U.S. patent application number 10/295308 was filed with the patent office on 2004-05-20 for composite propellant compositions.
This patent application is currently assigned to Estes-Cox Corporation. Invention is credited to Brown, Edwin Duane, Dixon, Scott James, Tunick, Barry Richard.
Application Number | 20040094250 10/295308 |
Document ID | / |
Family ID | 32297163 |
Filed Date | 2004-05-20 |
United States Patent
Application |
20040094250 |
Kind Code |
A1 |
Dixon, Scott James ; et
al. |
May 20, 2004 |
Composite propellant compositions
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: |
LAHIVE & COCKFIELD, LLP.
28 STATE STREET
BOSTON
MA
02109
US
|
Assignee: |
Estes-Cox Corporation
1295 H. Street
Penrose
CO
81240
|
Family ID: |
32297163 |
Appl. No.: |
10/295308 |
Filed: |
November 14, 2002 |
Current U.S.
Class: |
149/19.91 |
Current CPC
Class: |
C06B 45/10 20130101;
C06B 23/007 20130101 |
Class at
Publication: |
149/019.91 |
International
Class: |
C06B 045/10 |
Claims
1. A propellant composition comprising a solid inorganic oxidizing
agent; a nitrogen-containing fuel; and a burn rate catalyst,
wherein said burn rate catalyst is nanoparticulate.
2. The composition of claim 1 wherein said burn rate catalyst is an
oxide selected from the group consisting of copper, chromium,
cobalt, manganese, iron, vanadium, and mixtures thereof.
3. The composition of claim 1, wherein said material is a
propellant material, a pyrotechnic material, an explosive material,
a light generating material, a heat generating material, or a sound
generating material.
4. The composition of claim 1, wherein said nanoparticles are
irregularly shaped.
5. The composition of claim 2, wherein said burn rate catalyst is
about 0.1-15 wt % oxide.
6. The composition of claim 2, wherein said burn rate catalyst is
about 0.5-10 wt % oxide.
7. The composition of claim 1, further comprising a binder.
8. The composition of claim 7, wherein said binder is a combustible
organic polymer selected from the group consisting of
polyvinylchloride, polyvinylacetate, polyvinylalcohol, and
copolymers thereof including VAAR, VROH, VAGH, Geon-121, poly
(2-ethyl-2-oxazoline), and epoxy or acrylate resin.
9. The composition of claim 7, wherein said composition is about
0.5%-15% binder.
10. The composition of claim 7, wherein said composition is about
0.75%-12% binder.
11. The composition of claim 7, further comprising a
plasticizer.
12. The composition of claim 11, wherein said plasticizer is
selected from the group consisting of dioctyl adipate, dioctyl
sebacate, hydrocarbon ester tackifier, and combinations
thereof.
13. The composition of claim 1, wherein said solid inorganic
oxidizing agent is selected from the group consisting of potassium
perchlorate and ammonium perchlorate.
14. The composition of claim 1, wherein said composition is about
60-75 wt % solid oxidizing agent.
15. The composition of claim 1, wherein said composition is about
64-72 wt % solid oxidizing agent.
16. The compositions of claim 1, wherein said nitrogen-containing
fuel comprises a cyano-, amide-, or amine-containing material.
17. The composition of claim 16, where said nitrogen-containing
fuel is selected from the group consisting of acrylonitrile, amino
tetrazole, aminoguanidinium bitetrazole, ammonium dicyanamide,
bistriaminoguanidiniumdecacarborane, bis(trinitroethyl)nitramine,
calcium bitetrazole, dicyandiamide, nitroaminoguanidine,
triaminoguanidine, and triaminoguanidinedicyanamide.
18. The composition of claim 1, wherein said composition is about
10-30 wt % nitrogen-containing fuel.
19. The composition of claim 1, wherein said composition is about
15-23 wt % nitrogen-containing fuel.
20. The composition of claim 7, wherein said binder is selected
from the group consisting of epoxidized trimethylolpropane,
trimethylol ethane trigylcidyl ether, epoxidized soybean oil and
combinations thereof.
