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: An insensitive combined cycle missile propulsion system includes a solid fuel contained within a first section of the missile, a liquid oxidizer contained within a second section of the missile and a solid oxidizer contained within a third section of said missile. A first conduit has a first valve communicating the fuel and the oxidizer and a second conduit, spatially removed from the first conduit, has a second valve communicating the fuel and the oxidizer. An inlet system for delivering atmospheric oxygen for combustion with the fuel rich gases generated within the missile and a nozzle exhausts combustion products that result from combustion of the fuel, the liquid and solid oxidizers, and air.

Abstract: The present invention describes a hybrid rocket motor that includes a first solid reactant and at least one thrust nozzle and at least one moveable combustion control member within the hybrid rocket motor that restricts the contact of a first fluid reactant in the combustion chamber. In this way, it is then possible to regulate the exposure of the solid reactant to the fluid reactant and thus control thrust.

Abstract: The hybrid rocket system of this invention is characterized by use of an oxidizer tank having a cylindrical midsection surrounded by a skirt and bonded thereto by a layer of elastomeric adhesive. The skirt outer surface is in turn adhesively secured to a spacecraft inner surface. An elongated solid-fuel motor case is mechanically rigidly secured to a central rear surface of the tank, and the case terminates in a throat and nozzle. The elastomeric-adhesive bonding of tank to skirt, and rigid adhesion of skirt to spacecraft forms the sole support for the rocket system, and separate support for the motor case is not required.

Abstract: Activation of a propellant in a constant volume container causes a phase change material to rapidly expand so that the pressure in the container increases. Programmability and sequential actuation are enabled by patterning the phase change material into the integrated device. The pressure generated may be used to activate an energy transducer such as a high pressure turbine, a piezoelectric material, and an elastic strain material. This provides a hybrid actuation system of electrical energy, pneumatic and hydraulic power. The pressure change in the constant volume container is also harnessed to provide a microbattery.

Abstract: A heated O-ring that has particular application for providing sealing in the connector region of a compressed hydrogen storage tank in a fuel cell system. In one embodiment, an electrical heating wire is wound through the O-ring so that resistive heating is provide by applying an electrical potential to the wire so that the temperature of the O-ring does not decrease below a predetermined temperature. In another embodiment, electrical heating elements are provided adjacent to and in contact with the O-ring, where an electrical potential applied to the heating elements causes the heating elements to maintain the temperature of the O-ring.

Abstract: Propulsion from combustion of solid propellant pellet-projectiles for providing a useful propulsion that has the advantages of both solid and liquid propulsion engines, and also can make use of either solid chemical propellants or fissionable nuclear material as the fuel. Preferred methods and systems can include a storage chamber for storing solid propellant pellets, a feeding system having a pellet feeding channel and a pellet feeding mechanism connected to the storage chamber and to a gun assembly, which is positioned along a longitudinal axis to eject the pellets at a certain velocity. A triggering system positioned between gun assembly and thrust chamber can initiate the propellant pellet-projectile, and a thrust chamber having a combustion chamber for combustion of propellant pellet-projectile with an exhaust nozzle.

Abstract: Disclosed is a propulsion system having a structural configuration that provides easy and convenient access to the interior regions of a liquid fuel tank and a hybrid rocket motor case. In one embodiment, the propulsion system comprises: a hybrid rocket motor case and a fuel tank coupled to the hybrid motor case. The motor case is configured to hold solid rocket fuel and the fuel tank defines an internal volume configured to hold a fluid oxidizer. A bulkhead is interposed between the motor case and the fuel tank, wherein at least one access passageway extends through the bulkhead. The access passageway provides exterior access to the interior volume of the motor case or the internal volume of the storage tank while the hybrid rocket motor is coupled to the fuel tank.

Abstract: Propellant management systems and methods are provided for controlling the delivery of liquid propellants in a space launch vehicle utilizing multiple rockets. The propellant management systems and methods may be configured to enable substantial simultaneous depletion of liquid propellants in each of a plurality of active rockets during operation of various booster stages of the launch vehicle.

