Class / Patent application number | Description | Number of patent applications / Date published |
060205000 | By chemical reaction | 45 |
20090031700 | Mixtures of oxides of nitrogen and oxygen as oxidizers for propulsion, gas generation and power generation applications - This invention involves the mixtures of oxides of nitrogen and oxygen (O | 02-05-2009 |
20100218478 | Turbine engine compressor - A counter-rotating blade stage in lieu of a stator stage may compensate for relatively low rotational speed of a gas turbine engine spool. A first spool may have at least one compressor blade stage and at least one turbine blade stage. A combustor is located between the at least one compressor blade stage and the at least one turbine blade stage along a core flowpath. The at least one counter-rotating compressor blade stage is interspersed with the first spool at least one compressor blade stage. A transmission couples the at least one additional compressor blade stage to the first spool for counter-rotation about the engine axis. | 09-02-2010 |
20100326043 | Expander cycle rocket engine nozzle - An expander cycle rocket engine includes a primary nozzle with a heat exchanger formed therein to cool the nozzle and heat up a fluid used to drive a turbo-pump, and a secondary heat exchanger is located within the primary nozzle and includes passages to channel the fluid in order to add additional heat to the fluid used to drive the turbo-pump. The secondary heat exchanger can be a nozzle shaped heat exchanger located within the primary nozzle, and struts that secure the nozzle shaped heat exchanger within the primary nozzle and channel the fluid between nozzles. The concentric arrangement of first and second heat exchangers can transfer more heat from the combustion gases to the fluid that is used to drive the turbo-pump such that higher pressures can be obtained allowing for larger nozzles and much higher thrust than can be obtained with traditional nozzle engines, or provide significantly higher chamber pressures for engines in the prior art thrust class. | 12-30-2010 |
20120036829 | System, Method and Apparatus for Lean Combustion with Plasma from an Electrical Arc - The present invention provides a plasma arc torch that can be used for lean combustion. The plasma arc torch includes a cylindrical vessel, an electrode housing connected to the first end of the cylindrical vessel such that a first electrode is (a) aligned with a longitudinal axis of the cylindrical vessel, and (b) extends into the cylindrical vessel, a linear actuator connected to the first electrode to adjust a position of the first electrode, a hollow electrode nozzle connected to the second end of the cylindrical vessel such that the center line of the hollow electrode nozzle is aligned with the longitudinal axis of the cylindrical vessel, and wherein the tangential inlet and the tangential outlet create a vortex within the cylindrical vessel, and the first electrode and the hollow electrode nozzle create a plasma that discharges through the hollow electrode nozzle. | 02-16-2012 |
20120079803 | Exhaust plume heat effect reducing method and apparatus - An apparatus for reducing heating effects of an exhaust plume of a jet engine on an impinged surface includes fluid injectors disposed adjacent and aimed into an exhaust plume zone that's to be occupied by an exhaust plume when the engine is running. A flow generator transmits fluid flow into such an exhaust plume through the injectors. Each injector emits fluid in at least two divergent directions to increase the cross-sectional area of the exhaust plume by forming fluidic lobes in the exhaust plume. | 04-05-2012 |
20130145742 | CATALYTICALLY ENHANCED GAS GENERATOR SYSTEM FOR ROCKET APPLICATIONS - A rocket engine system with a fuel conversion system in communication with a gas generator. | 06-13-2013 |
20130186059 | DUAL FUEL AIRCRAFT SYSTEM AND METHOD FOR OPERATING SAME - A dual fuel propulsion system comprising a gas turbine engine configured to generate a propulsive thrust using a cryogenic liquid fuel. | 07-25-2013 |