21. The composition of claim 1, wherein said composition has a
specific impulse of at least 100 seconds at 100 psia.
22. The composition of claim 1, wherein said composition has a
specific impulse of at least 120 seconds at 100 psia.
23. The composition of claim 1, wherein the average molecular
weight of the exhaust gases of said composition is less than about
45.
24. The composition of claim 1, wherein the average molecular
weight of the exhaust gases of said composition is less than about
40.
25. The composition of claim 1, wherein the combustion temperature
is at least about 3000.degree. F.
26. The composition of claim 1, wherein the combustion temperature
is at least about 4000.degree. F.
27. The composition of claim 1, wherein the burn rate of said
composition is r.sub.h=a(P.sub.c).sup.n, wherein a is the burn rate
coefficient, P.sub.c is the chamber pressure, and n is the burn
rate exponent having a value of less than about 0.5.
28. The composition of claim 1, wherein the burn rate of said
composition is r.sub.h=a(P.sub.c).sup.n, wherein a is the burn rate
coefficient, P.sub.c is the chamber pressure, and n is the burn
rate exponent having a value of less than about 0.35.
29. The composition of claim 1, wherein the burn rate of said
composition is r.sub.h=a(P.sub.c).sup.n, wherein a is the burn rate
coefficient having a value of greater than about 0.1, P.sub.c is
the chamber pressure, and n is the burn rate exponent.
30. The composition of claim 1, consisting of about 62-72% wt.
potassium perchlorate, about 15-30% wt. dicyandiamide, about 2-10%
wt. iron oxide, and about 3-12% wt. VROH base stock, wherein said
VROH base stock comprises a mixture of VROH and dioctyl
sebacate.
31. The composition of claim 30, wherein the relative amounts of
VROH and dioctyl sebacate are about 3:2.
32. The composition of claim 1, consisting of about 68% wt.
potassium perchlorate, about 19% wt. dicyandiamide, about 5% wt.
iron oxide, and about 7.5% wt. VROH base stock, wherein said VROH
base stock comprises a mixture of VROH and dioctyl sebacate.
33. The composition of claim 32, wherein the relative amounts of
VROH and dioctyl sebacate are about 3:2.
34. The composition of claim 1, having a specific impulse of at
least about 120% of the specific impulse of black powder.
35. The composition of claim 1, wherein the molar %. of
non-expandable solid by-products in the exhaust stream is less than
about molar 5%.
36. The composition of claim 1, wherein said composition is adapted
to be pressed or molded in the form of a grain.
37. The composition of claim 36, wherein said composition is
adapted to be pressed directly into a rocket engine casing or
molded to fit an engine by means of a mold.
38. A method of manufacturing a rocket engine comprising a step of
loading a rocket engine chamber with a propellant composition
according to claim 1.
39. A method of manufacturing a rocket engine comprising a step of
loading a rocket engine chamber with a propellant composition
according to claim 32.
40. The method according to claim 38, wherein said method is
automated.
41. A rocket engine packed with a propellant composition comprising
a solid inorganic perchlorate oxidizing agent; a
nitrogen-containing fuel; and a burn rate catalyst, wherein said
burn rate catalyst is particulate.
42. The rocket engine of claim 41, wherein said burn rate catalyst
is nanoparticulate.
43. The rocket engine of claim 41 wherein said burn rate catalyst
is an oxide selected from the group consisting of copper, chromium,
cobalt, manganese, iron, vanadium, and mixtures thereof.
44. The rocket engine of claim 41, wherein said composition further
comprises a binder.
45. The rocket engine of claim 44, wherein said binder is a
combustible organic polymer selected from the group consisting of
polyvinylchloride, polyvinylacetate, polyvinylalcohol and
copolymers thereof including VAAR, VROH, VAGH, Geon-121, poly
(2-ethyl-2-oxazoline), and epoxy or acrylate resin.
46. The rocket engine of claim 41, wherein said composition further
comprises a plasticizer.