Abstract: Disclosed is a propulsion system having a structural configuration that provides easy and convenient access to the interior regions of a liquid fuel tank and a hybrid rocket motor case. The system operates with a high oxidizer-to-fuel ratio and a high bulk density propellant combination that has a near uniform specific impulse over a large oxidizer-to-fuel ratio range. The system has an increased propellant mass fraction and reduced propellant residuals. This improves the performance of the hybrid propulsion system.

Abstract: Propulsion from combustion of solid propellant pellet-projectiles for providing a useful propulsion that has the advantages of both solid and liquid propulsion engines, and also can make use of either solid chemical propellants or fissionable nuclear material as the fuel. Preferred methods and systems can include a storage chamber for storing solid propellant pellets, a feeding system having a pellet feeding channel and a pellet feeding mechanism connected to the storage chamber and to a gun assembly, which is positioned along a longitudinal axis to eject the pellets at a certain velocity. A triggering system positioned between gun assembly and thrust chamber can initiate the propellant pellet-projectile, and a thrust chamber having a combustion chamber for combustion of propellant pellet-projectile with an exhaust nozzle.

Abstract: Disclosed is a propulsion system for a spacecraft. The propulsion system includes a supply of oxidizer and at least one nozzle. A conduit fluidly couples the supply of oxidizer and the nozzle. The conduit provides a pathway for oxidizer to flow in a downstream direction from the supply of oxidizer toward and into the nozzle. A pressure regulator is coupled to the conduit and is interposed between the supply of oxidizer and the nozzle, wherein the pressure regulator regulates the pressure of oxidizer flowing through the conduit and downstream of the pressure regulator to a pressure at or below the pressure required to maintain the oxidizer in a gas state to ensure that the any oxidizer flowing through the conduit is in a gas state prior to entering the nozzle. The conduit supplies oxidizer from the supply of oxidizer to a hybrid rocket motor.

Abstract: A super-staged rocket includes at least approximately 50 rocket engines, where the engines are distributed according to at least one of: at least five multi-engine stages connected in series, each stage including at least ten engines connected in parallel; and at least five multi-stage units connected in parallel, each unit including at least five engines connected in series.

Abstract: Propulsion from combustion of solid propellant pellet-projectiles for providing a useful propulsion that has the advantages of both solid and liquid propulsion engines, and also can make use of either solid chemical propellants or fissionable nuclear material as the fuel. Preferred methods and systems can include a storage chamber for storing solid propellant pellets, a feeding system having a pellet feeding channel and a pellet feeding mechanism connected to the storage chamber and to a gun assembly, which is positioned along a longitudinal axis to eject the pellets at a certain velocity. A triggering system positioned between gun assembly and thrust chamber can initiate the propellant pellet-projectile, and a thrust chamber having a combustion chamber for combustion of propellant pellet-projectile with an exhaust nozzle.

Abstract: The engine for a rocket is suitable for using for an educational program. The engine uses a liquid phase propellant having a predetermined boiling point and a heating substance having a temperature higher than the boiling point. The engine has an inner wall having a circumferential surface; an outer wall surrounding the inner wall, the outer wall having an interior surface spaced from the circumferential surface of the inner wall by a predetermined distance such that the space between the circumferential surface and the interior surface form a mixing chamber having an opening; and injector for injecting the liquid-phase propellant and the heating substance into the mixing chamber so that the propellant is evaporated by the heating substance thereby creating a jet stream moving from the opening of the mixing chamber to the outside of the mixing chamber.

Abstract: The invention concerns a rocket engine wherein the combustion chamber includes at least one first monolithic component made of a thermostructural composite material comprising a porous wall through which the fuel is introduced in the core of the combustion chamber. A small part of the fuel is directed towards the neck for it to be cooled.

Abstract: An ignition system for a rocket motor includes a soft plastic tube that extends up into the combustion chamber and is coupled to an oxidizer source. Ignition source wires extend through the tube and terminate at a first end at a location which is set back from the end of the tube, and have a second end coupled to an electric power supply. In operation, an oxidizer is introduced into the tube simultaneously with activation of the power supply. The set back portion of the plastic tube becomes fuel for the oxidizer and is consumed, introducing a fire plume into the combustion chamber. The tube introduces additional fuel distinct from the fuel grain or propellant which is in contact with the both the ignition source wires and oxidizer. In addition, the tube will not damage the nozzle as it is being blown of the rocket during the main propulsion phase.

