Abstract: A spaceplane suitable for aeronautical flight comprising a body and a wing defining a lower airfoil surface in addition to attitude control means that comprise one or a plurality of shutters disposed under the lower airfoil surface of same and maneuverable between a stowed position and an inclined extended position for aerodynamic braking during the transition from a phase of space flight to a phase of aeronautical flight of the aircraft.

Abstract: A bomb for deployment from an air vehicle includes a rocket motor for propelling the bomb. The rocket motor includes propellant which at least partially defines a void downstream of an initial burning surface. As the propellant is burned the void will be exposed, increasing the surface area of the burning surface of the propellant to increase the thrust of the rocket motor.

Abstract: A hybrid rocket motor is manufactured by photopolymerizing the solid fuel grain in a stereolithography method, wherein fuel grains in a plastic matrix are deposited in layers for building a solid fuel rocket body in three dimensions for improved performance and for a compact design, the hybrid rocket motor including buried radial channels for defining a desired burn profile including the oxidizer to fuel burn ratio.

Abstract: A thruster stack assembly and method for manufacturing the same is provided. The thruster stack assembly includes a plurality of grain elements having a common core region (e.g., a channel or passageway), where each grain element comprises a volume of electrically ignitable propellant. The thruster stack assembly further includes electrodes associated with the plurality of grain elements, the electrodes adapted for selectively igniting the plurality of grain elements. In one example, one or more of the grain elements may be ignited and combusted without igniting or damaging adjacent grain elements of the stack. The core region serves to channel combustion gases and exhaust from the thruster stack. Multiple stacks may be assembled together to form three-dimensional thruster arrays.

Abstract: An end-burning grain of a solid rocket motor or other gas-generating device is supplemented with one or more sticks of high-burn-rate propellant embedded in a matrix of a relatively low-burn-rate propellant. The sticks increase the burning surface area as the grain burns by forming conical indentations in the surface.

Abstract: The present invention provides an improved kinematic countermeasure flare and method of constructing the same wherein the housing has integrated internal features. The flare nose weight is set at the end of a die core prior to molding for integration with the housing. Longitudinal grooves and other recesses in the die core allow for the formation of internal longitudinal ribs which propellant can bond to and a retaining bead for retaining the nose weight in the housing which are integrated with the housing. Propellant is cast into the formed housing. A propellant shaping mandrel is then inserted into the formed housing thereby forcing the propellant into the internal cavity created by the integrally molded longitudinal ribs, the shaping mandrel and the wall of the flare housing such that the propellant is bonded to the interior housing wall and the longitudinal ribs.

Abstract: A solid fuel rocket motor, a castellated propellant cartridge and a method of controlling a pressure differential in a cartridge-loaded rocket motor are disclosed. The rocket motor may include a housing having an inside surface, a plurality of propellant cartridges disposed within the housing, an igniter disposed to ignite propellant material within the propellant cartridges, and a nozzle disposed to exhaust combustion gases out of the housing. At least some of the propellant cartridges may be castellated propellant cartridges.

Abstract: A solid fuel rocket burn rate control employing electrical energy preheating of the fuel grain at a plurality of locations dispersed over the grain burn face cross-section as a thrust control arrangement. The electrical energy of the preheating is generated by electrical battery cells which are dispersed across the grain cross-section and are triggered from an inert state into energy generation by burn face heat. The preheating situs moves continually in response to movement of the grain burn face; this is achieved through the use of high temperature electrical battery cell materials responding to the approach of the grain burn face by increased chemical activity and decreased electrical resistance. Preferably, the preheating is directly accomplished by the heat from electrical and other losses in the battery cells.

Abstract: A thruster stack assembly and method for manufacturing the same is provided. The thruster stack assembly includes a plurality of grain elements having a common core region (e.g., a channel or passageway), where each grain element comprises a volume of electrically ignitable propellant. The thruster stack assembly further includes electrodes associated with the plurality of grain elements, the electrodes adapted for selectively igniting the plurality of grain elements. In one example, one or more of the grain elements may be ignited and combusted without igniting or damaging adjacent grain elements of the stack. The core region serves to channel combustion gases and exhaust from the thruster stack. Multiple stacks may be assembled together to form three-dimensional thruster arrays.