20140137539 | THRUST-PRODUCING DEVICE WITH DETONATION MOTOR - A detonation thrust-producing device includes an explosive located in a recess in an external surface of a body. Detonation of the explosive expels material out of the recess, providing thrust to the body in an opposite direction. A mass, such as a metal disk, may be placed blocking or covering the external opening. The body may be a part of a vehicle, such as an airborne projectile. The thrust-producing device may include multiple detonation motors arrayed around the body, capable of being individually or multiply detonated. Such thrust-producing devices may be used for attitude adjustment, steering, or other control of the flight of the projectile or other air vehicle. The detonation thrust-producing devices have the advantage of a faster-response time than propellant-based devices, and do not need the nozzles that are used with many propellant-based devices. | 05-22-2014 |
20140305098 | HYBRID-CYCLE LIQUID PROPELLANT ROCKET ENGINE - Systems and methods are described herein for a hybrid liquid propellant rocket engine. In an embodiment, the engine includes a first pump powered by a first turbine, a second pump powered by a second turbine, and a gas generator. An output of the gas generator is connected to the first turbine and the second turbine. The engine further includes a third pump powered by a third turbine, a fourth pump powered by a fourth turbine, and a nozzle having an expander cycle in a wall and a combustion chamber. An output of the third pump is connected to the expander cycle and an output of the wall is connected to the third turbine and the fourth turbine. An output of the fourth pump, an output of the third turbine, and an output of the fourth turbine are connected to the combustion chamber. | 10-16-2014 |
20150013305 | Dual-Mode Combustor - A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated. | 01-15-2015 |
20150354452 | STARTER DEVICE FOR ROCKET MOTOR TURBOPUMP - The subject of the invention is a device for starting a turbopump of a rocket motor of an aircraft including a turbine engine for propelling the aircraft and a rocket motor, which includes a pneumatic supply of compressed air to a turbine of the turbopump, this compressed air being tapped from a tapping on a compressor stage of the aircraft propulsion turbine engine upstream of the combustion chamber of the turbine engine. It applies notably to an aircraft of the space airplane type. | 12-10-2015 |
060206000 | Utilizing indirect heat exchange | 3 |
20100326044 | METHOD FOR COOLING ROCKET ENGINES - A coolant system for a rocket propulsion system having a rocket engine and a propellant tank includes a first heat exchanger that is operatively coupled to the rocket engine. The cooling system also includes a second heat exchanger that is operatively coupled to the propellant tank, a coolant tank that is configured to hold a coolant, and a pump. The cooling system also includes a flow circuit, through which the coolant flows with the aid of the pump. The flow circuit is defined by the pump, the first heat exchanger, the second heat exchanger and the coolant tank. The coolant is pumped by the pump from the coolant tank to the first heat exchanger. The heat of combustion from the rocket engine is transferred to the coolant to cool the rocket engine. The heated coolant transfers the heat therein to the propellant with the second heat exchanger, thereby causing a pressure of the propellant in the propellant tank to be maintained at a desired level. | 12-30-2010 |
20110005193 | Method and apparatus for simplified thrust chamber configurations - The invention of this disclosure is methods and apparatuses improving the ease of fabrication and delivered specific impulse performance of simplified rocket engine thrust chambers. Included are a method and apparatus for a pool-boiling cooling system rocket thrust chamber. This cooling system utilizes a convective coolant flowing in a continuous or semi-continuous coolant loop. In addition the convective coolant itself is cooled in a pool-boiling heat exchanger by the evaporation of a propellant that functions as a boiling coolant. The invention also includes a method and apparatus for a shortened, simplified, conical expansion nozzle for a rocket thrust chamber that can operate with reduced specific impulse losses due to nozzle configuration and the use of film coolant in the thrust chamber. | 01-13-2011 |