Abstract: This invention comprises a new process for developing high regression rate propellants for application to hybrid rockets and solid fuel ramjets. The process involves the use of a criterion to identify propellants which form an unstable liquid layer on the melting surface of the propellant. Entrainment of droplets from the unstable liquid-gas interface can substantially increase propellant mass transfer leading to much higher surface regression rates over those that can be achieved with conventional hybrid propellants. The main reason is that entrainment is not limited by heat transfer to the propellant from the combustion zone. The process has been used to identify a new class of non-cryogenic hybrid fuels whose regression rate characteristics can be tailored for a given mission. The fuel can be used as the basis for a simpler hybrid rocket design with reduced cost, reduced complexity and increased performance.

Abstract: A hybrid rocket motor is provided with a precombustionchamber supplied with propellant from separate fuel and oxidizer sources. The propellant can be in the form of gas or liquid and injected substantially tangentially into the head end of the hybrid motor adjacent the oxidizer injector to form a propellant swirl. As the hybrid motor oxidizer is injected into the swirl, it is heated and gasified, and assumes a swirling motion which increases the oxidizer path length and thereby increases the dwell time of the oxidizer. The increased dwell time increases combustion efficiency and permits multiple restarts of the hybrid motor. The propellant may also be a combination of solid and fluid reactants. In one embodiment, the oxidizer injector is extended into the combustion chamber to form a toroidalprecombustion chamber which has an annular nozzle adjacent a face of the oxidizer injector.

Abstract: A hybrid rocket motor is provided with a precombustion chamber supplied with propellant from separate fuel and oxidizer sources. The propellant can be in the form of gas or liquid and injected substantially tangentially into the head end of the hybrid motor adjacent the oxidizer injector to form a propellant swirl. As the hybrid motor oxidizer is injected into the swirl, it is heated and gasified, and assumes a swirling motion which increases the oxidizer path length and thereby increases the dwell time of the oxidizer. The increased dwell time increases combustion efficiency and permits multiple restarts of the hybrid motor. The propellant may also be a combination of solid and fluid reactants. In one embodiment, the oxidizer injector is extended into the combustion chamber to form a toroidal precombustion chamber which has an annular nozzle adjacent a face of the oxidizer injector.

Abstract: A flame holder is provided at a head end of a combustion chamber of a hybrid rocket motor. The flame holder includes a high-temperature casing defining a cavity, and a solid propellant within cavity around or near the injector. The propellant may be provided in an annulus within the casing, such that the flame plume from the burning propellant is substantially parallel to the flow of the oxidizer, or may be generally cylindrical within a cylindrical casing such that the flame plume of the burning propellant is substantially perpendicular to the oxidizer flow. The solid propellant is preferably ignited substantially simultaneously with the ignition of the hybrid motor. The burning of the solid propellant prevents the flame from combustion of the fluid oxidizer and solid fuel in the hybrid motor from drifting and thereby stabilizes the flame front.

Abstract: A hybrid rocket engine and a method for propelling a rocket utilizing a vortex flow field. The flow field includes an outer fluid vortex spiraling toward a closed end of the flow field generating apparatus and an inner fluid vortex substantially concentric with the outer vortex spiraling away from the closed end and toward an outlet opening in which the inner vortex spirals in the same direction as the outer vortex, but in the opposite axial direction. The invention also relates to a rocket propulsion system utilizing the flow field in which the propulsion system includes a combustion chamber with a fuel source and an oxidizer source that deliver the said fuel and said oxidizer to the said outer and inner vortexes in a manner to support a combustion process while flowing along the flow field.

Abstract: A hybrid rocket motor is provided with a precombustion chamber supplied with propellant from separate fuel and oxidizer sources. The propellant can be in the form of gas or liquid and injected substantially tangentially into the head end of the hybrid motor adjacent the oxidizer injector to form a propellant swirl. As the hybrid motor oxidizer is injected into the swirl, it is heated and gasified, and assumes a swirling motion which increases the oxidizer path length and thereby increases the dwell time of the oxidizer. The increased dwell time increases combustion efficiency and permits multiple restarts of the hybrid motor. The propellant may also be a combination of solid and fluid reactants. In one embodiment, the oxidizer injector is extended into the combustion chamber to form a toroidal precombustion chamber which has an annular nozzle adjacent a face of the oxidizer injector.