Abstract: The invention relates to an increment charge (4) to be placed around a tail shaft of a fin-stabilized mortar projectile (1), the increment charge (4) having a centrally located space for the tail shaft and a mounting opening in the space for mounting the increment charge (4). On the opposite side of the increment charge (4) there is provided a protrusion that fits into the mounting opening of an adjacent increment charge (4) for locking them in relation to each other.

Abstract: A thruster stack assembly and method for manufacturing the same is provided. The thruster stack assembly includes a plurality of grain elements having a common core region (e.g., a channel or passageway), where each grain element comprises a volume of electrically ignitable propellant. The thruster stack assembly further includes electrodes associated with the plurality of grain elements, the electrodes adapted for selectively igniting the plurality of grain elements. In one example, one or more of the grain elements may be ignited and combusted without igniting or damaging adjacent grain elements of the stack. The core region serves to channel combustion gases and exhaust from the thruster stack. Multiple stacks may be assembled together to form three-dimensional thruster arrays.

Abstract: An insulation composition that comprises at least one nitrile butadiene rubber (“NBR”) having an acrylonitrile content that ranges from approximately 26% by weight to approximately 35% by weight and polybenzoxazole (“PBO”) fibers. The NBR may be a copolymer of acrylonitrile and butadiene and may be present in the insulation composition in a range of from approximately 45% by weight to approximately 56% by weight of a total weight of the insulation composition. The PBO fibers may be present in a range of from approximately 3% by weight to approximately 10% by weight of a total weight of the insulation composition. A rocket motor including the insulation composition and a method of insulating a rocket motor are also disclosed.

Abstract: A pyrotechnic charge structure, in particular for use as propellant in a rocket motor, is disclosed. The charge structure has a predetermined longitudinal extension and cross sectional area adapted to the intended purpose. The pyrotechnic charge structure (1; 2; 3; 4) is in cross sectional view in cellular form aimed at having similar wall thickness for all the cells of the charge structure. The wall (6) of the cells comprises the pyrotechnic charge, i.e. the propellant, and each cavity of the cells occupies air or oxygen-rich gas.

Abstract: A protector for propelling charge increments on a mortar shell includes a semi-cylindrical body, the semi-cylindrical body arranged around a bottom and sides of the charge increments; two generally horseshoe shaped end plates attached to a front and a rear of the semi-cylindrical body, the end plates defining openings therein, the tail tube of the mortar fitting in the opening in each end plate, the charge increments being disposed between the two end plates; a wedge-shaped boom attached to the rear end plate, the wedge-shaped boom fitting between two fins on the fin portion of the mortar shell; a stop tab attached to a rear end of the wedge-shaped boom, the stop tab extending substantially perpendicular to a narrow side of the wedge-shaped boom to contact an end of the fin portion thereby minimizing longitudinal motion of the charge increments; and a nose cone attached at the front of the semi-cylindrical body, the nose cone defining an opening at its front end, the middle portion of the mortar shell fitting

Abstract: A disk-shaped propellant module (6) for a propellant charge (2), which can be composed of several modules of this type. To ensure a uniform and symmetrical ignition of the propellant charge (2) for a cartridge with increased charge density (>1.2 g/cm3), the propellant module (6) is provided with at least three axially extending ignition channels (9), uniformly distributed over a partial circle, and at least three radial ignition channels (10) that extend from the inside or interior of the propellant module (6) to its outer edge surface (11). The center of the circle for the axial ignition channels (9) is positioned on the central axis (100) for the propellant module (6) and each of the radial ignition channels (10) intersects with at least one of the axial ignition channels (9).

Abstract: Insulation for a rocket motor is provided, as is a method for insulating a rocket motor. The insulation includes a cured elastomer and vapor-grown carbon fibers dispersed in the cured elastomer. The cured elastomer is preferably formed from a precursor composition comprising an EPDM terpolymer. Generally, the vapor-grown carbon fibers have an internal graphitized tube surrounded by a sheath of vapor-deposited amorphous carbon.

Abstract: An oxidizer package for use in a solid fuel propellant system. The oxidizer package comprises a solid oxidizer in the form of discrete pellets of predetermined geometric shape. The pellets are arranged in an array with spaces amongst the pellets. A holder is provided for maintaining the pellets in the array for receipt of a binder introduced to spaces amongst the array of pellets. Preferred propellants include hydrazinium nitroformate and nitronium perchlorate. This arrangement for the pellets can achieve loadings as high as 85% by weight for the pellets.