20120060464 | SYSTEMS, METHODS AND APPARATUS FOR PROPULSION - In some implementations a propulsion system includes a thrust chamber comprised of a combustion chamber and an expansion nozzle. The thrust chamber has an interior and exterior surfaces and a main propellant injector mounted to the thrust chamber to inject an oxidizer and a fuel into the interior of the thrust chamber. The total fluid flowing to the rocket engine is compromised of oxidizer, fuel, internal film coolant, and external convective coolant. The internal film coolant ranges from about 1% to about 10% of the total fluid. Reduced coolant tubing circumscribes the exterior of the thrust chamber to circulate an external convective coolant, and a nozzle film coolant manifold mounted to the expansion nozzle injects the external convective coolant onto the interior wall of the expansion nozzle, the external convective coolant being about 1% to about 10% of the total fluid flow to the thrust chamber. | 03-15-2012 |
060207000 | Utilizing plural reaction zones within a system | 3 |
20150047316 | MULTI-PULSE GAS GENERATOR AND OPERATION METHOD THEREOF - A multi-pulse gas generator includes a pressure vessel, an outer propellant arranged in the pressure vessel and which has a tubular shape, an intermediate propellant arranged inside the outer propellant and which has a tubular shape, an inner propellant arranged inside the intermediate propellant and which has a tubular shape, an internal structure arranged inside the inner propellant and fixed to the pressure vessel, a first barrier membrane arranged between the outer propellant and the intermediate propellant so as to isolate the outer propellant and the intermediate propellant from each other, and a second barrier membrane arranged between the intermediate propellant and the inner propellant so as to isolate the intermediate propellant and the inner propellant from each other. The outer propellant is supported on its outer surface by the pressure vessel. The inner propellant is supported on its inner surface by the internal structure. | 02-19-2015 |
20160090330 | OXIDIZER-RICH LIQUID MONOPROPELLANTS FOR A DUAL MODE CHEMICAL ROCKET ENGINE - The subject invention relates to oxidizer-rich liquid monopropellants based on ADN or HAN for a dual mode bipropellant chemical rocket engine. Such engines may be part of propulsion systems to be used in aerospace applications for 1) orbit raising, orbit manoeuvres and maintenance, attitude control and deorbiting of spacecraft, and/or 2) propellant settling, attitude and roll control of missiles, launchers and space planes. | 03-31-2016 |
20160377295 | GAS TURBINE ENGINE WITH SELECTIVE FLOW PATH - A gas turbine engine includes a gas generator with at least one compressor rotor, at least one gas generator turbine rotor and a combustion section. There is a duct located downstream of the gas generator. The duct is configured to move between a first position and a second position. A fan drive turbine is positioned downstream of a path of the products of combustion and passes over at least one gas generator turbine rotor when the duct is in the first position. The fan drive turbine is for driving a shaft and the shaft is for engaging gears to drive at least two fan rotors. An augmentor section is positioned downstream of a path of products of combustion and passes over at least one gas generator turbine rotor when the duct is in second position. | 12-29-2016 |
060208000 | Injecting air into the reaction zone | 2 |
20150369134 | Reverse Core Gas Turbine Engine with High Temperature Third Stream - A gas turbine engine has a fan rotor for delivering air axially downstream into a core engine duct, which sequentially passes a turbine, a combustor, and a compressor. The core engine duct extends to a turning supply duct configured to turn the core air flow axially upstream so that the core air sequentially passes through the compressor, combustor and turbine, and into an exhaust conduit which turns the core airflow radially outwardly and axially downstream into an exhaust duct. A door is selectively opened to communicate a portion of the core airflow in the core engine duct to an augmentor with the exhaust duct isolated from the augmentor. The door is at a location prior to the core airflow reaching the compressor. A method is also disclosed. | 12-24-2015 |
