| Class / Patent application number | Description | Number of patent applications / Date published |
|---|
| 060257000 | Liquid oxidizer | 33 |
| 20090120060 | Catalytically activated transient decomposition propulsion system - A catalytically activated transient decomposition propulsion system provides thrust by decomposing flow controlled propellant in contact with a catalyzing agent using a fixed volume of liquid propellant that is placed in contact with the catalyst within the decomposition chamber by a calibrated flow control valve. After injecting the liquid propellant into the decomposition chamber, the valve returns to the closed position while surface tension holds the liquid within the decomposition chamber until complete decomposition and exhaust of the warm gaseous products through a converging and diverging nozzle occurs. The increasing and decreasing transient pressure in the decomposition chamber changes each cycle in response to flow control valve actuation as the decomposition process is repeated. | 05-14-2009 |
| 20100037589 | COMPONENT CONFIGURED FOR BEING SUBJECTED TO HIGH THERMAL LOAD DURING OPERATION - A component configured for being subjected to a high thermal load during operation includes a wall structure with cooling channels adapted for handling a coolant flow. At least one first cooling channel is adapted to convey the coolant from a first portion of the component to a second portion of the component. At least one second cooling channel in the second portion is closed so that the coolant is at least substantially-prevented from entering the closed second cooling channel from a cooling channel in the first portion. | 02-18-2010 |
| 20100107601 | ELECTROLYTIC IGNITER FOR ROCKET ENGINES USING MONOPROPELLANTS - The electrolytic ignitor comprises an injector ( | 05-06-2010 |
| 20100170223 | CONTROL AND/OR DRIVE DEVICE FOR A FLYING BODY - Control and/or drive device for a flying body for ejecting hot gas streams of a combusted fuel combination of at least a first and second component. Device includes a first hollow chamber body structured and arranged to contain first component, a second hollow chamber body structured and arranged to contain second component, a controllable fuel valve arranged between first hollow chamber body and second hollow chamber body to control feed of first component to second hollow chamber body, and a plurality of outlets structured and arranged to eject respective hot gas streams for influencing a flight path of flying body. Second hollow chamber body is formed as a combustion chamber for combusting the at least first and second components within second hollow chamber body to generate respective hot gas streams, and plurality of outlets are connected to the second hollow chamber body. | 07-08-2010 |
| 20110005195 | ALUMINUM POROUS MEDIA - Disclosed are materials of variable density or tiered porosity micro-fluidic porous media structures of sintered metal or other materials, and methods of making same. An embodiment discloses an aluminum porous media element of variable density having a tiered porosity micro-fluidic media structure. A method of making the aluminum porous media element disclosed herein includes mixing a binding agent with a metal powder to generate a first mixture, heating the first mixture to a sub metal sintering temperature to get a homogeneous composite of the metal powder and heating the homogeneous composite to a metal sintering temperature to sinter-bond the metal powder to get a porous media of first porosity. | 01-13-2011 |
| 20120131903 | ROCKET ENGINE WITH CRYOGENIC PROPELLANTS - A cryogenic-propellant rocket engine includes: at least a first tank for a first liquid propellant; a second tank for a second liquid propellant; a third tank for an inert fluid; an axisymmetrical nozzle including a combustion chamber, a device for injecting first and second liquid propellants into the combustion chamber, a nozzle throat, and a divergent section; and a heater device including at least one duct for conveying the inert fluid and arranged outside the nozzle in immediate proximity thereof, but without making contact therewith, to recover energy of thermal radiation emitted when the rocket engine is in operation and to heat the inert fluid. | 05-31-2012 |
| 20120144799 | Oxidizer Compound for Rocket Propulsion - A rocket propulsion oxidizer compound that is a mixture that is a homogenous and stable liquid at room temperature that includes nitrous oxide and nitrogen tetroxide. | 06-14-2012 |
| 20130205754 | OXIDIZER COMPOUND FOR ROCKET PROPULSION - The present disclosure generally pertains to a rocket propulsion oxidizer compound that is a solution, is a homogenous and stable liquid at room temperature and includes nitrous oxide and nitrogen tetroxide. In addition, an apparatus is provided for burning a fuel and nitrous oxide/nitrogen tetroxide. The apparatus has a combustor, a catalyst, a nitrous oxide/nitrogen tetroxide supply passage for directing the nitrous oxide/nitrogen tetroxide to a contact position with the catalyst, and a fuel supply passage for supplying the fuel to the combustor. The catalyst acts to facilitate decomposition of the nitrous oxide/nitrogen tetroxide, while the combustor burns the fuel, the decomposed nitrous oxide/nitrogen tetroxide and/or nitrous oxide/nitrogen tetroxide decomposed in the reaction. | 08-15-2013 |
