Abstract: A method to form a sealed, partially or fully densified fiber reinforced composite preform which comprises: impregnating said preform with a major portion of a liquid, ceramic-forming precursor; initiating cure of the outer surface of said precursor with radiation, such as actinic radiation; a thereby forming a sealing barrier or skin on the surface of said preform. In a second embodiment the preform is infiltrated with a thermoplastic as liquid, frozen, excess frozen infiltrate optionally removed from the preform surface, then the preform is coated with a liquid which is curable with radiation, cure is initiated with radiation and the preform is sealed with a barrier coat.

Abstract: A composite tubular article such as a nozzle or a polar boss for a rocket motor wherein straight matrix impregnated yarn portions extend from one end portion to the other in a non-cylindrical ply to achieve increased strength while affording lighter weight. The matrix material may be at least partially curable by actinic radiation to anchor the yarn portions in position as they are applied on a mandrel so that radial rods or the like for anchoring the yarn may be eliminated. For high performance rocket motor nozzle components and the like, the yarn portions may be straight or non-straight and applied by braiding or the like, and the matrix material may be decomposed to a char adequate to anchor the yarn portions and subjected to one or more densification cycles whereby the matrix material may be carbonized.

Abstract: A nozzle assembly including a nozzle structure and liner is disclosed. The nozzle structure is made of at least one carbon-based material and includes a nose tip region, a restricted cross-sectional throat region, and an exit cone region that collectively provide an interior surface configured to define a converging-diverging pathway. The liner includes leg and body portions. The leg portion protrudes over an edge or into a groove of the nozzle structure to engage the liner to the nozzle structure. The body portion of the liner covers at least the throat region of the nozzle structure along the flow path to obstruct high temperature combustion products from causing recession of the nozzle structure. The erosion-resistant liner has at least one irregularity that extends, in a continuous manner, radially at least along the leg portion and, optionally, longitudinally along the body portion.

Abstract: One embodiment of this novel rocket system employs at least one rocket motor closure that includes low and high load capability retainers that, respectively, provide an insensitive munitions system during storage and positive locking of a nozzle assembly at lunch. Another embodiment of this novel rocket involves an insensitive munitions approach that does not require use of such retainers by developing different normal and abnormal burn paths. The rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well varying gas pressure inherent in gun-launched systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: A rocket system employing at least one rocket motor closure that includes low and high load capability retainers that, respectively, provide an insensitive munitions system during storage and positive locking of a nozzle assembly at launch. The rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well as varying gas pressure inherent in gun-launched rocket systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: A rocket motor or other gas generator the case of which is ventable at a predetermined temperature substantially greater than the ambient temperature range for storage and operation thereof and substantially less than the auto-ignition temperature of gas generant material contained therein so as to prevent catastrophic propulsion or burst in case of fire. In order to provide such venting without requiring any shaped charges or other auxiliary system for that purpose at all, the case is composed of a matrix-impregnated fibrous material wherein the fibrous material is characterized by having a melting point which is substantially higher than the ambient temperature range and which is substantially lower than the auto-ignition temperature of the gas generant material. The matrix material is preferably actinic radiation-curable at a temperature which is substantially lower than the melting point temperature of the fibrous material.

Abstract: A composite structure for rocket motors includes a longitudinal element of the composite structure, which is an assembly of segments of clamshell or barrel stave form. The element is held in place by a circumferential element, which is a composite overwrap of reinforcing fibers in a resin matrix. The matrix may be thermoplastic or thermosetting. The reinforcing fiber being high strength polyolefin thereby providing the rocket motor with insensitive munitions capability.

Abstract: A membrane of refractory material having opposed arcuate edges one of which is longer than the other, having opposed straight edges which when welded together in edge-to-edge relationship form the frustum of cone, is deformed by brake or die press means over a portion adjacent the longer arcuate edge to provide thereon a pattern comprising two sets of creases in each of five equal sectors, a first set of which creases points at the viewer from the exterior of the membrane (as it subsequently is formed into the frustum of a cone) and a second set which points away from the viewer. The creases of the first set run both circumferentially and longitudinally with respect to the surface of the cone while the creases of the second set run longitudinally only.

