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.

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™ 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° 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° F., and in some cases the combustion temperature is greater than 4000° 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 ⁢   ⁢ γ γ - 1 ) ⁢ ( R ⁢   ⁢ T c M ) ⁡ [ 1 - ( P e P c ) ( γ - 1 ) γ ]

[0025] where

[0026] &ggr; is the ratio of the specific heats of the combustion gases (Cp/Cv),

[0027] R is the universal gas constant,

[0028] Tc is the absolute combustion temperature,

[0029] g is the gravitational acceleration constant,

[0030] M is the average molecular weight of the exhaust gases,

[0031] Pc is the combustion chamber pressure, and

[0032] Pe 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™ 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 rh=a(Pc)n, wherein a is the burn rate coefficient, Pc 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. 1 TABLE I Comparative Burn Rates Burn Rate (above Burn Rate Burn Rate about 100 psia) Exponent Coefficient Black Powder r = 0.5974 Pc0.0789 0.05-0.11 0.50-0.70 Typical Castable r = 0.0482 Pc0.3607 0.32-0.70 0.02-0.06 Composition Vulcanite ™ EB-75 r = 0.1033 Pc0.3613 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 m2/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™ 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)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 %, 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™ 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™ 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™ EB-75 Composition B)

[0064] Following the steps described above, a propellant composition B was manufactured using 68.04% potassium perchlorate, 19.22% dicyandiamide (≧7&mgr;), 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™ 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™ 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 (Isp) 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 ™ EB-75 Black Powder Composition C Combustion temperature 2221° F. 4313° 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 Isp 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 &Dgr;Isp=80.13% increase in specific impulse (Isp) over black powder. This was accompanied by a &Dgr; solids of −83.65% reduction in total exhaust solids, and a &Dgr;T 94.19% increase in combustion chamber temperature.

Example V

Combustion and Safety Characteristics

[0072] The granular propellant composition Vulcanite™ EB-75 Composition B of Example II, consisting of 68.04% potassium perchlorate, 19.22% dicyandiamide (≧7&mgr;), 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° F.

[0074] r=0.1033 Pc0.3613.

[0075] Failed to explode when subjected to DOT Impact Test.

[0076] Only 0.067% weight loss per 200 g at 75° C. after 48 hours.

[0077] The flame temperature of the composition was calculated to be 4250° F., the molecular weight of the exhaust mixture was 38.5, and Cp/Cv 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% triaminoguanidinedicyanamide 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 ™ EB-75 (Vulcanite ™ EB-75 (Vulcanite ™ EB-75 Composition C) Composition D) Composition E) Combustion 4313° F. 4091° F. 4365° 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 Cp/Cv 1.2059 1.2383 1.2275 Isp 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.

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° F.

26. The composition of claim 1, wherein the combustion temperature is at least about 4000° F.

27. The composition of claim 1, wherein the burn rate of said composition is rh=a(Pc)n, wherein a is the burn rate coefficient, Pc 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 rh=a(Pc)n, wherein a is the burn rate coefficient, Pc 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 rh=a(Pc)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.

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° F.

89. The composition of claim 71, wherein the burn rate of said composition is rh=a(Pc)n, wherein a is the burn rate coefficient, Pc 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 rh=a(Pc)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.