Abstract: A rocket motor includes at least one inert rod. The inert rod has a groove that extends along the length of the inert rod. The groove may be machined by a lathe, die, and/or CNC machine, such that the groove is configured helically around the rod. A reactive wire is inserted into the groove along the length of the inert rod. The grooved inert rod, and the reactive wire together constitute the burn rate enhancer assembly. The rocket motor is configured such that the burn rate enhancer assembly is inserted into a rocket motor casing. The rocket motor casing is then filled with a burnable propellant grain, the highly loaded grain, which is in a liquid or semi-solid state. The highly loaded grain then cures in the rocket motor casing around the burn rate enhancer assembly.

Abstract: A flight test system uses a flight termination destruct charge that is configured to overpressurize a pressure vessel in a rocket motor to terminate thrust. The flight termination destruct charge is an electroexplosive detonator arranged on a final burn surface of a propellant contained in the pressure chamber. In a multi-pulse rocket motor, one of the pulses is ignited by the activation of the detonator. The activated detonator is configured to ignite the propellant grain without venting of the gas resulting from the burning of the propellant. Due to the burning of the propellant, the surface area in the pressure vessel is increased which causes increased pressure in the pressure vessel until a critical pressure is reached. When the critical pressure is reached, the rocket motor casing structural capabilities are exceeded. The overpressurized rocket motor casing then ruptures and thrust of the rocket motor is terminated.

Abstract: A method for non-destructively determining a mechanical property of a solid rocket motor propellant grain may comprise applying, via a gas, a force to a surface of the solid rocket motor propellant grain, wherein a deformation is formed on the surface of the solid rocket motor propellant grain in response to the applying, and measuring a pressure of the gas. This process may be performed over time to determine a lifespan of the propellant grain.

Abstract: A satellite has thrusters that are integral parts of its frame. The frame defines cavities therein where thrusters are located. The thrusters may include an electrically-operated propellant and electrodes to activate combustion in the electrically-operated propellant. The frame may be additively manufactured, and the propellant and/or the electrodes may also be additively manufactured, with the frame and the propellant and/or the electrodes also being manufactured in a single process. In addition the thrusters may have nozzle portions through which combustion gases exit the thrusters. The thrusters may be located at corners and/or along edges of the frame, and may be used to accomplish any of a variety of maneuvers for the satellite. The satellite may be a small satellite, such as a CubeSat satellite, for instance having a volume of about 1 liter, and a mass of no more than about 1.33 kg.

Abstract: A method for non-destructively determining a mechanical property of a solid rocket motor propellant grain may comprise applying a force to a surface of the solid rocket motor propellant grain, wherein a deformation is formed on the surface of the solid rocket motor propellant grain in response to the applying, and calculating a value of the mechanical property of the solid rocket motor propellant grain based on the deformation. This process may be performed over time to determine a lifespan of the propellant grain.

Abstract: A propellant is made from a flexible sheet that in some examples is nitrocellulose. An ignitable material is deposited on one side of the flexible sheet. The ignitable material is a series of triangles having a base adjacent to one edge of the sheet, and an apex adjacent to the other side of the sheet. Some examples of the ignitable material may be thermite compositions. The flexible sheet is rolled around a nonburnable tube and placed within a firearm casing, with the triangle bases being adjacent to the back of the casing, and the triangle apexes being adjacent to the front of the casing. The nonburnable tube is disposed over the primer pocket, so that ignition products from the primer travel through the tube, igniting the propellant adjacent to the front of the casing.

Abstract: A hybrid-like liquid fuel motor (the “motor”) may include a port surrounded by a wall. Surrounding the wall are a plurality of chambers and segmented walls to separate the chambers. In some instances, a single helix chamber may surround the wall, and may operate similar to that of a segmental chamber. During operation of the motor, gas flows from one end of the port to another end of the port. As the walls surrounding the port begin to disintegrate, liquid fuel within chambers begins to begin to mix with the flow of gas. As the segmented walls between the chambers begin to disintegrate, liquid from the other chambers begin to mix with the flow of gas, creating a metering of the liquid fuel.

Abstract: A pressure vessel includes a shell, a first boss, a second boss, and a reinforcement support. The shell defines an internal cavity. The first and second bosses are disposed within the cavity and respectively extend through opposing longitudinal ends of the shell. The first boss defines a longitudinally extending central orifice. The reinforcement support is disposed within the cavity, is secured to the first boss radially outward of the central orifice, extends from the first boss to the second boss, and is secured to the second boss.