20160053722 | OPTIMAL FEEDBACK HEAT ENERGY INTERNAL COMBUSTION ENGINE AND APPLICATIONS - An internal combustion engine wherein a thermo potential heat flow in combustion is maximized by providing a feedback of an optimized amount of thermo potential heat flow that is modulated in the exhaust media, into the air intake, and a method of providing feedback comprises producing a shock wave of pulse of exhaust media and pulse of intake air on the opposite side of a high temperature shock tube thereby transferring the thermo potential heat energy flow from the exhaust media to the air intake. | 02-25-2016 |
060211000 | Injecting separate streams of fuel and oxidizer (e.g., hypergole, etc.) into the reaction zone | 9 |
20100043391 | INJECTOR ASSEMBLY HAVING MULTIPLE MANIFOLDS FOR PROPELLANT DELIVERY - There is provided an injector assembly having two or more oxidizer manifolds and/or two or more fuel manifolds for delivery of liquid propellants to a combustion chamber such that combustion instability is reduced or eliminated during throttling. Delivery of the oxidizer to the oxidizer manifolds is controlled by an oxidizer valve, which may comprise an integral valve. The oxidizer passes from the oxidizer manifolds into the oxidizer element and then into the combustion chamber. The multiple oxidizer manifolds allow the oxidizer to be provided through selective openings of the oxidizer element thus reducing the change in pressure drop across the oxidizer element to thereby reduce or eliminate combustion instability and other problems. Additionally, the injector assembly may also include a lift-off seal or a filler fluid source to fill any temporarily unused oxidizer manifolds with an oxidizer or filler fluid. | 02-25-2010 |
20100257840 | INJECTION DEVICE FOR COMBUSTION CHAMBERS OF LIQUID-FUELED ROCKET ENGINES - An injection device including at least one injection plate adjacent a combustion space of a combustion chamber, and at least one first injection nozzle including a first entry bore having a first discharge into the combustion chamber, and a first orifice bore, having a cross-sectional dimension less than or equal to the first entry bore, coaxially arranged with the first entry bore and remote from the first discharge. At least one second injection nozzle includes a second entry bore having a second discharge into the combustion chamber, and a second orifice bore, having a cross-sectional dimension less than or equal to the second entry bore, coaxially arranged with the second entry bore and remote from the second discharge. The instant abstract is neither intended to define the invention disclosed in this specification nor intended to limit the scope of the invention in any way. | 10-14-2010 |
20120067023 | ROCKET ENGINE AND METHOD FOR CONTROLLING COMBUSTION IN THE ROCKET ENGINE ITSELF - Supply of a liquid component in a combustion chamber of a rocket engine is controlled by a feed valve provided with an obturator mobile between a pen position and a closed position of at least one supply pipe, which has an inlet that communicates with a tank for containing the liquid component and an outlet that communicates with the combustion chamber; the displacement of the obturator from its closed position to its open position being triggered by a pressurized fluid supplied to the outlet of the supply pipe. | 03-22-2012 |
20130199155 | Rocket Propulsion Systems, and Related Methods - In an aspect of the invention, a system for rocket propulsion includes a heater operable to generate thermal energy from energy supplied from a non-chemical energy source, and to supply the thermal energy to a non-cryogenic fuel to thermally decompose the fuel into components that include at least a first component and a second component. The rocket propulsion system also includes a combustion chamber and a nozzle. The combustion chamber is operable to receive an oxidizer and at least a portion of the thermally decomposed fuel, and allow the two to combust. The nozzle generates thrust by directing the products of the combustion out of the system. | 08-08-2013 |
20140290212 | REACTANTS SPRAYED INTO PLASMA FLOW FOR ROCKET PROPULSION - Specific impulse and rocket engine efficiency can be improved by injecting reactants, e.g., a propellant combination or a monopropellant and a catalyst, into a plasma flow of a rocket engine. In some aspects, a catalyst or a propellant is carried by plasma formed by passing a flow of a feed gas through an electrical arc. In some aspects, reactants are combusted in supersonic plasma flow to generate combustion ionization in the plasma flow. | 10-02-2014 |