47. The compositions of claim 46, wherein said plasticizer is
selected from the group consisting of dioctyl adipate, dioctyl
sebacate, hydrocarbon ester tackifier, and combinations
thereof.
48. The rocket engine of claim 41, wherein said solid inorganic
oxidizing agent is selected from the group consisting of potassium
perchlorate and ammonium perchlorate.
49. The rocket engine of claim 41, wherein said nitrogen-containing
fuel comprises a cyano-, amide-, or amine-containing material.
50. The rocket engine of claim 49, where said nitrogen-containing
fuel is selected from the group consisting of acrylonitrile, amino
tetrazole, aminoguanidinium bitetrazole, ammonium dicyanamide,
bistriaminoguanidiniumdecacarborane, bis(trinitroethyl)nitramine,
calcium bitetrazole, dicyandiamide, nitroaminoguanidine,
triaminoguanidine, and triaminoguanidinedicyanamide.
51. A high burn rate, high combustion temperature propellant
composition comprising a solid inorganic perchlorate oxidizing
agent; a nitrogen-containing fuel consisting essentially of
dicyandiamide; a burn rate catalyst; and a binder.
52. The composition of claim 51 wherein said burn rate catalyst is
selected from the group consisting of oxides of copper, chromium,
cobalt, manganese, iron, vanadium, and mixtures thereof.
53. The composition of claim 52, wherein said oxide is
particulate.
54. The composition of claim 53, wherein said particles are
irregularly shaped.
55. The composition of claim 52, further comprising a
plasticizer.
56. The composition of claim 51, wherein said solid inorganic
perchlorate oxidizing agent is selected from the group consisting
of potassium perchlorate and ammonium perchlorate.
57. The composition of claim 51, wherein said binder is selected
from the group consisting of polyvinylchloride, polyvinylacetate,
polyvinylalcohol and copolymers thereof including VAAR, VROH, VAGH,
Geon-121, poly (2-ethyl-2-oxazoline), and epoxy or acrylate
resin.
58. The composition of claim 51, wherein said composition is about
60-75 wt % solid oxidizing agent.
59. The composition of claim 51, wherein said composition is about
10-30 wt % nitrogen-containing fuel.
60. The composition of claim 52, wherein said composition is about
0.1-15 wt % oxide.
61. The composition of claim 51, wherein said composition is about
0.5%-15% binder.
62. A high burn rate, high combustion temperature propellant
composition comprising about 64-72 wt. % solid inorganic
perchlorate; about 15-23 wt. % nitrogen-containing fuel; about
0.5-10 wt. % oxide of copper, chromium, cobalt, manganese, iron,
vanadium, or a mixture thereof; and about 0.75-12 wt. % binder.
63. The composition of claim 62, wherein said oxide is
particulate.
64. The composition of claim 63, wherein said oxide particles are
irregularly shaped.
65. The composition of claim 62, further comprising a
plasticizer.
66. The composition of claim 65, wherein said plasticizer is
selected from the group consisting of dioctyl adipate, dioctyl
sebacate, hydrocarbon ester tackifier, and combinations
thereof.
67. The composition of claim 62, wherein said solid inorganic
perchlorate is selected from the group consisting of potassium
perchlorate and ammonium perchlorate.
68. The composition of claim 62, wherein said nitrogen-containing
fuel comprises a cyano-, amide-, or amine-containing material.
69. The composition of claim 68, where said nitrogen-containing
fuel is selected from the group consisting of acrylonitrile, amino
tetrazole, aminoguanidinium bitetrazole, ammonium dicyanamide,
bistriaminoguanidiniumdecacarborane, bis(trinitroethyl)nitramine,
calcium bitetrazole, dicyandiamide, nitroaminoguanidine,
triaminoguanidine, and triaminoguanidinedicyanamide.
70. The composition of claim 62, wherein said binder is selected
from the group consisting of polyvinylchloride, polyvinylacetate,
polyvinylalcohol and copolymers thereof including VAAR, VROH, VAGH,
Geon-121, poly (2-ethyl-2-oxazoline), and epoxy or acrylate
resin.