| 20140260186 | ROCKET ENGINE SYSTEMS WITH AN INDEPENDENTLY REGULATED COOLING SYSTEM - An improved rocket engine system including a rocket engine, a coolant source, and a propellant source. Each of coolant source and propellant source is in operative communication with rocket engine whereby there is independent regulation or control of coolant source flow relative to propellant source flow. The improved rocket engine system may further include one or more of: at least one power source, at least one power source motor, at least one pump, at least one controller, and a propellant pressurizing source. The propellant source preferably includes a fuel source and an oxidizer source. An embodiment includes a cooling system with a coolant source, a fuel system with a fuel source, an oxidizer system with an oxidizer source, a propellant pressurizing system with a propellant pressurizing source, multiple pumps, an assembly of power source and controller, and a main rocket engine assembly. | 09-18-2014 |
| 060258000 | Including injector means | 16 |
| 20080256925 | Compact, high performance swirl combustion rocket engine - A compact rocket engine with advanced swirl combustion can generate vacuum thrust in the 500 lb | 10-23-2008 |
| 20090007543 | Pintle injector tip with active cooling - A bi-propellant rocket engine may include a primary propellant flowing in a central passageway, a secondary propellant flowing in a secondary passageway generally coaxial with central passageway and a pintle tip having a central chamber sidewall coaxial with the primary passageway and surrounding a central chamber, the central chamber sidewall having a first plurality of apertures there through so that some of the primary propellant exits the central chamber transverse to the flow of the secondary propellant in the secondary passageway. The pintle tip may have a secondary chamber sidewall, substantially thicker than the primary chamber sidewall, surrounding a secondary chamber downstream of and in fluid communication with the primary chamber, the secondary chamber sidewall having a second plurality of apertures there through so that some of the primary propellant exits the secondary chamber transverse to the flow of the secondary propellant in the secondary passageway. The pintle tip may have an end wall generally traverse to the flow of the primary propellant in the central passageway and adjacent the secondary chamber sidewall so that the flow of primary propellant through the secondary chamber sidewall cools a downstream face of the end wall during combustion of the mixed propellants adjacent the downstream face of the end wall. The pintle tip may be used for mixing a first liquid with a second liquid. | 01-08-2009 |
| 20090211228 | HIGH PERFORMANCE LIQUID FUEL COMBUSTION GAS GENERATOR - A gas generation system includes a fuel source, an oxidizer source, and a combustion chamber. The fuel source is operable to supply a flow of a lithium fuel, and the oxidizer source is operable to supply a flow of a fluorinated carbon oxidizer. The combustion chamber is coupled to receive the flow of lithium fuel and the flow of the fluorinated carbon oxidizer and, upon receipt thereof, supplies a combustion gas. The combustion chamber is formed, at least partially, of a carbon material. | 08-27-2009 |
| 20090241511 | HEAT EXCHANGE INJECTOR FOR USE IN A ROCKET ENGINE - A heat exchange injector assembly includes a heat exchange element comprising a fuel sleeve, a liquid oxidizer post disposed in the fuel sleeve, and a multi-passage swirl member such as a double helix member, disposed in the liquid oxidizer post. | 10-01-2009 |
| 20090320447 | Coaxial ignition assembly - A bi-propellant injector includes first and second injector elements and a spark exciter assembly. The first injector element has a conductive layer electrically connected to the spark exciter assembly and a nonconductive layer disposed on an exterior portion of the conductive layer. The second injector element comprises a conductive material and has an opening therethrough in fluid communication with a combustion chamber. An end of the first injector element is positioned at or near the opening in the second injector element. The spark exciter assembly can generate an electrical arc between the conductive layer of the first injector element and the second injector element. | 12-31-2009 |
| 20100005779 | Device for injecting a mono-propellant with a large amount of flow rate modulation - The device for injecting a liquid mono-propellant with a large amount of modulation of its flow rate and disposed at an upstream end of the wall of a combustion chamber of a rocket engine has a feed channel for feeding a mono-propellant from a tank. The device includes a single annular speed-up channel connected to the feed channel and having its outlet opening out via an annular injection section, the speed-up channel and the annular injection section being defined firstly by a first wall forming a stationary surface of revolution situated level with said upstream end, and secondly by a second wall forming a surface of revolution that is on a part that is movable in translation relative to the first wall forming a stationary surface of revolution. | 01-14-2010 |