20160032867 | STABLE HYBRID ROCKET TECHNOLOGY - A hybrid rocket engine is described that achieves stable, highly efficient hybrid combustion by having a core flow of fuel-rich gas generator gases, with the flow being surrounded with an annular injection of oxidizer. The fuel-rich gas serves to vaporize and decompose the oxidizer, such as nitrous oxide, and prepare it for effective, stable combustion. In one embodiment, this is done at the head-end of a combustion chamber. The combustion products can then be expanded through a nozzle to create thrust. The engine can be an upper stage engine that can include modular thrust chambers and an integrated aerospike nozzle. The thrust chambers can be arranged in an array that rings the top of the aerospike nozzle. | 02-04-2016 |
060212000 | Using igniter aid | 2 |
060213000 | Injected separately | 2 |
20130205750 | IGNITER FOR A ROCKET ENGINE, METHOD FOR IGNITION OF A ROCKET ENGINE - An igniter for a rocket engine or motor comprising a combustion chamber with a solid fuel, an inlet for supplying an oxidizer to the combustion chamber to ignite the solid fuel and an outlet for discharging exhaust gas, wherein the igniter is arranged to discharge exhaust gas to the rocket engine for igniting the rocket engine. The igniter can be used for multiple ignitions and can also be re-used after re-filling. | 08-15-2013 |
20130205751 | Fast Ignition and Sustained Combustion of Ionic Liquids - A catalyst free method of igniting an ionic liquid is provided. The method can include mixing a liquid hypergol with a HAN-based ionic liquid to ignite the HAN-based ionic liquid in the absence of a catalyst. The HAN-based ionic liquid and the liquid hypergol can be injected into a combustion chamber. The HAN-based ionic liquid and the liquid hypergol can impinge upon a stagnation plate positioned at top portion of the combustion chamber. | 08-15-2013 |
060214000 | Oxidizer in the form of a mixture | 1 |
20110167789 | RAYLEIGH-TAYLOR ASSISTED COMBUSTION AND COMBUSTORS ADAPTED TO EXPLOIT RAYLEIGH-TAYLOR INSTABILITY FOR INCREASING COMBUSTION RATES THEREIN - A method for improving interpenetration, mixing, and combustion between a fuel and oxidizer reactant mixture and combustion product gases in engines having a combustor includes the step of providing an engine having a combustor in which the reaction mixture and product gases are subjected to acceleration directed transverse to a direction along which the reactant mixture flows through the combustor during combustion. One or more catalyst elements are positioned within the combustor to generate Rayleigh-Taylor instability and thereby enhance interpenetration of the reactant mixture and product gases within the combustor chamber during combustion. | 07-14-2011 |
060217000 | Injecting mixture of fuel and oxidizer into the reaction zone | 4 |
20120279197 | NITROUS OXIDE FLAME BARRIER - Propellants flow through specialized mechanical hardware that is designed for effective and safe ignition and sustained combustion of the propellants. By integrating a micro-fluidic porous media element between a propellant feed source and the combustion chamber, an effective and reliable propellant injector head may be implemented that is capable of withstanding transient combustion and detonation waves that commonly occur during an ignition event. The micro-fluidic porous media element is of specified porosity or porosity gradient selected to be appropriate for a given propellant. Additionally the propellant injector head design integrates a spark ignition mechanism that withstands extremely hot running conditions without noticeable spark mechanism degradation. | 11-08-2012 |
20150354503 | SYSTEM AND A METHOD FOR FEEDING A ROCKET ENGINE - The invention relates to the field of rocket engines, and more particularly to a system for feeding a rocket engine ( | 12-10-2015 |
20150354504 | ROCKET ENGINE, ROCKET AND START METHOD OF ROCKET ENGINE - A rocket engine has fuel passage of pipes | 12-10-2015 |