71. 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.
72. The composition of claim 71, wherein said burn rate catalyst is
an oxide selected from the group consisting of copper, chromium,
cobalt, manganese, iron, vanadium, and mixtures thereof.
73. The composition of claim 71, wherein said material is a
propellant material, a pyrotechnic material, an explosive material,
a light generating material, a heat generating material, or a sound
generating material.
74. The composition of claim 71, wherein said nanoparticles are
irregularly shaped.
75. The composition of claim 72, wherein said burn rate catalyst is
about 0.1-15 wt % oxide.
76. The composition of claim 71, further comprising a binder.
77. The composition of claim 76, wherein said binder is a
combustible organic polymer selected from the group consisting of
polyvinylchloride, polyvinylacetate, polyvinylalcohol, and
copolymers thereof including VAAR, VROH, VAGH, Geon-121, poly
(2-ethyl-2-oxazoline), and epoxy or acrylate resin.
78. The composition of claim 76, wherein said composition is about
0.5%-15% binder.
79. The composition of claim 76, further comprising a
plasticizer.
80. The composition of claim 79, wherein said plasticizer is
selected from the group consisting of dioctyl adipate, dioctyl
sebacate, hydrocarbon ester tackifier, and combinations
thereof.
81. The composition of claim 71, wherein said composition is about
60-75 wt % solid oxidizing agent.
82. The compositions of claim 71, wherein said nitrogen-containing
fuel comprises a cyano-, amide-, or amine-containing material.
83. The composition of claim 82, where said nitrogen-containing
fuel is selected from the group consisting of acrylonitrile, amino
tetrazole, aminoguanidinium bitetrazole, ammonium dicyanamide,
bistriaminoguanidiniumdecacarborane, bis(trinitroethyl)nitramine,
calcium bitetrazole, dicyandiamide, nitroaminoguanidine,
triaminoguanidine, and triaminoguanidinedicyanamide.
84. The composition of claim 71, wherein said composition is about
10-30 wt % nitrogen-containing fuel.
85. The composition of claim 76, wherein said binder is selected
from the group consisting of epoxidized trimethylolpropane,
trimethylol ethane trigylcidyl ether, epoxidized soybean oil and
combinations thereof.
86. The composition of claim 71, wherein said composition has a
specific impulse of at least 100 seconds at 100 psia.
87. The composition of claim 71, wherein the average molecular
weight of the combustion gases of said composition is less than
about 45.
88. The composition of claim 71, wherein the combustion temperature
is at least about 3000.degree. F.
89. The composition of claim 71, wherein the burn rate of said
composition is r.sub.h=a(P.sub.c).sup.n, wherein a is the burn rate
coefficient, P.sub.c is the chamber pressure, and n is the burn
rate exponent having a value of less than about 0.5.
90. The composition of claim 71, wherein the burn rate of said
composition is r.sub.h=a(P.sub.c).sup.n, wherein a is the burn rate
coefficient having a value of greater than about 0.1, PC is the
chamber pressure, and n is the burn rate exponent.
91. The composition of claim 71, consisting of about 62-72% wt.
potassium perchlorate, about 15-30% wt. dicyandiamide, about 2-10%
wt. iron oxide, and about 3-12% wt. VROH base stock, wherein said
VROH base stock comprises a mixture of VROH and dioctyl
sebacate.
92. The composition of claim 71, having a specific impulse of at
least about 120% of the specific impulse of black powder.
93. The composition of claim 71, wherein the molar %. of
non-expandable solid by-products in the combustion gas is less than
about molar 5%.
94. The composition of claim 71, wherein said composition is
adapted to be pressed or molded in the form of a grain.
Description
BACKGROUND
[0001] Composite propellant compositions typically contain separate
fuel and oxidizer components that are intimately mixed. Black
powder, sometimes referred to as gunpowder, invented by the ancient
Chinese, 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.