Abstract: A modular solid-fuel rocket propellent charge has a combustion-chamber wall, a plurality of separate elements including frozen fuel or oxidizer, and respective rigid casings completely surrounding and enclosing the elements. Respective formations on the casing and on the casings directly engage each other for supporting the elements via the respective casings directly an the wall.

Abstract: A gas generating eject motor includes a case containing an ignitable low temperature gas generant material that does not produce toxic gases upon the combustion thereof. The gas generant material is generally contained with a screen enclosure housed within the case. An igniter is disposed within the gas generant material for selectively igniting the gas generant to thereby generate combustion gases. A nozzle is disposed within an open aft end of the case for focusing and directing the combustion gases generated by the ignited gas generant material. The case is constructed and arranged to be separably attached to the aft end of a rocket to be launched from a launch platform, so that, upon ignition of the gas generant, the combustion gases focused by the nozzle will apply a thrust to the rocket and thereby propel, or eject, the rocket from the launch platform, at which time the combustible propellant of the rocket motor will ignite and the eject motor will be separated from the rocket.

Abstract: The present invention is directed to a propellant composition made from a lacquer, the lacquer comprising: (a) from about 15 to about 70 wt % of an organic solvent; (b) from about 0.1 to about 2.5 wt % of a stabilizer; (c) optionally, from about 0% to about 40 wt % of an energetic plasticizer; (d) optionally, from about 0 to about 10 wt % of a nonenergetic plasticizer; (e) optionally, from about 0 to about 10 wt % water; (f) optionally, from about 0 to about 15 wt % of additional additives; and balance being nitrocellulose; all weight percents based on the total weight of the composition, and wherein the lacquer has a viscosity of less than 10 million centipoise when processed. The present invention is also directed to an efficient and safe method of making perforated propellant grains using the above lacquer composition, as well as propellant grains having ellipsoidal cross sections or outwardly extending ridges.

Abstract: A flashless rocket propellent has a continuous potassium sulfate rod extending though the length of the propellent component. The flashless rocket propellent is particularly useful in the MK 66 Rocket Motor.

Abstract: This rocket motor propellant grain is designed to reduce recession of at least the converging portion of a converging/diverging rocket motor nozzle. The propellant grain has an oxidizer-deficient grain portion constructed and arranged in relation to the nozzle to create, upon ignition of the propellant, a boundary layer of oxygen-deficient combustion products that flows through the nozzle throat passage and over an internal surface region of the nozzle. The boundary layer functions to at least substantially isolate the internal surface region of the nozzle from oxidizers contained in the combustion products generated from portions of the propellant grain having higher concentrations of oxygen than the oxidizer-deficient grain portion. As a result, the oxygen-rich portions of the propellant are unable to react with and thereby cause recession of the nozzle internal surface.

Abstract: A rocket according to the invention includes a motor having a fuel chamber and an oxidizer tank, a nose portion provided forward of the motor, and a nozzle surrounded by fins provided aft of the motor. The fuel chamber is provided with a relatively central solid propellant, and a hybrid fuel surrounds the solid propellant. The oxidizer tank is filled with a reactant and coupled to the forward end of the fuel chamber. A pathway is provided between the tank and the fuel chamber for the passage of reactant therethrough. At least one of a valve and a barrier is coupled in the pathway to prevent passage of the reactant into the fuel chamber until after the solid propellant is at least partially consumed. After the solid propellant is at least partially consumed, the valve is opened and/or the barrier is removed to permit the passage of reactant into the fuel chamber.

Abstract: A method for manufacturing solid rocket motors comprises placing a mandrel having at least one collapsible solid slot former positioned thereon substantially centrally along the axis of a rocket case, casting uncured solid rocket propellant about the mandrel and the collapsible solid slot former, curing the solid propellant, removing the mandrel and pressurizing the propellant and slot former, causing the slot former to collapse, and easily removed from the cured propellant.

Abstract: Solid propellant combustion apparatus such as a rocket motor, incorporating a combustion termination valve for providing two combustion phases from a single propellant charge system, the valve comprising a valve body (50) surrounding an opening (52) in the pressure vessel of the combustion apparatus, a valve member (60), retaining means (65) for releasably retaining said valve member in sealing engagement with the valve body whereby to hold the valve closed during a first combustion phase, valve opening means (68) operable on command for releasing said retaining means whereby to cause the valve member (61) to disengage from the valve body thereby opening the valve to terminate said first combustion phase by rapid de-pressurisation of the combustion gases through said opening, and valve closure means (72) operable on command to move the valve member (61) into sealing engagement with the valve body whereby to close the valve prior to re-ignition of the propellant charge for a second combustion phase.