| 20100037590 | LOW VELOCITY INJECTOR MANIFOLD FOR HYPERGOLIC ROCKET ENGINE - A fuel manifold for a thrust chamber assembly includes a main fuel chamber which is generally frustro-conical in shape. The main fuel chamber provides a resonance frequency that is at least an order of magnitude lower than an acoustic resonance frequency of a combustion chamber. | 02-18-2010 |
| 20100064661 | INJECTOR HEAD OF LIQUID ROCKET ENGINE - An injector head of a liquid rocket engine in which a mixture of oxidizer and fuel is supplied to a combustion chamber, including an injection plate having through-holes, a partition plate mounted above the injection plate so as to define a space between itself and the injection plate, a fuel injector which injects the fuel supplied from a fuel manifold into the combustion chamber, with an upper portion thereof engaged with a lower surface of the partition plate and a lower portion thereof inserted into the through-holes of the injection plate, and an oxidizer injector engaged with an upper surface of the partition plate so as to inject the oxidizer supplied from an oxidizer manifold, wherein the partition plate is provided with an oxidizer supply hole through which the oxidizer injected from the oxidizer injector is delivered towards the fuel injected from the fuel injector. | 03-18-2010 |
| 20100107602 | ELECTROLYTIC IGNITER FOR ROCKET ENGINES USING LIQUID PROPELLANTS - The electrolytic ignitor comprises an injector ( | 05-06-2010 |
| 20100115917 | Method and Device for Supplying a Space Propulsion Engine with Liquid Cryogenic Propellants - According to the invention, upstream of the injection means, the combustible propellant and the oxidant propellant are mixed at constant pressure; and to inject the said mixture of propellants at constant pressure into the combustion chamber ( | 05-13-2010 |
| 20110219743 | INJECTOR ASSEMBLY FOR A ROCKET ENGINE - A cap for a liquid propellant injection assembly of a rocket engine includes a cap body and a valve assembly. The cap body extends between first and second ends. The cap body has a bore that fluidly connects one or more inlets to an outlet. The inlets are disposed in a tubular sidewall of the cap body. The outlet is disposed in the second end of the cap body. The valve assembly includes a valve cap disposed around the first end of the cap body. The valve assembly is adapted to selectively regulate flow of a propellant through the inlets in the cap body as a function of pressure exerted by the propellant against the valve assembly. | 09-15-2011 |
| 20120167552 | ROCKET ENGINE SYSTEM FOR REALIZING HIGH-SPEED RESPONSE - Disclosed is a turbo pump in which a pump impeller is connected to one end of a rotary shaft and a turbine is connected to the other end of the rotary shaft. The turbo pump is designed such that an equivalent region, between a turbine efficiency curve obtained on the basis of a conditional expression where the number of rotations of the rotary shaft is maintained constant regardless of a pump flow rate and a turbine efficiency curve of an actual machine, becomes an operation region. | 07-05-2012 |
| 20130160426 | ROCKET ENGINE INJECTOR ASSEMBLY WITH CRYOGENIC CAVITY INSULATION - An injector assembly for a rocket engine includes a thermal insulating layer adjacent to an oxidizer cavity. | 06-27-2013 |
| 20130239545 | ROCKET ENGINE PRESSURE SENSE LINE - A rocket engine with a manifold is in communication with a combustion chamber. A sense line extends through the propellant manifold and into the combustion chamber. The sense line includes a venturi arranged downstream from the combustion chamber, and at least one aperture fluidly connecting the propellant manifold to a sense-line passageway downstream from the venturi. A method of sensing conditions in a combustion chamber includes exposing an end of a sense line to the combustion chamber, creating a low static pressure in the sense line at a location upstream from the end, introducing a fluid at the location to purge the sense line, and sensing the conditions downstream from the location. | 09-19-2013 |
| 20150027102 | TRI-PROPELLANT ROCKET ENGINE FOR SPACE LAUNCH APPLICATIONS - A tri-propellant rocket engine for space launch applications is disclosed. The tri-propellant rocket engine comprises three main assemblies: an injector, a chamber head, and a chamber. | 01-29-2015 |
| 20160153400 | GAS GENERATOR | 06-02-2016 |
| 060259000 | Including pressurizing means | 5 |