[0002] Black powder model rocket engines (or 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 is tightly packed into a casing,
usually a multi-layered paper tube. An electrical igniter is used
to ignite the model rocket engine. Because the black powder is
tightly packed to a uniform density, it burns evenly and produces
thrust as the hot expanding gases escape the rocket engine via the
nozzle. Because black powder is granular and pressable, rocket
engine production can be easily automated by means of multiple
feed, hydraulic pressing machinery.
[0003] Although black powder is relatively inexpensive and readily
available, it has a relatively low specific impulse. Other solid
composite propellants have been invented 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.
[0004] 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.
[0005] To achieve useful levels of energy and density, castable
propellants are formulated in such a manner as to obtain high
solids loadings (82%-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.
[0006] 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 4 and 10 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.
[0007] 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
(2 to 4 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.
[0008] Model rocket engines using castable composites are 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 model rocket engines is therefore labor
intensive, equipment limited, and difficult to automate.
SUMMARY OF THE INVENTION
[0009] 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. 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.
[0010] In one aspect, the invention relates to propellant
compositions comprising a solid inorganic oxidizing agent; a
nitrogen-containing fuel; and a burn rate catalyst. As used herein,
it is to be understood that propellant compositions of the present
invention can be used as a propellant material, a pyrotechnic
material, an explosive material, a light generating material, a
heat generating material, or a sound generating material.
[0011] The invention also relates to a rocket engine (sometimes
referred to as a rocket "motor") packed with a composition
comprising a solid inorganic perchlorate oxidizing agent, a
nitrogen-containing fuel, and a burn rate catalyst.
[0012] In 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.
[0013] In a further embodiment, 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.
[0014] 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 DRAWING
[0015] FIG. 1 depicts the specific impulse of 3 propellant
compositions of the present invention described in Table III,
below, against black powder and a typical castable composite.
[0016] FIG. 2 depicts the density impulse of 3 propellant
compositions of the present invention described in Table III,
below, against black powder and a typical castable composite.
DETAILED DESCRIPTION OF THE INVENTION
[0017] 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 used
today 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.
[0018] The individual components of the present invention were
selected based on 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 have
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. Propellant compositions of the present invention
allow higher levels of automation; therefore, they allow greater
production rates accompanied by a significant decrease in
work-in-process materials.
[0019] 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
motor production is easily automated by means of single or multiple
feed, hydraulic pressing machinery.
[0020] Particularly regarding its use in manufacturing of model
rocket engines, the propellant compositions can be used as granules
or pellets and pressed into the engine casing, or they can be
pressed into molds to make "grains," the mass of solid propellant
used in model rocket engines or other applications. Grains can then
be inserted into engine casings or sold separately, allowing the
consumer to load the casing on an "as needed" basis. 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 ignition
transfer pellets used in piccolo tubes, pyrotechnics, artillery
shell igniters, hand signals, maritime smokes and signals,
expelling charges and various other military applications.
[0021] The propellant compositions of the invention may be used as
a high burn rate (at least {fraction (5/10)} inch per second at 150
pounds per square inch absolute (psia)) propellant for 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.
[0022] The propellant compositions of the invention also have a
high combustion temperature (as high as 4000.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 II,
where the Isp of the propellant composition of the present
invention is about 1.73 times that of black powder and produces a
combustion temperature nearly 1.5 times that of black powder.
Generally the combustion temperature is at least 3500.degree. F.,
and in some cases the combustion temperature is greater than
4000.degree. F. Less propellant composition 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.
[0023] In addition, the propellant compositions of the invention
have improved ballistic properties. Compositions of the present
invention have a specific impulse of at least 100 to 120 seconds at
100 psia (pounds per square inch absolute), and thus are more
energetic than black powder, which typically has a specific impulse
of about 80-101 at 100 psia. In selected embodiments, the
propellant compositions of the invention have a specific impulse of
at least 120% that of black powder. See Table II, below.