Abstract: A hybrid rocket motor is provided with a precombustion chamber supplied with propellant from separate fuel and oxidizer sources. The propellant can be in the form of gas or liquid and injected substantially tangentially into the head end of the hybrid motor adjacent the oxidizer injector to form a propellant swirl. As the hybrid motor oxidizer is injected into the swirl, it is heated and gasified, and assumes a swirling motion which increases the oxidizer path length and thereby increases the dwell time of the oxidizer. The increased dwell time increases combustion efficiency and permits multiple restarts of the hybrid motor. The propellant may also be a combination of solid and fluid reactants. In one embodiment, the oxidizer injector is extended into the combustion chamber to form a toroidal precombustion chamber which has an annular nozzle adjacent a face of the oxidizer injector.

Abstract: The present invention comprises a rocket motor nozzle that replaces the fixed-wall throat with a “wall” created by injecting gas radially into the nozzle. By injecting gas into the flow of combustion products that are going through the nozzle, this will deflect the combustion product flow, restricting and accelerating such flow, just as a fixed-wall does in a standard nozzle. However, by using the “gas-wall” described herein, no erosion will result at the restriction area.

Abstract: A hybrid rocket motor includes a storage tank which stores an oxidizer under relatively low pressure, a turbopump preferably directly coupled to an outlet of the storage tank which pressurizes the oxidizer to a relatively high pressure, a combustion chamber including a solid fuel, and an injector between the turbopump and combustion chamber through which the oxidizer is injected into the combustion chamber. According to a preferred aspect of the invention, the turbopump is operated by an expander cycle of a heat exchanger. According to another preferred aspect of the invention, the fluid flowing through the heat exchanger is oxidizer tapped from the storage tank. A barrier is maintained between an oxidizer feed line from the turbopump and the injector until sufficient pressure is created by the turbopump to pump the oxidizer at the requisite pressure into the injector.

Abstract: This invention comprises a new process for developing high regression rate propellants for application to hybrid rockets and solid fuel ramjets. The process involves the use of a criterion to identify propellants which form an unstable liquid layer on the melting surface of the propellant. Entrainment of droplets from the unstable liquid-gas interface can substantially increase propellant mass transfer leading to much higher surface regression rates over those that can be achieved with conventional hybrid propellants. The main reason is that entrainment is not limited by heat transfer to the propellant from the combustion zone. The process has been used to identify a new class of non-cryogenic hybrid fuels whose regression rate characteristics can be tailored for a given mission. The fuel can be used as the basis for a simpler hybrid rocket design with reduced cost, reduced complexity and increased performance.

Abstract: A hybrid rocket motor is provided with a precombustion chamber supplied with propellant from separate fuel and oxidizer sources. The propellant can be in the form of gas or liquid and injected substantially tangentially into the head end of the hybrid motor adjacent the oxidizer injector to form a propellant swirl. As the hybrid motor oxidizer is injected into the swirl, it is heated and gasified, and assumes a swirling motion which increases the oxidizer path length and thereby increases the dwell time of the oxidizer. The increased dwell time increases combustion efficiency and permits multiple restarts of the hybrid motor. The propellant may also be a combination of solid and fluid reactants. In one embodiment, the oxidizer injector is extended into the combustion chamber to form a toroidal precombustion chamber which has an annular nozzle adjacent a face of the oxidizer injector.

Abstract: The present invention relates to a propulsion system for transporting a crew transfer vehicle. The propulsion system has a casing which defines a chamber, a solid propellant system positioned within the chamber for generating one of emergency escape propulsion during an emergency portion of an ascent flight and orbital injection propulsion during normal flight operations, and a sustain propulsion system communicating with the chamber for sustaining one of the emergency escape propulsion during the emergency portion of the ascent flight and orbital injection propulsion during the normal flight operations. In one embodiment of the present invention, the sustain propulsion system comprises a hybrid solid fuel grain and liquid oxidizer system. In a second embodiment of the present invention, the secondary propulsion system comprises a liquid fuel and liquid oxidizer system.