| 20120198818 | AUGMENTATION OF A MONOPROPELLANT PROPULSION SYSTEM - A propulsion system of a spacecraft includes a main tank adapted to contain a volume of propellant and a pressurising gas which applies pressure to the propellant. The main tank includes a membrane delimiting an upper volume to contain the pressurising gas and an inferior volume to contain the propellant. The system further includes an auxiliary tank adapted to contain pressurising gas and connected directly to the main tank by a pressurisation circuit. In operation, gas contained in the auxiliary tank expands continuously with the gas contained in the upper volume of the main tank, the pressures prevailing in the upper volume of the main tank, in the pressurisation circuit and in the auxiliary tank being identical. The auxiliary tank is dimensioned so that the maximal volume of propellant of the main tank is greater than the volume of propellant which the main tank can contain without an auxiliary tank. | 08-09-2012 |
| 20140245717 | DEVICE FOR HEATING A FLUID - A device heating a fluid and usable in a rocket launcher to pressurize a liquefied propellant. The device includes a first burner performing first combustion between a limiting propellant and an excess propellant; a first heat exchanger in which first burnt gas from the first combustion transfers heat to the fluid; at least one second burner into which both the first burnt gas and some limiting propellant are injected to perform second combustion between the limiting propellant and at least a portion of unburnt excess propellant present in the first burnt gas. The second burnt gas from the second combustion flows through a second heat exchanger to transfer heat to the fluid. Burnt gas from each combustion flows in respective burnt gas tubes within a common overall heat exchanger including the heat exchange units, the gas transferring heat to the fluid, the fluid flowing between the burnt gas tubes. | 09-04-2014 |
| 20150143797 | TURBOPUMP - A feed method for feeding reaction engines including off-loading a secondary flow of a first propellant downstream from a first pump but upstream from a first turbine that is driven by expansion of the first propellant and that drives at least the first pump. The off-loading is controlled in such a manner as to achieve equilibrium between power generated by the first turbine and power consumed by the first pump, thereby stopping a rise in speed of the first turbine and the first pump at a predetermined speed lower than a nominal speed. | 05-28-2015 |
| 20160016677 | SATELLITE PROPULSION DEVICE ALLOWING PASSIVE ELIMINATION OF PRESSURIZING GAS - A satellite propulsion system comprises: at least one tank containing a propulsion reagent, and at least one tank containing a pressurizing gas that pressurizes the propulsion reagent, at least one transport means for transporting the pressurizing gas from the pressurizing gas tank to the propulsion reagent tank, the transport means comprising at least one opening allowing a continuous leak-off of pressurizing gas, and at least one device intended, after the satellite has entered operational orbit, to isolate a zone comprising the opening from the rest of the propulsion system. | 01-21-2016 |
| 20160195039 | DEVICE FOR FEEDING A ROCKET ENGINE WITH PROPELLANT | 07-07-2016 |
| 060260000 | Including heating means | 3 |
| 20100218482 | SYSTEM AND METHOD FOR COOLING ROCKET ENGINES - A propulsion system for a rocket engine and a method of cooling a rocket engine includes a propellant tank fluidically coupled to the rocket engine to hold a pressurized propellant, a coolant tank to hold a coolant, a first heat exchanger thermally coupled to the rocket engine and fluidically coupled to the coolant tank, a second heat exchanger thermally coupled to the propellant tank and fluidically coupled to the first heat exchanger, and a third heat exchanger disposed inside the propellant tank to thermally couple a propellant withdrawn from the tank for combustion to a propellant disposed inside the tank. The coolant flows from the coolant tank to the first heat exchanger and through the first heat exchanger to cool the rocket engine. The propellant withdrawn from the propellant tank receives heat from the propellant disposed inside the tank through the third heat exchanger to convert to a gaseous propellant when withdrawn from the propellant tank as a liquid propellant. The coolant flows from the first heat exchanger to the second heat exchanger and through the second heat exchanger to heat the propellant disposed inside the propellant tank. | 09-02-2010 |
| 20150128563 | 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, which reactor further comprises an inner reactor housing accommodating a heat bed and a catalyst bed, and separating the heat bed and catalyst bed from contact with the inner surface of the reactor housing. | 05-14-2015 |
| 20150308384 | PROPULSION ASSEMBLY FOR ROCKET - A propulsion assembly for a rocket including a tank for liquid oxygen, an engine having a combustion chamber, and a “heater” heat exchanger suitable for vaporizing liquid oxygen. The assembly has a vaporized oxygen circuit suitable for directing the oxygen vaporized by the heater either to the combustion chamber or to the tank. When the vaporized oxygen is directed to the combustion chamber, the engine advantageously develops low thrust. | 10-29-2015 |