20160160801 | PREMIXED LIQUID PROPELLANT PROPULSION SYSTEM AND METHOD WITH ANTI-FLASHBACK QUENCHING LIQUID INJECTOR - A liquid injector system for a combustion engine, having a single feed inlet configured to receive a premixed liquid propellant under pressure or a purge gas under pressure, and having a liquid injector assembly. The assembly has a liquid injector having a hollow dome and injector holes configured to receive and inject the premixed liquid propellant or the purge gas through the liquid injector and into a combustion chamber. The liquid injector system has a liquid-to-gas zone between an injector outlet side and a flame front. A pressure gradient decrease between the liquid injector and the combustion chamber causes the premixed liquid propellant to expand from liquid to gas phases, which causes a temperature decrease at the liquid-to-gas zone, wherein the pressure gradient decrease and the temperature decrease prevent or mitigate the flame front from propagating upstream of the combustion chamber, which achieves an anti-flashback quenching liquid injector design. | 06-09-2016 |
060218000 | Decomposing a compound in the reaction zone | 5 |
20110056184 | EXTENDED ALTITUDE COMBUSTION SYSTEM - A combustion system for performing stable combustion and flame stabilization at high altitudes is described. A primary liquid hydrocarbon fuel is atomized and vaporized within the main combustor chamber to produce a primary fuel vapor. When the combustion system operates at a high altitude, a secondary gaseous fuel is fed into the inlet air port such that the secondary fuel mixes with air, thereby enabling the mixture of the air and the secondary fuel to combust in a catalytic reactor to produce high temperature, oxygen-rich gases that flow into the main combustor chamber. Proper proportional amounts of the two fuels are determined as a function of altitude. | 03-10-2011 |
20130014487 | PACKAGED PROPELLANT AIR-INDUCED VARIABLE THRUST ROCKET ENGINE - This invention is a packaged propellant air-induced variable thrust rocket engine that has a vast number of uses and applications for this invention. The primary purpose of the device described here is to provide a light weight, torque and vibration free thrust generator for the propulsion of aircraft. This device will facilitate the fabrication of very light weight aircraft because of the lack these forces. This device can also be used anywhere high velocity air flow and or the resulting thrust is needed. The invention uses aerodynamic principles to compress and accelerate the incoming air, prior to it being heated and accelerated by a short duration burst of thermal and kinetic energy from discrete packets of a mixture of oxidizable fuels. The heated and accelerated air then expands as it travels thru the device providing thrust. | 01-17-2013 |
20130042594 | TERRESTRIAL POWER AND PROPULSION FROM NUCLEAR OR RENEWABLE METAL FUELS WITH MAGNETOHYDRODYNAMICS - A propulsion system is disclosed that uses metal fuel particles heated to a range from 3000° K. to 6000° K. by reaction with oxygen in air inside a cyclone combustor to form metal oxides that are retained and removed from the combustor for re-conversion to metal, while nitrogen in the air is heated to the temperatures that on supersonic exhaust propels 50,000 ton monohull ships to from 50 to 100 knots. The ship can accelerate by electromagnetic force from a MHD generator-accelerator rendered electrically conducting by alkali metal seed injection into the gas. Also disclosed are 10 MW to 1000 MW Closed Cycle MHD power plants fired by natural gas into a top half of a falling pebble bed heat exchanger that transfers 2000° K. to 3000° K. heat to a noble or diatomic gas in a bottom half of the exchanger that on exit is seeded with an alkali metal rendering the gas conducting. | 02-21-2013 |
20130305685 | Novel Ionic Micropropellants Based on N2O for Space Propulsion - Novel monopropellants are provided. The monopropellants are based on N | 11-21-2013 |
20150121843 | REACTOR FOR AMMONIUM DINITRAMIDE-BASED LIQUID MONO-PROPELLANTS, AND THRUSTER INCLUDING THE REACTOR - The present invention relates to a reactor for the decomposition of ammonium dinitramide-based liquid monopropellants into hot, combustible gases for combustion in a combustion chamber, and a rocket engine or thruster comprising such reactor, wherein the reactor comprises a heat bed exhibiting catalytic activity. | 05-07-2015 |
060219000 | Using solid material in reaction zone | 8 |