Abstract: A hybrid rocket engine and a method for propelling a rocket utilizing a vortex flow field. The flow field includes an outer fluid vortex spiraling toward a closed end of the flow field generating apparatus and an inner fluid vortex substantially concentric with the outer vortex spiraling away from the closed end and toward an outlet opening in which the inner vortex spirals in the same direction as the outer vortex, but in the opposite axial direction. The invention also relates to a rocket propulsion system utilizing the flow field in which the propulsion system includes a combustion chamber with a fuel source and an oxidizer source that deliver the said fuel and said oxidizer to the said outer and inner vortexes in a manner to support a combustion process while flowing along the flow field.

Abstract: A flame holder is provided at a head end of a combustion chamber of a hybrid rocket motor. The flame holder includes a high-temperature casing defining a cavity, and a solid propellant within cavity around or near the injector. The propellant may be provided in an annulus within the casing, such that the flame plume from the burning propellant is substantially parallel to the flow of the oxidizer, or may be generally cylindrical within a cylindrical casing such that the flame plume of the burning propellant is substantially perpendicular to the oxidizer flow. The solid propellant is preferably ignited substantially simultaneously with the ignition of the hybrid motor. The burning of the solid propellant prevents the flame from combustion of the fluid oxidizer and solid fuel in the hybrid motor from drifting and thereby stabilizes the flame front.

Abstract: A hybrid rocket motor is provided with a precombustion chamber supplied with propellant from separate fuel and oxidizer sources. The propellant can be in the form of gas or liquid and injected substantially tangentially into the head end of the hybrid motor adjacent the oxidizer injector to form a propellant swirl. As the hybrid motor oxidizer is injected into the swirl, it is heated and gasified, and assumes a swirling motion which increases the oxidizer path length and thereby increases the dwell time of the oxidizer. The increased dwell time increases combustion efficiency and permits multiple restarts of the hybrid motor. The propellant may also be a combination of solid and fluid reactants. In one embodiment, the oxidizer injector is extended into the combustion chamber to form a toroidal precombustion chamber which has an annular nozzle adjacent a face of the oxidizer injector.

Abstract: A hybrid rocket motor includes a storage tank which stores an oxidizer under relatively low pressure, a turbopump preferably directly coupled to an outlet of the storage tank which pressurizes the oxidizer to a relatively high pressure, a combustion chamber including a solid fuel, and an injector between the turbopump and combustion chamber through which the oxidizer is injected into the combustion chamber. According to a preferred aspect of the invention, the turbopump is operated by an expander cycle of a heat exchanger. According to another preferred aspect of the invention, the fluid flowing through the heat exchanger is oxidizer tapped from the storage tank. A barrier is maintained between an oxidizer feed line from the turbopump and the injector until sufficient pressure is created by the turbopump to pump the oxidizer at the requisite pressure into the injector.

Abstract: A hybrid rocket motor includes a storage tank which stores an oxidizer under relatively low pressure, a turbopump preferably directly coupled to an outlet of the storage tank which pressurizes the oxidizer to a relatively high pressure, a combustion chamber including a solid fuel, and an injector between the turbopump and combustion chamber through which the oxidizer is injected into the combustion chamber. According to a preferred aspect of the invention, the turbopump is operated by an expander cycle of a heat exchanger. According to another preferred aspect of the invention, the fluid flowing through the heat exchanger is oxidizer tapped from the storage tank. A barrier is maintained between an oxidizer feed line from the turbopump and the injector until sufficient pressure is created by the turbopump to pump the oxidizer at the requisite pressure into the injector.

Abstract: A propulsion-assisted projectile has a body, a cowl forming a combustion section and a nozzle section. The body has a fuel reservoir within a central portion of the body, and a fuel activation system located along the central axis of the body and having a portion of the fuel activation system within the fuel reservoir. The fuel activation system has a fuel release piston with a forward sealing member where the fuel release piston is adapted to be moved when the forward sealing member is impacted with an air flow, and an air-flow channel adapted to conduct ambient air during flight to the fuel release piston.