Abstract: A rocket propelled vehicle includes a lower stage solid propellant rocket motor and at least one upper stage solid propellant rocket motor. Each of the rocket motors includes a case, a nozzle and a solid propellant having a central combustion chamber, with the case of the lower stage motor, in a preferred embodiment, being open at the forward end and the case of the upper stage motor being open at the aft end and forming a continuation of the open end of the lower stage case, with a layer of slow burning material cast on the surface defining the central combustion chamber of the upper stage solid propellant, and with ignition of all stages being accomplished approximately simultaneously through operation of a single ignition system.

Abstract: A rocket motor extendible nozzle is made by folding a sheet stock (metal, etc.) development of a membrane in the form of a cone or conical frustum (with a straight meridian) in such manner that the aft or larger diameter region of the membrane that is constructed by joining the two longitudinal edges may be collapsed for stowage with vastly reduced overall length, maximum diameter, and with substantially increased rigidity in comparison with the fully deployed unit. In the stowed condition, the material comprising the aft portion of the membrane is situated in multiple layers within the forward portion so that deployment is induced by forward-to-aft gas flow of the rocket motor within the cone.

Abstract: A rocket motor extendible nozzle exit cone, in one embodiment, is made by folding a sheet stock (metal, etc.) development of a membrane in the form of a cone or conical frustum (with a straight meridian) in such manner involving the provision of a plurality of crease lines therein that the aft or larger diameter region of the membrane that is constructed by joining the two longitudinal edges may be collapsed for stowage with vastly reduced overall length, maximum diameter, and with substantially increased rigidity in comparison with the fully deployed unit. In other embodiments, the extendible nozzle exit cone is made as an integral, seamless, free-standing member. In the stowed condition of each of the embodiments, the material comprising the aft portion of the membrane is situated in multiple layers within the forward portion so that deployment is induced by forward-to-aft gas flow of the rocket motor within the cone.

Abstract: A rocket motor extendible nozzle, in a first embodiment of the invention, is made by folding a sheet stock (metal, etc.) development of a membrane in the form of a cone or conical frustum (with a straight meridian) in such manner that the aft or larger diameter region of the cone that is constructed by joining the opposed edges may have the aft portion collapsed for stowage in a single layer of fluted form contained between two planes perpendicular to the axis of the cone with reduced overall length, and the forward portion extensively folded for substantially increased rigidity in comparison with the fully deployed unit. A second embodiment of the invention differs from the first embodiment in the further development of the membrane such that the aft end of the cone may be collapsed for stowage in multiple layers with vastly reduced overall length while retaining the increased rigidity of the forward portion in the stowed condition.

Abstract: A rocket motor extendible nozzle, in a first embodiment of the invention, is made by folding a sheet stock (metal, etc.) development of a membrane in the form of a cone or conical frustum (with a straight meridian) in such manner that the aft or larger diameter region of the cone that is constructed by joining the opposed edges may have the aft portion collapsed for stowage in a single layer of fluted form contained between two planes perpendicular to the axis of the cone with reduced overall length, and the forward portion extensively folded for substantially increased rigidity in comparison with the fully deployed unit. A second embodiment of the invention differs from the first embodiment in the further development of the membrane such that the aft end of the cone may be collapsed for stowage in multiple layers with vastly reduced overall length while retaining the increased rigidity of the forward portion in the stowed condition.

Abstract: A thrust reversal system for reducing, neutralizing and/or reversing the thrust vector resultant of a solid propellant rocket motor utilizes a movable conical piston or plenum cover that fits over a plurality of small ports that are arranged in a circumferential band around the aft dome of the rocket motor. Before actuation, the cover seals the ports. When actuated, the cover is retracted aft a pre-set distance to uncover the ports, shock and the stroke time being controlled by a shock absorber which in one embodiment of the invention is of the hydraulic type and in other embodiments is of the mechanical type. The cover becomes a baffle which directs the combustion chamber gas exiting the ports outward and forward. Suitable control of the flow characteristics may be exercised by the setting of the travel distance of the piston cover and the number, size and orientation of the ports.