Abstract: Certain embodiments of the invention may include systems, methods, and apparatus for providing a multi-fuel hybrid rocket motor. According to an example embodiment of the invention, a method is provided for producing a multi-fuel hybrid motor. The method can include forming a body, where the body includes one or more intake ports; one or more exit nozzles; one or more channels connecting the one or more intake ports with the one or more exit nozzles; and a plurality of cavities comprising segment walls in communication with the one or more channels. The method also includes depositing a propellant fuel within the plurality of cavities, wherein at least a portion of the propellant fuel is exposed to the one or more channels and wherein the propellant fuel has a higher burn consumption rate than the segment walls.

Abstract: Embodiments of a flight vehicle are provided, as are embodiments of a method for manufacturing a flight vehicle. In one embodiment, the flight vehicle includes a solid-propellant rocket motor, control circuitry, and an electrically-interconnective support structure. The electrically-interconnective support structure includes a load-bearing frame and a plurality of electrical conductors embedded within the load-bearing frame. The solid-propellant rocket motor is mounted to the load-bearing frame, and the plurality of electrical conductors embedded within the frame electrically couples the solid-propellant rocket motor to the control circuitry.

Abstract: The present invention provides a motor including a combustion chamber, an oxidizer injector, a vortex generators and a nozzle. The combustion chamber can be used to dispose a solid fuel, and the oxidizer injector is used to control the flow rate of an oxidizer and to inject the oxidizer into the combustion chamber. The vortex generators is disposed on the inner wall of the combustion chamber for generating eddies to enhance the mixing of the fuel and the oxidizer. Additionally, the nozzle is connected with the combustion chamber for exhausting the gas generated by the combustion of the propellants.

Abstract: The invention relates to a pressure vessel consisting of a housing and an insulating layer lining the inner surface of the housing. The housing is formed from a fiber composite material which has a high mechanical and thermal resistance and the insulating layer is formed from a gas-tight silicone elastomer, with the housing and the insulating layer being bonded to each other. The invention also relates to a method for the manufacture of such pressure vessels.

Abstract: A hybrid rocket motor includes a supply of oxidizer, a first solid fuel element positioned around the supply of oxidizer, a second solid fuel element positioned concentrically around the first solid fuel element, and a combustion port positioned between the first and second solid fuel elements. The oxidizer interacts with the first and second solid fuel elements within the combustion port to produce a combustion product. A nozzle is in communication with the combustion port for combustion discharge of the combustion product.

Abstract: Embodiments of the invention relate to systems and methods for fabricating hybrid rocket motor fuel grains. In one embodiment, a method for fabricating a rocket motor fuel grain can be provided. The method can include providing a platform operable to support a rocket motor fuel grain during fabrication. The method can also include disposing at least one fuel material onto the platform to form a material layer. Further, the method can include successively disposing additional fuel material onto the material layer in small amounts, wherein the rocket motor fuel grain is formed on the platform.

Abstract: Certain embodiments of the invention may include systems, methods, and apparatus for providing a multi-fuel hybrid rocket motor. According to an example embodiment of the invention, a method is provided for producing a multi-fuel hybrid motor. The method can include forming a body, where the body includes one or more intake ports; one or more exit nozzles; one or more channels connecting the one or more intake ports with the one or more exit nozzles; and a plurality of cavities comprising segment walls in communication with the one or more channels. The method also includes depositing a propellant fuel within the plurality of cavities, wherein at least a portion of the propellant fuel is exposed to the one or more channels and wherein the propellant fuel has a higher burn consumption rate than the segment walls.

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: Embodiments of a flight vehicle are provided, as are embodiments of a method for manufacturing a flight vehicle. In one embodiment, the flight vehicle includes a solid-propellant rocket motor, control circuitry, and an electrically-interconnective support structure. The electrically-interconnective support structure includes a load-bearing frame and a plurality of electrical conductors embedded within the load-bearing frame. The solid-propellant rocket motor is mounted to the load-bearing frame, and the plurality of electrical conductors embedded within the frame electrically couples the solid-propellant rocket motor to the control circuitry.

Abstract: Propulsion systems and method for making a propulsion system include additively manufacturing a casing body into a single-piece structure having no bonded or bolted joints. The casing body defines a combustion chamber therein and is at least partially composed of a material useful as a solid rocket fuel and capable of being consumed during combustion. Other embodiments are also described.