Abstract: A rocket engine (10) generates a flow of hot propulsion fluid through a nozzle (14N) for generating propulsive thrust along a thrust vector 8. Hybrid exhaust gas generators (36,38) have their exhausts (44,54) through the side of the nozzle. Each gas generator includes a fuel grain (46,56) and a source of oxidizer (16,40,50). The fuel grain is kept hot by either or both (a) direct radiation or conduction from the hot propulsion fluid, or (b) by a trickle of oxidizer. The fuel grain can thus react quickly to a substantial flow of oxidizer.

Abstract: A hybrid propulsion system comprises a liquid fuel section and a solid fuel section. The liquid fuel section contains an aqueous solution of hydrogen peroxide. An injector system is located between the liquid fuel section and the solid fuel section. The injector system injects a stream of hydrogen peroxide or a decomposed stream of hydrogen peroxide at elevated temperatures into the solid fuel section to effect combustion of the fuel grain in the solid fuel section.

Abstract: A solid fuel propulsion system for a ran jet rocket having a combustion chamber (B) surrounded by a tubular casing (2) and a gas generator (G) disposed upstream from the combustion chamber (B) and surrounded by a tubular casing (1) for generation of a combustible gas from a solid fuel, disposed between the gas generator (G) and the combustion chamber (B) is a gas-stream regulator unit (R) for regulation of the flow of the combustible gas from the gas generator (G) to the combustion chamber (B). The propulsion system has a middle section (4) in which the gas stream regulator unit (R) is housed. The middle section is connected in a load-bearing manner with the combustion chamber casing (2) and the gas generator casing (1) and is provided with a first pressure head (8) sealing the gas generator (G), and a second pressure head (9) sealing the combustion chamber (B). Disposed between the pressure heads (8, 9) is a base unit which houses the gas stream regulator unit (R) and braces the pressure heads (8, 9).

Abstract: A hybrid rocket engine and a method for propelling a rocket utilizing a vortex flow field. The flow field includes an outer fluid vortex spiraling toward a closed end of the flow field generating apparatus and an inner fluid vortex substantially concentric with the outer vortex spiraling away from the closed end and toward an outlet opening in which the inner vortex spirals in the same direction as the outer vortex, but in the opposite axial direction. The invention also relates to a rocket propulsion system utilizing the flow field in which the propulsion system includes a combustion chamber with a fuel source and an oxidizer source that deliver the said fuel and said oxidizer to the said outer and inner vortexes in a manner to support a combustion process while flowing along the flow field.

Abstract: This propulsion system of a rocket motor assembly includes an array of attitude-control rocket engines, one or more oxidizer-fluid sources, one or more ignition-fluid sources, and, optionally, one or more primary rocket engines. Each of the attitude-control rocket engines has a respective combustion chamber and is offset from the longitudinal axis of the rocket motor assembly so that when a selected one or group of the attitude-control rocket engines is fired, the flight path of the assembly is diverted and/or the rocket assembly spins. The oxidizer-fluid and ignition-fluid sources are in operative communication with the attitude-control rocket engines to respectively permit oxidizer fluid and ignition fluid to be supplied to selected ones or groups of the attitude-control rocket engines. Optionally, a portion of the ignition fluid from the ignition-fluid source can be cooled and used to pressurize the oxidizer-fluid source.

Abstract: A turbojet engine includes a housing having a forward inlet nozzle, an aft exhaust nozzle, a combustion chamber therebetween, a compressor between the inlet nozzle and the combustion chamber, and a turbine between the combustion chamber and the exhaust nozzle. A liquid fuel injector is provided to inject atomized or vaporized fuel into the chamber for combustion therein. In addition, a solid fuel is provided in the aft exhaust nozzle. The solid fuel may be used to assist the starting of the engine, particularly useful under adverse starting conditions. That is, when the compressor begins to spin up to pump air through the engine, the solid fuel in the exhaust nozzle is ignited by an igniter which begins to combust with the oxygen rich air introduced by the compressor. This combustion increases the temperature in the entire engine, such that when liquid fuel is injected in the combustion chamber, the temperature is sufficiently high to allow the liquid fuel to burn in a self-sustaining manner.