Abstract: Structure, composition and method in which a fiber reinforcement is impregnated with an interpenetrating network of a light cure resin and a thermosetting resin which are independently curable, enabling autobody component repair, electronic assembly bonding, etc. by selective cure first of one or the other resin, and the support against fluid flow of the uncured resin by the first cured resin.

Abstract: A propellant grain composition for use in solid fuel rocket engines comprises a first reactant component consisting of an oxidizer. A polymeric barrier coating encapsulates the first reactant component and a second reactant component comprising a reducer fuel reactable with the oxidizer is coated thereon, A final polymeric coating is placed over the entire propellant grain to yield a unitary metal fuel-oxidizer propellant grain structure for use as a solid rocket fuel.

Abstract: There is disclosed a propellent having a deterred burn rate. The propellent is a particulate containing an energetic binder base and a thermoplastic burn deterrent. The energetic binder has in impetus (energy) in excess of 200,000 foot pounds per pound (mass) and the thermoplastic burn deterrent is both a solid at room temperature and soluble in an organic solvent. The burn deterrent is gradationally dispersed within the particulate with the greatest concentration of burn deterrent at the particulate periphery.

Abstract: There is disclosed a propellent having a deterred burn rate. The propellent is a particulate containing a nitrocellulose base and a cellulosic thermoplastic deterrent, preferably cellulose acetate butyrate or cellulose acetate propionate. The deterrent is gradationally dispersed within the particulate with the greatest concentration of deterrent at the particulate periphery.

Abstract: With the intention of achieving more stable combustion in solid-fuel ramjets, and particularly in the case of ramjets with spinning application, it is proposed that the inner surface of the ramjet fuel element (8) is provided with raised surfaces, or lands (14), preferably formed from the fuel element material. In one preferred embodiment, the lands (14) are straight and extend parallel with the longitudinal axis of the fuel element and have the same length as the fuel element. The tangenital component of the rotating gasflow generated by rotation, forces the peripheral part of the gasflow to pass over these raised surfaces (14), thereby generating standing, stabilizing vortices downstream of the raised surfaces.

Abstract: An elastomeric prepreg wherein a carrier, whether roving, yarn, broad goods or tape, is impregnated with an uncured elastomeric rubber system which may include the addition of modifiers to enhance the properties of the coating. The prepreg is capable of being spooled and de-spooled for subsequent filament winding or laid up into laminates. In either process no environmental control equipment is required as the volatiles have previously been removed, leaving the elastomer in the generally uncured state.

Abstract: Compositions and methods for increasing the storage life of a propellant grain by reducing the rate of decomposition of its nitrate ester components. A nitrate ester stabilizer is placed in a layer adjacent to a nonburning surface of the propellant grain. The stabilizer migrates from the layer into the grain's nonburning surface to provide additional stabilization for the entire grain, especially portions of the grain adjacent to the protected nonburning surface.

Abstract: Rocket motors are constructed by encasing a solid porous poly(vinyl alcohol) core in the shape of the propellant grain with motor casing material in the same manner as motor casing material is conventionally applied over a disposable or removable mandrel. The motor is then completed by filling the open passages of the core with liquid oxidizer. When the liquid oxidizer permeates the poly(vinyl alcohol), the result is a solid continuous propellant grain of rubbery consistency. Aluminum powder may be incorporated in either the porous poly(vinyl alcohol) or the liquid oxidizer to achieve a fuel-enhanced propellant.

Abstract: Methods and materials for modifying rocket motors to render them safer are disclosed. The invention is primarily related to placing within the bore of a rocket motor a sufficient quantity of a material mitigant such that if the rocket motor receives an unintended impact, the material mitigates the effects of the impact, and prevents ignition or explosion of the propellant. The material mitigant may take a number of forms; however, polymer foams and fibers of various types have been found to work well. It is presently preferred that the material mitigants have a density of not less than approximately 0.04 grams/cc. Examples of such materials include polyurethane foam, polystyrene, and Kevlar.RTM. fibers. It is also preferred that the material mitigant not be reactive with the propellant in the rocket motor.