Abstract: There is disclosed a solid fuel rocket motor which may include a case and a nozzle coupled to the case. A plurality of fuel pellets may be disposed within the case. An igniter may be disposed to ignite at least a portion of the fuel pellets. A pellet retainer may be positioned within the case to retain the plurality of fuel pellets within the case. The pellet retainer may be perforated to allow exhaust gases to flow from the ignited fuel pellets to the nozzle while preventing unburned fuel pellets from being expelled through the nozzle.

Abstract: An integral composite rocket motor nozzle. The novel nozzle includes a first layer of a first reinforcement material, a second layer of a second reinforcement material, and a common matrix material surrounding the first and second reinforcement materials such that the reinforcement materials and matrix material form a single integral composite structure. In an illustrative embodiment, the first reinforcement material includes graphite fibers for providing structural support, and the second reinforcement material includes glass or quartz fibers for providing thermal insulation on a first side of the first layer. The nozzle may also include a third layer of a third reinforcement material for providing thermal insulation on a second side of the first layer. In a preferred embodiment, the first layer is shaped to form an integrated dome and nozzle structure.

Abstract: A fixed sized bell rocket nozzle is lined with a layer of combustible material that is ignited during launch ignition and burns to outgas into the rocket exhaust for spatially variably confining the exhaust and perfecting an effective variably sized altitude compensating exhaust nozzle that maximizes lift during the launch of a spacecraft into orbit.

Abstract: There is disclosed a solid fuel rocket motor which may include a case and a nozzle coupled to the case. A plurality of fuel pellets may be disposed within the case. An igniter may be disposed to ignite at least a portion of the fuel pellets. A pellet retainer may be positioned within the case to retain the plurality of fuel pellets within the case.

Abstract: A rocket motor case includes a plurality of plank sections which provide a structural key for locking metal and/or composite end enclosures in place as part of the motor case. The plank sections are sized longitudinally to form venting spaces in the motor case sidewall to allow passive venting in the event of propellant ignition during cook-off. The plank sections have plank legs which are surrounded by tailorable roll-bonded sheet material to complete formation of the motor case sidewall.

Abstract: There is disclosed a solid fuel rocket motor which may include a grain disposed within a case. Prior to ignition, a burnable surface of the grain may be defined, at least in part, by a non-circular center perforation. The interior surface of the case may conform to the expected shape of the burnable surface of the grain at a predetermined time after ignition.

Abstract: An insulation material comprising a low-density EPDM polymer, at least one flame-retardant, and an organic filler. The insulation material is used in an insulation layer of a rocket motor. The organic filler is a polymeric, organic filler such as polyvinyl chloride. A rocket motor comprising an insulation material is also disclosed. The insulation material comprises a low-density EPDM polymer, at least one flame-retardant, and a polymeric, organic filler and is applied between an inner surface of a case of the rocket motor and a propellant. A method of insulating a rocket motor is also disclosed.

Abstract: A long range artillery shell having a rocket motor which comprises a combustion chamber (shown filled with solid propellant 9, burn inhibitor 10, igniter material 13 and propellant support 11,12) which has at one end a plenum chamber 8 and a rocket nozzle 14 for venting the plenum chamber 8. The propellant 9 is arranged in an end-burn configuration and housed within the combustion chamber and the propellant support 11,12 is located between the propellant 9 and the plenum chamber 8 and is capable of preventing entry of unburnt propellant into the plenum chamber 8 while permitting, in use, the substantially unhindered transfer of propellant combustion products to the plenum chamber 8.

Abstract: A case-burning rocket booster includes a combustible case containing a solid rocket fuel, a combustion chamber adjustably coupled to the case for burning the case and the solid rocket fuel, a nozzle connected to the combustion chamber for expelling the burned case and fuel to generate thrust, and a drive system for pulling the combustion chamber and nozzle up the case as the case and fuel burn. A hybrid version of the case-burning rocket also includes an oxidizer supply system for supplying varying amounts of oxidizer to the combustion chamber to vary the thrust generated by burning the case and the fuel.

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 rocket motor insulation including an elastomer-based polymer is improved in its processability by the addition of silica particles treated with a hydrophobic coating. The insulation also preferably includes a metallic coagent cross-linker, which when used in combination with the hydrophobic silica particles increases the tear strength and the elasticity of the insulation, while at the same time not adversely affecting the bonding characteristics of the insulation.