[0024] The specific impulse, Isp, may be calculated according to
the following formula: 1 Isp = 1 g ( 2 - 1 ) ( R T c M ) [ 1 - ( P
e P c ) ( - 1 ) ]
[0025] where
[0026] .gamma. is the ratio of the specific heats of the combustion
gases (Cp/Cv),
[0027] R is the universal gas constant,
[0028] T.sub.c is the absolute combustion temperature,
[0029] g is the gravitational acceleration constant,
[0030] M is the average molecular weight of the exhaust gases,
[0031] P.sub.c is the combustion chamber pressure, and
[0032] P.sub.e is the exhaust gas pressure at the nozzle exit.
[0033] In order 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
most rocketry applications including the present invention as it
pertains to model rocketry, the ratio of chamber pressure to exit
pressure is determined by the design of the rocket engine itself in
conjunction with the ballistic properties of the propellant.
[0034] 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 such
as black powder.
[0035] 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. In consequence, the compositions of the present invention do
not exhibit the same batch to batch variability.
[0036] The compositions of the invention produce exhaust gases
(combustion gases) of low average molecular weight, less than 45,
and preferably about 39-40. 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 model rocketry
applications, there is less solid residue that can alter the
performance of a model rocket engine. Reducing the solid residue
also is desirable because it produces less build-up on launch
equipment.
[0037] The very low percentage of non-expandable solid by-products
in the exhaust stream (preferably less than about 5 molar %)
increases efficiency. By comparison, black powder engines typically
have 15.96 molar % solid by-products in the exhaust stream.
Castable compositions typically have about 1 molar % 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.
[0038] The Vulcanite.TM. EB-75 propellant compositions of the
invention also have a high burn rate coefficient as compared to
typical castable compositions. The burn rate may be expressed as
r.sub.h=a(P.sub.c).sup.n, 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 I, below, in comparison with typical castable
compositions, these propellant compositions 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
with relatively high burn rate coefficient is both unique and
beneficial. It is beneficial because, inter alia, it allows for
higher mass flow rates at low chamber pressures of about 50-200
psia, pressures at which known castable compositions typically do
not function reliably.
1TABLE I Comparative Burn Rates Burn Rate (above Burn Rate Burn
Rate about 100 psia) Exponent Coefficient Black Powder r = 0.5974
P.sub.c.sup.0.0789 0.05-0.11 0.50-0.70 Typical Castable r = 0.0482
P.sub.c.sup.0.3607 0.32-0.70 0.02-0.06 Composition Vulcanite .TM.
EB-75 r = 0.1033 P.sub.c.sup.0.3613 0.36-0.50 >about 0.10
Propellant Composition
[0039] The propellant compositions of the invention also have a low
burn rate exponent, which means that there is less 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 model rocket engines where
the inner diameter of the casing may vary from casing to casing and
even within individual casings. The burn rate exponent of
compositions of the invention may be less than about 0.5, or
preferably less than about 0.35.
[0040] In particular, the invention relates to propellant
compositions comprising a solid inorganic perchlorate oxidizing
agent; a nitrogen-containing fuel; and a burn rate catalyst, which
is preferably an oxide of copper, chromium, cobalt, manganese,
iron, vanadium, or a mixture thereof. The burn rate catalyst is
preferably a high surface area particulate. Similarly, the
invention pertains to rocket engines packed with propellant
compositions comprising a solid inorganic perchlorate oxidizing
agent; a nitrogen-containing fuel; and a burn rate catalyst which
is preferably a high surface area particulate metal oxide of
copper, chromium, cobalt, manganese, iron, vanadium, or a mixture
thereof.
[0041] More particularly, 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. The burn rate catalyst is preferably an
oxide of copper, chromium, cobalt, manganese, iron, vanadium, or a
mixture thereof.
[0042] A specific advantageous example of the propellant
compositions of the invention is a high burn rate, high combustion
temperature composition comprising about 64-72 wt. % solid
inorganic perchlorate; about 15-23 wt. % nitrogen-containing fuel;
about 0.5-10 wt. % oxide of copper, chromium, cobalt, manganese,
iron, vanadium, or a mixture thereof; and about 0.75-12 wt. %
combustible binder.