Abstract: An igniter for solid fuel rockets in which hot gases from an initiator are hanneled to drive a piston against a reservoir of combustion modifying fluid, like ethylene glycol, so as to spray the fluid onto the propellant grain just before ignition. After the piston moves, the hot gases escape the igniter as to ignite the propellant directly or through an intermediate main ignition charge.

Abstract: An interpenetrating network combination of ultraviolet and thermally curable rocket motor liner formulation comprising an optional solid filler, at least one ultraviolet curable polymer, an optional ultraviolet reactive diluent, at least one thermally curable polymer and at least one curing agent.

Abstract: A method is described for improving the mechanical stressability of ammunition bodies, particularly shaped charges, which are highly accelerated and/or subjected to shock waves by coating the contacting portion of the cavity. In a first step of the method, the charge liner is coated on its outer surface and the case is coated on its inner surface with a styrene-containing copolymer; in a further step of the method, the explosive charge is cast or compressed in a known manner. The method finds use particularly for the production of tandem shaped charge projectiles.

Abstract: The invention relates to a low-vulnerability explosive munitions element comprising a casing containing a multicomposition explosive charge, the innermost layer of which is a composite explosive comprising a filled polyurethane or polyester polymer matrix, the filling of which contains more than 20% by weight of organic nitrate explosive, and the peripheral layer of which is a pyrotechnic composition of the family of composite solid propellants comprising a filled polyurethane or polyester polymer matrix the filling of which contains at least one mineral oxidant and less than 17% by weight of organic nitrate explosive. The blast and/or bubble effect produced is close to that produced by the much more-vulnerable charge of monocomposition composite explosive of equivalent mass. The invention also relates to a method for obtaining a blast and/or bubble effect by releasing in the casing of an aforementioned munitions element according to the invention, and then rupture of the casing.

Abstract: A unitary propellant charge module for use in separate ammunition comprises a combustible container having a generally hollow cylindrical shape, a generally cylindrical charge body of interconnected compacted spheroidal propellant grains disposed within the container, each of the grains having an uncompacted propellant grain diameter of at least 100 mils, and at least one generally tubular center core igniter body of interconnected compacted spheroidal propellant grains having an unrolled grain size between about 20 to less than 100 mils in diameter. The igniter body is disposed within a central bore extending along the longitudinal axis of the charge body.

Abstract: The present invention relates to propellent powders in multi-perforated sticks (1) divided by radial slits (7) cutting the peripheral channels (6) of the stick (1) but allowing a continuous cover (8) to remain around its central channel (5). The invention also relates to equipment for obtaining such sticks, comprising:--a drive shaft (17) carrying circular cutting members (18),--a positioning shaft (19) carrying rollers (20) and a thrust shaft (21) carrying rollers (22). The sticks (1) advantageously constitute bundles for artillery ammunition which divide up into elementary particles at the moment of firing.

Abstract: The invention relates to a low-vulnerability explosive munitions element comprising a casing containing a multicomposition explosive charge, the innermost layer of which is a composite explosive comprising a filled polyurethane or polyester polymer matrix, the filling of which contains more than 40% by weight of organic nitrate explosive, and the peripheral layer of which is a pyrotechnic composition of the family of composite solid propellants comprising a filled polyurethane or polyester polymer matrix the filling of which contains at least one mineral oxidant and less than 10% by weight of organic nitrate explosive. The blast and/or bubble effect produced is close to that produced by the much more-vulnerable charge of monocomposition composite explosive of equivalent mass. The invention also relates to a method for obtaining a blast and/or bubble effect by releasing in the casing of an aforementioned munitions element according to the invention, and then rupture of the casing.

Abstract: The trimer of 1,6-hexanediisocyanate, which has the chemical name in acconce with (International Union of Pure and Applied Chemistry) (IUPAC) nomenclature of 7-aza-8-oxo-7[1-isocyanato-6-oxohexyl]pentadecanediisocyanate, is employed in a method of chemical bonding of solid propellant grains to the internal insulation of an interceptor motor. Both the propellant composition and the insulation are chemically reactive with the trimmer identified hereinabove. The method comprises providing a solid propellant rocket motor case having the internal insulation installed therein. Following degreasing of the insulation outer surface, the trimer identified hereinabove is spray coated onto the insulation, and the solid propellant composition is placed onto the trimer. A chemical bond is formed between the propellant and the trimer and the insulation to achieve the chemical bonding of an isocyanate curable solid propellant composition to the internal insulation.