Abstract: A vehicle includes at least one polyoxymethylene structural support member. The polyoxymethylene structural support member includes a polyoxymethylene component that is a propellant that provides thrust to the vehicle upon pyrolysis or combustion of the polyoxymethylene component of the product of pyrolysis of the polyoxymethylene component. The vehicle can be a satellite or other type of spacecraft.

Abstract: The present invention comprises a device which is inert and “thermally-equivalent” to actual ordnance. The device can travel with live ordnance and track the propellant temperatures in order to get a more precise propellant temperature for the ordnance. The device comprises a thermally equivalent inert grain instrumented with thermocouples, connected to a data recorder, and packaged in scaled-down ordnance hardware. The hardware is scaled down to enable the device to more easily travel with live munitions.

Abstract: A combustion chamber, in particular for a rocket drive, comprises at least one jacket made of a composite material with a ceramic matrix. The composite material contains a fibrous structure made of carbon-containing fibers, and the fibrous structure comprises layers of fibers that form a three-dimensional matrix.

Abstract: A rocket motor insulation including an elastomer-based polymer is improved in its processability by the addition of silica particles treated with a hydrophobic coating. The insulation also preferably includes a metallic coagent cross-linker, which when used in combination with the hydrophobic silica particles increases the tear strength and the elasticity of the insulation, while at the same time not adversely affecting the bonding characteristics of the insulation.

Abstract: A rocket motor insulation including an elastomer base polymer is improved in its processability by the addition of silica particles treated with a hydrophobic coating. The insulation also preferably includes a metallic coagent cross-linker, which when used in combination with the hydrophobic silica particles increases the tear strength and the elasticity of the insulation, while at the same time not adversely affecting the bonding characteristics of the insulation.

Abstract: This rocket motor propellant includes a combustible double-base propellant and non-carbonized, non-graphitized polymeric fibers dispersed in the double-base propellant. The double-base propellant is formed from a composition comprising nitrocellulose and at least one nitrate ester. Representative polymeric fibers include polyethylene, polypropylene, polyesters, polyamides, polyacrylonitriles and combinations thereof.

Abstract: An elastomerized phenolic resin ablative insulation particularly suitable for use in connection with the thermal insulation of selected components of rocket motors. A composition for making the elastomerized ablative insulation is disclosed. Furthermore, an associated method of forming calendered sheets of material formed of the composition is disclosed. The preferred ingredients of the disclosed composition include acrylonitryle butadiene rubber, zinc borate, and phenol formaldehyde resin which can be cured and bonded to structures such as domes of open-ended rocket motors and other rocket motor components. The subject elastomerized ablative insulation is well suited for use independently or in connection with other insulative materials to form a thermal barrier which is highly resistant to the heat and the erosive nature associated with the combustion of propellant fuels, for example.

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: A towing rocket motor assembly has a tubular solid propellant grain defined by exterior surface area and interior surface area. The exterior surface area is inhibited from ignition within a tubular region of a housing. The housing includes a first end and a second end at either end of the tubular region. The first end defines a plenum having an annular region defined by a diameter larger than that of the tubular region. A plurality of holes formed in the first end communicate with the annular region and are directed substantially in the direction of the second end of the housing. A burn-inhibiting baffle is fitted in the propellant grain wherein an annular spacing is defined between the baffle and the interior surface area of the propellant grain. An ignition assembly, mounted at the second end of the housing, is coupled to the interior surface area of the propellant grain so that the propellant grain burns from the inside out upon ignition.

Abstract: A solid rocket motor for accelerating a payload comprises a motor casing and a solid propellant matrix, utilizing a high burn-rate fuel. The use of a high burn rate fuel allows the rocket motor to operate in an end-burning fashion without the use of a constricting aperture to increase the back-pressure upon the burn-front of the fuel matrix. The exhaust gas produced from combustion of the propellant matrix exits directly to the ambient environment through a simple aperture without the use of an expansion nozzle. By eliminating the mass of the nozzle and allowing the use of lighter, less structurally robust motor casings, the needed acceleration of the vehicle can be achieved while using less propellant and a lighter launch vehicle.

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: The present invention comprises a direct replacement for a charge support spacer currently used in rocket motors. The present invention comprises a vented spacer that cools combustion gases through contact and expansion prior to equalizing pressure by transferring these gases from the inside of the propellant grain to the outside of the grain. The vented spacer takes gases from combustion of the propellant from the core flow inside the propellant grain and cools these gases by contact with and ablation of the spacer material. The gases are cooled to the point that they present no hazard to the rocket motor casing or the outside of the propellant grain. The vented spacer accomplishes these tasks in the same amount of area as the current charged support spacer.