20090056305 | SCALABLE FLAT-PANEL NANO-PARTICLE MEMS/NEMS THRUSTER - A scalable flat-panel nano-particle MEMS/NEMS thruster includes a grid having a plurality of electrodes to establish electrical fields. A liquid is disposed in a liquid reservoir of the grid. The liquid is positioned within the electrical fields. A plurality of nano-particles are suspended in the liquid. A plurality of MEMS and NEMS micron-size vias are disposed in the grid. The electrical fields extract the plurality of nano-particles from the liquid and accelerate the nano-particles in the vias to provide propulsion system thrust. | 03-05-2009 |
20090205311 | Combined cycle missile engine system - 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. | 08-20-2009 |
20090320443 | Propulsion system, opposing grains rocket engine, and method for controlling the burn rate of solid propellant grains - A solid propellant thrust control system, method, and apparatus for controlling combustion of solid propellants in an opposing grains solid propellant rocket engine (OGRE) is provided. In particular, an opposing grains rocket engine and propulsion system is provided, in which actuator means connected are connected to solid propellant grains disposed in the pressure vessel of the engine. The actuator means are operable to selectively move the solid propellant grains together or apart relative to one another, such that the burning ends of the solid propellant grains decrease or increase relative to one another. This action controls the rate of combustion of the solid propellant grains by varying spacing distance between the burning ends of the solid propellant grains, and enables extinguishment and reignition of the OGRE. Further, a method is providing for controlling the burn rate of solid propellant grains undergoing combustion in an opposing grains rocket engine. | 12-31-2009 |
20100077723 | MOTOR WITH NOTCHED ANNULAR FUEL - A motor includes an annular solid fuel between an inside diameter and an outside diameter. The solid fuel has a series of radial notches that define segments of fuel between them. The notches allow for faster burning of the fuel, while still allowing structural integrity of the fuel segments to be retained during the burning process. | 04-01-2010 |
20120304620 | CATALYST, GAS GENERATOR, AND THRUSTER WITH IMPROVED THERMAL CAPABILITY AND CORROSION RESISTANCE - A catalyst includes a carrier of essentially hafnia, up to an equal part zirconia, and optionally additional stabilizers, upon the surface of which is deposited an active metal suitable to promote the reaction of propellants to be used in gas generators and thrusters. | 12-06-2012 |
20130097995 | SYSTEM AND METHOD FOR CONTROLLING AN OBJECT TRAVELING THROUGH EXOATMOSPHERIC SPACE - A system for controlling an object traveling through exoatmospheric space is disclosed herein. The system includes, but is not limited to, a pressure vessel, a pair of solid propellant grains associated with the pressure vessel that is configured to direct a gas into the pressure vessel during combustion, an exhaust nozzle that is in fluid communication with the pressure vessel, and a valve that is coupled with the exhaust nozzle and that is configured to selectively obstruct the gas from venting through the exhaust nozzle. | 04-25-2013 |
20150059314 | ELECTRICALLY IGNITED AND THROTTLED PYROELECTRIC PROPELLANT ROCKET ENGINE - According to one aspect, an apparatus and method for electrically igniting and throttling pyroelectric propellant, e.g., in a rocket engine, are provided. In one example, an apparatus includes an injector body for supplying an electrically ignitable propellant to a combustion chamber and a opposing electrodes. A first electrode may be included with the injector body and a second electrode positioned relative to the first electrode to cause ignition of the electrically ignitable propellant as the electrically ignitable propellant flows thereby. | 03-05-2015 |
060220000 | Including injecting modifying fluid | 1 |
20090007541 | THRUSTER USING NITROUS OXIDE - A thruster is provided that makes it possible to reduce and eventually eliminate the toxicity of storable liquid propellant and improve the low-temperature environment adaptability of propulsion system using the storable liquid propellant. The thruster produces thrust by using catalytic decomposition gas obtained by catalytically decomposing nitrous oxide with a nitrous oxide decomposition catalyst. | 01-08-2009 |