Abstract: Propellant grain consisting essentially of:A. from about 50% to about 75% by weight of an oxidizer consisting essentially of ammonium nitrate;B. from about 1% to about 20% by weight of a cured polymeric binder;C. from 0% to about 30% by weight of a nitrate ester plasticizer;D. from about 3.5% to about 8% by weight of boron in free elemental form; andE. from 0% to about 3% by weight of aluminum metal.The grain has a pressure exponent not exceeding about 0.30 at a combustion pressure between about 2500 and about 7000 pounds per square inch (between about 1700 and about 4800 N/cm.sup.2). Also, an uncured pourable slurry having the formula stated above, except that the binder is uncured. Also, a method for forming a propellant grain having the formula stated above, comprising the steps of providing the uncured pourable slurry defined above, pouring the slurry into a casing, and curing the slurry in situ to form a cured propellant grain having a pressure exponent not exceeding about 0.

Abstract: This invention generally relates to aircraft countermeasures and more particularly to an infrared decoy flare which operates in a substantially aerodynamic configuration and effectively shields aerodynamic cooling associated with its operation when ejected into the adjacent windstream of a tactical aircraft. The decoy provides an enhanced infrared signature to IR and heat-seeking threats to effectively decoy them away from the protected aircraft.

Abstract: A method of making an elastomer lined vessel or portion of such vessel is disclosed. The method comprises causing a thin, tacky ribbon comprising an elastomer (preferably curable) to continuously so encircle a center rotational axis of a mandrel as to adjacently position and tack together integral segments of the ribbon substantially circumferentially relative to this center rotational axis thereby forming a layer of the elastomer about the mandrel which may be the shell of the vessel in case of internal application of the insulator or may be a workpiece around which the shell is built over the so-laid ribbon.

Abstract: A low freezing point eutectic solid propellant plasticizer composition component comprising a mixture of 1,5-dinitrato-3-nitrazapentane and 1,5-diazido-3-nitrazapentane having a eutectic freezing point below about -25.degree. C.

Abstract: A method of installing a polysiloxane insulation in an interceptor rocket motor comprises of applying an adhesive composition, which is compatible with a composite rocket motor case and polysiloxane with which it interfaces; applying a polysiloxane in an uncured state and any added reinforcing fibers to the inside surface of rocket motor case containing the adhesive coating; curing the polysiloxane to form a polysiloxane insulation; casting a solid propellant composition on the cured siloxane insulation; and curing the cast solid propellant composition to complete the method of installing a polysiloxane insulation. The polysiloxane insulation enables the use of a wider selection of reinforcing fibers, even fragile fibers, since the fibers can be incorporated with less energy and the mixing and extruding of the polysiloxane does not destroy the fiber length.

Abstract: An annular bayonet vacuum cup allows a vacuum to be pulled on an annular bayonet and the bayonet to be filled with propellant that is devoid of air. The vacuum cup is transparent/translucent and allows the propellant to be seen as it leaves the exit end of the bayonet. This alerts an operator to cut off the vacuum pressure thereby to keep the excess of propellant flow to a minimum, which excess drops into the vacuum cup and is prevented from contaminating the vacuum line. The vacuum cup is disposable (an O-ring sealing ring being reusable) and eliminates propellant clean-up.

Abstract: Non-asbestos elastomeric insulating materials for rocket motors are disclosed. The insulating materials are low in density (between about 0.035 and 0.050 lb./cubic inch) and comprise crosslinked elastomeric polymers (100 parts by weight of a crosslinkable elastomeric polymer that is substantially saturated and about 5-50 parts by weight of a crosslinkable, substantially unsaturated, elastomeric polymer) in which are dispersed between about 10 and 100 (preferably 15-75) parts by weight of a char forming organic fiber selected from polyaramide pulps, between 1 and 15 parts by weight of a peroxide crosslinker, and between about 0 and 150 parts by weight organic and/or inorganic particulate. The insulating materials issue little smoke, have notable erosion resistance and can be tailored to have thermal, mechanical and other properties of desired character. Ingredients such as supplemental elastomers, oils and lubricants enhance the processability and green properties such as tack of the insulating materials.

Abstract: A method of and apparatus for casting uncured solid propellant in solid propellant rocket motors includes a casting bayonet having annular slits formed in the exit end thereof that improve the fluidity and self-leveling behavior of the uncured propellant. Both of these effects eliminate or reduce the trapping of air in the cast propellant, thus providing a better product.