Abstract: A propulsion system with at least one storage chamber containing at least one solid propellant and at least one fluid propellant is disclosed. The fluid propellant is retained under pressurized conditions, such that depressurization of the storage chamber substantially homogeneously disperses the at least one solid propellant in the at least one fluid propellant. A mixed-phase propellant can thereby be fed to a combustion chamber. The pressurized conditions under which the at least one fluid propellant is retained can include supercritical or critical conditions, saturation conditions, and conditions sufficient to provide a compressed gas.

Abstract: A rocket engine (10) generates a flow of hot propulsion fluid through a nozzle (14N). Hybrid exhaust gas generators (36,38) have their exhausts (44,54) through the side of the nozzle. Each gas generator includes a fuel grain (46,56) and a source of oxidizer (16,40,50). The fuel grain is kept hot by either or both (a) direct radiation or conduction from the hot propulsion fluid, or (b) by a trickle of oxidizer. When the thrust vector is to be modified, the appropriate one of the hybrid gas generators receives a flow of oxidizer, and the resulting exhaust gas is injected through the side of the nozzle.

Abstract: The blanking-cap system (16) includes a glass blanking cap (15) and a destruction device (17) which comprises a projectile (18) capable of destroying the blanking cap (15), and a controllable projection means (19), which is capable of projecting the said projectile (18) and which is arranged outside the conduit (7) while being oriented in such a way as to be able to project the said projectile (18) onto the said blanking cap (15).

Abstract: The invention is an airframe which includes a vehicle (12) having a solid propellant rocket engine (14) and a ramjet or scramjet engine (16); a thrust plug (18) extending from an end (20) of the vehicle which directs combustion gases (23 and 64) produced by the solid propellant rocket engine or ramjet/scramjet engine to produce forward thrust; a longitudinal passage (38) extending from the end of the vehicle to an opening (30) forward of the end which receives external air directed by forward movement of the vehicle and in which solid propellant (32) of the solid propellant rocket engine is located, and wherein during rocket operation solid propellant is combusted to produce the combustion gases in the longitudinal passage which are conveyed by the longitudinal passage into contact with the thrust plug and during ramjet/scramjet operation the longitudinal passage is open to flow of external air after operation of the solid propellant rocket engine is completed and which supports mixing and combustion of the a

Abstract: Batch-mode and continuous-mode decomposition of nitrous oxide is used to provide multiple ignitions of a solid-propellant gas generator and subsequently control its output gas temperature and flow rate, respectively. To reignite the solid-propellant gas generator, a controlled mass of a reactive oxidizer, such as hot nitrous oxide decomposition products, is injected into the gas generator chamber.

Abstract: A family of water-based gas-generating liquid compositions is described. A composition of the present invention includes: hydrogen peroxide; ammonium nitrate; and water. Compositions of the present invention may be mixed with fuels to make monopropellants or used in bipropellant or hybrid systems. Alternative uses of the present invention include breathable gas generation.

Abstract: A monopropellant assisted solid rocket engine of the present invention is comprised of a combination of a solid rocket engine with nitromethane monopropellant injected into the rocket motor. By controlling the rate at which the monopropellant is injected the thrust of the rocket engine can be controlled.

Abstract: The present invention relates to a shut-off system for an orifice for introducing combustion air into a ramjet with a consumable auxiliary motor (14).According to the invention, this shut-off system comprises:an elastic system (28, 29, 31, 32) connected to the shut-off flap (20A, 20B); anda retaining element (33) for keeping the said elastic system in the said tense state during the initial phase of rocket operation, the said retaining element being sensitive to the hot gases emitted by the said consumable auxiliary motor (14) so that on completion of combustion thereof, the said retaining element (33) releases the said elastic system which spontaneously changes from its tense state to its relaxed state, bringing the said flap from its shut-off position to its open position.

Abstract: A cryogenic solid hybrid engine with a solid propellant chamber, a first propellant within such chamber in which the first propellant is in solid form in the chamber and is in fluid form at room temperature, a coolant fluid chamber and a coolant fluid in the coolant fluid chamber being maintained at a temperature blow the freezing point of the first propellant. The invention also relates to a method for propelling a rocket and a method for forming a solid propellant grain for use in a cryogenic solid hybrid rocket engine.