[0043] The propellant compositions of the invention contain a burn
rate catalyst that provides a large surface area to increase
catalytic activity. The average size of the catalyst is generally
small and therefore the particles may be referred to as
"nanoparticles." The surface area of the catalyst is preferably
greater than about 50 m.sup.2/g. In compositions of the invention,
the burn rate catalyst may constitute about 0.1-15 wt % of the
compositions. The preferred catalysts are metal oxides that
accelerate the reduction/oxidation (redox) reactions associated
with combustion and the catalyst is instrumental in lowering the
burn rate exponent. The catalyst is preferably a high surface area
particulate, and is irregularly shaped. A preferred metal oxide
catalyst of the invention is an iron oxide known as Sicotrans.TM.
L-2715-D iron oxide, available from BASF.
[0044] The solid inorganic oxidizing agent may be potassium
perchlorate or ammonium perchlorate. In some embodiments, the
oxidizing agent is about 60-75 wt %, or more particularly, about
64-72 wt % of the composition.
[0045] 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)nitra- mine,
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 %, or more
particularly about 15-23 wt %.
[0046] A binder may also be used in the propellant compositions of
the invention. A binder may be added to prevent the resulting
granules from being easily degraded, e.g., broken down into
potentially dangerous dust during the manufacturing processes. The
binder may also increase the mechanical strength of the
composition, such as by cross-linking, and upon curing or aging
react to form a cross-linked network within the composition,
thereby imparting greater mechanical strength, and aiding in
pressing of the composition. The binder may be modified through the
use of a plasticizer that further makes the composition easier to
process. Preferred binders are combustible and produce low
molecular weight by-products, and are preferably combustible
organic polymers. Unlike inorganic binders such as silica, the
preferred binders do not increase the quantity of non-expandable
solid combustion products. The preferred binders are understood to
assist, rather than interfere with, the propellant functionality.
Examples of binders effective in the practice of the present
invention include polyvinylchloride, polyvinylacetate,
polyvinylalcohol and/or copolymers thereof such as the solution
vinyl resins VAAR, VROH, and VAGH (Union Carbide/DOW Chemical),
Geon-121 (B. F. Goodrich), poly (2-ethyl-2-oxazoline), and epoxy or
acrylate resin, epoxidized trimethylolpropane, trimethylol ethane
triglycidyl ether (such as HELOXY Modifier-44 from Shell Chemical
Company), epoxidized soybean oil and combinations thereof. The
propellant compositions of the invention may be 0.5%-15%, or more
particularly 0.75%-12% 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.
[0047] 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 VROH and a plasticizer mixed in a
solvent and subsequently evaporated. The VROH stock solution
comprises a mixture of 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 58.8:41.2 VROH to dioctyl sebacate.
[0048] The propellant compositions of the invention may be used as
a propellant material, a pyrotechnic material, an explosive
material, a light generating material, a heat generating material,
or a sound generating material.
[0049] 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.
[0050] Processed in the dry powder or granular form, the pressable
propellant composition of the present invention is not subject to
the viscosity, air entrapment or pot-life problems inherent to
castable propellants. As such, pressable propellant rocket motor
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".
[0051] Because the propellant composition of the present invention
contains no moisture sensitive raw materials and is post processed
in dry powder or granular form; it is not subject to the same
processing constraints as castable compositions. Processing is
straightforward and very flexible with respect to suitable mixing
equipment.
[0052] 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 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.
[0053] 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 10.0 N-S therefore, 7.5 grams of the propellant composition is
used.
[0054] 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 using known trajectory analysis equations or programs,
which take into account, inter alia, the power of the engine, the
weight of the rocket, the drag coefficient, and the diameter of the
rocket.
[0055] 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.
[0056] 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.
[0057] This invention is further illustrated by the following
examples, which should not be construed as limiting.
EXAMPLES
Example I
Preparing a Propellant Composition of the Invention (Vulcanite.TM.
EB-75 Composition A)
[0058] 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.
[0059] 0.92% VROH (Dow Chemical)
[0060] 64.22% potassium perchlorate (Service Chemical Corp.)
[0061] 27.52% dicyandiamide (Air Products Corp.)
