combination of two or more features, systems, subsystems, and/or
methods, if such features, systems, subsystems, and techniques are
not mutually inconsistent, is considered to be within the scope of
the invention as embodied in the claims.
[0144] Use of ordinal terms such as "first," "second," "third," and
the like in the claims to modify a claim element does not by itself
connote any priority, precedence, or order of one claim element
over another or the temporal order in which acts of a method are
performed, but are used merely as labels to distinguish one claim
element having a certain name from another element having a same
name (but for use of the ordinal term) to distinguish the claim
elements.
[0145] Those skilled in the art should appreciate that the
parameters and configurations described herein are exemplary and
that actual parameters and/or configurations will depend on the
specific application in which the systems and techniques of the
invention are used. Those skilled in the art should also recognize
or be able to ascertain, using no more than routine
experimentation, equivalents to the specific embodiments of the
invention. It is therefore to be understood that the embodiments
described herein are presented by way of example only and that,
within the scope of the appended claims and equivalents thereto;
the invention may be practiced otherwise than as specifically
described.
* * * * *