Abstract: Modular, cryogenic, solid rocket propellants of different fuel components, such as fuels, oxidizers, energy-increasing admixtures, binders, additives, etc., for all applications of solid rockets produce a uniform, stable and complete combustion, This objective is accomplished since one of the fuel elements, due to the special selection of its composition, is provided as a permanent igniter generator of the modular propellant.

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: The present invention relates to a low-vulnerability solid-propellant motor (1) comprising a charge (11) of solid propellant, a heat shield (2) covering the inside of a metal structure with a front end wall (3) and a rear end wall (4) which carries at least one nozzle (10), the said front and rear end walls being fastened to a cylindrical shell (5) having wide apertures which are circumferentially distributed and temporarily sealed by sealing plugs (8). The object of the present invention is to provide a motor which has a simple structure and avoids the use of detection and actuation means for freeing the plugs. The motor according to the invention is such that the sealing plugs (8) are held in the apertures between the heat shield and a composite winding (7) of fibres impregnated in a matrix which surrounds the cylindrical shell (5).

Abstract: The present invention concerns an erosion resistant, low signature liner for solid propellant rocket motors and rocket motors having such a low signature liner. The liner combustion products in the exhaust plume have low visibility in the IR and visible light spectra which is essential for use in low signature solid propellant rocket motors. The liner also forms an in-situ protective layer at elevated temperatures for enhanced erosion resistance. These erosion resistant properties eliminate the need for supplemental case insulation which would contribute additional smoke and combustion products to the exhaust plume as well as add additional costs and inert weight to the motor.

Abstract: The disclosed launch systems are formed from mass producible elements that integrate propulsion, pneumatics and structural systems. In one embodiment, a launch system (10) includes multiple stages (12-16) where each stage includes a number of generally wedge-shaped segments (54). Alternating segments (54) form liquid fuel and oxidizer tanks. Each fuel and oxidizer tank pair is associated with a thruster (70). The resulting plurality of thrusters (70) are disposed about a common aerospike thrust structure (30-34). The segments (54) are connected to an internal, hollow structural spine (26) that can also interface with an internal umbilical tower (46). The segments (54) together with their associated thrusters and related components can be mass produced, thereby increasing competition among suppliers, reducing design costs and timeframes.

Abstract: A solid fuel propellant motor casing is described, the casing comprising a body portion for containing the propellant, the body portion also including at least one detachable end closure portion, the closure portion being maintained in position relative to said body portion by retaining means, the retaining means being held in position with engagement means on an inner surface of said body portion and located radially with respect to said body portion by locating means against resilient biasing means, said locating means being responsive to temperature such that said locating means melt at a predetermined temperature allowing said resiliently biased retaining means to move radially inwardly of said body portion and allow said at least one closure portion to be ejected from the casing.

Abstract: The present invention is directed to a novel pressure vessel and process for manufacturing the pressure vessel. In one embodiment, the invention includes a case for a rocket motor which includes a cylindrical, metal shell which has been configured with a plurality of slots positioned in axially extending rows. A fiber reinforced plastic layer is attached to the metal shell. Each slot has a length (L) and is separated by adjacent slots in the same row by an axial distance (A) with adjacent rows of slots separated by a circumferential distance (C) and configured such that L+A.gtoreq.6C and L.gtoreq.4A. The method of the invention comprises using a CO.sub.2 laser to cut a keyhole in the shell and then cutting axially outwardly from the keyhole to form the slot, resulting in slots having a width equal to the cutting width of the laser. The beam is then defocused and used to heat a portion of the metal shell surrounding the ends of each slot to produce a tempered zone.

Abstract: A pulse detonation engine is provided with at least one detonation combustor selectively coupled to an air inlet and fuel source. The detonation combustors are equipped with either active or passive means to dissipate the heat of detonation. Fuel, air, and an oxidizer can be fed to said detonation combustors either through a rotary valve or through a conically shaped injector head.

Abstract: A self-aligning adapter used for attaching a booster to a missile includes a hollow tapered receptacle positioned on the missile. A tapered member sized and shaped to fit into and mate with the receptacle is positioned on the booster. The tapered member fits into the tapered receptacle to align and attach the booster to the missile. A linear explosive charge is positioned on the booster around where the tapered member is joined to the booster. Explosion of the charge severs the tapered member from the booster releasing the booster from the missile.