[0062] 7.34% iron oxide nanoparticles (Sicotrans.TM. L-2715-D,
BASF)
[0063] 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 II
Preparing Another Propellant Composition of the Invention
(Vulcanite.TM. EB-75 Composition B)
[0064] 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 III
Engine Manufacture Using a Vulcanite.TM. EB-75 Propellant
Composition of the Invention
[0065] Clay (about 2.5 grams) is fed and pressed into the
multi-layered paper casing for the model 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, 7.5 grams of propellant composition was used.
[0066] 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 trajectory
analysis calculations. Delay powder used in this example was
Pyrodex HF-20.TM. made by Hodgdon Powder Company, but may be
selected from other slow burning compositions (0.05 inches per
second) as is known in the art.
[0067] 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.
[0068] 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.
[0069] 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 IV
Comparison of Specific Impulse (I.sub.sp) of Black Powder and
Compositions of the Invention
[0070] The combustion characteristics of black powder and a
propellant composition of Example VI were compared based on
theoretical values.
2 TABLE II Vulcanite .TM. EB-75 Black Powder Composition C
Combustion temperature 2221.degree. F. 4313.degree. F. Exhaust gas
pressure at nozzle exit 12.10 psia 14.70 psia Combustion chamber
pressure 100 psia 50 psia Molecular weight of exhaust 48.341 39.159
mixture Total Exhaust solids 15.96 molar % 2.61 molar % Cp/Cv
1.1318 1.2059 Delivered I.sub.sp 75-80 lb. 148.6 lb. seconds/lb.
seconds/lb.
[0071] As can be seen from the table, the propellant composition C
of the invention (Table III, 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 V
Combustion and Safety Characteristics
[0072] The granular propellant composition Vulcanite.TM. EB-75
Composition B of Example II, 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:
[0073] Auto-ignition temperature: >500.degree. F.
[0074] r=0.1033 P.sub.c.sup.0.3613.
[0075] Failed to explode when subjected to DOT Impact Test.
[0076] Only 0.067% weight loss per 200 g at 75.degree. C. after 48
hours.
[0077] 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 VI
Calculated Thermodynamic Data for Various Compositions of the
Invention
[0078] 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 III, below:
3 TABLE III 68.04% potassium perchlorate 56.04% potassium
perchlorate 60.04% potassium perchlorate 19.22% dicyandiamide
33.58% triaminoguanidine 29.58% triaminoguanidinedicyanamid- e
5.66% iron oxide 5.66% iron oxide 5.66% iron oxide 4.16% Geon-121
2.78% Geon-121 2.78% Geon-121 2.92% dioctyl adipate 1.94% dioctyl
adipate 1.94% dioctyl adipate (Vulcanite .TM. EB-75 (Vulcanite .TM.
EB-75 (Vulcanite .TM. EB-75 Composition C) Composition D)
Composition E) Combustion 4313.degree. F. 4091.degree. F.
4365.degree. F. temperature Exhaust gas 14.70 psia 14.7 psia 14.7
psia pressure at nozzle exit Combustion 50 psia 50 psia 50 psia
chamber pressure Avg. 39.159 28.90 33.466 molecular weight of
exhaust products C.sub.p/C.sub.v 1.2059 1.2383 1.2275 I.sub.sp
148.6 lb. seconds/lb. 166.5 lb. seconds/lb. 160.5 lb.
seconds/lb.
Example VI
Comparison of Specific Impulse and Density Impulse
[0079] The graphs in FIGS. 1 and 2 plot the specific impulse and
density impulse of Propellant Compositions C-E, above, against that
of 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.
[0080] INCORPORATION BY REFERENCE
[0081] The entire contents of all patents, published patent
applications and other references cited herein are hereby expressly
incorporated herein in their entireties by reference.
Equivalents
[0082] Those skilled in the art will recognize, or be able to
ascertain using no more than routine experimentation, numerous
equivalents to the specific procedures described herein. Such
equivalents are considered to be within the scope of this invention
and are covered by the following claims.
* * * * *