Abstract: An injection device for injecting an oxidizer for a liquid rocket includes a housing, a plate disposed inside the housing and having an injection hole to eject an oxidizer, a duct disposed above the plate to guide the oxidizer, and a manifold with one end connected to the injection hole of the plate and the other end connected to the duct, wherein the oxidizer may be distributed to the injection hole at an equal flow rate.

Abstract: A thrust reverser of a nacelle may include a translating sleeve and an inner fixed structure a pressure relief assembly. A pressure relief mechanism may include a pressure relief door coupled via a hinge and a latch to the inner fixed structure. Alternatively, a pressure relief door may be coupled to a frame via a hinge and a latch to the inner fixed structure. The pressure relief assembly may limit deflections between the thrust reverser and the pylon in response to a burst duct. The pressure relief door may release from the latch automatically in response to an over pressurization event.

Abstract: Provided are a precursor supply unit, a substrate processing system, and a method of fabricating a semiconductor device using the same. The precursor supply unit may include an outer container, an inner container provided in the outer container and used to store a precursor source, a gas injection line having an injection port, which is provided below the inner container and in the outer container and is used to provide a carrier gas into the outer container, and a gas exhaust line having an exhaust port, which is provided below the inner container and in the outer container and is used to exhaust the carrier gas in the outer container and a precursor produced from the precursor source.

Abstract: A turbofan engine may comprise an inlet and a fan case coupled to the inlet. An engine case may be coupled to the fan case via a vane extending between the fan case and the engine case. A strut apparatus may extend from the fan case and limit deflection of the fan case. The strut apparatus may comprise a first end proximate the fan case, and a second end coupled to at least one of the engine case or a structure for mounting the turbofan engine to an aircraft.

Abstract: A pressure relief arrangement for a gas turbine engine comprises a hinged door and a mount spaced from the door. An expandable structure is connected to the door and to the mount such that when the door hinges open the expandable structure expands.

Abstract: Gas turbine engine systems involving I-beam struts are provided. In this regard, a representative strut assembly for a gas turbine engine includes a first I-beam strut having first and second flanges spaced from each other and interconnected by a web, the first strut exhibiting a twist along a length of the web.

Abstract: Electrically conductive aerogel and methods of making the same are disclosed. A solution is provided. The solution is cured to form a polymer. The polymer is carbonized to form the conductive aerogel.

Abstract: A gas turbine engine includes a first annular portion that is stationary and adapted for partially surrounding an engine core. The first annular portion includes a fore pylon connecting portion. The gas turbine engine also includes a rail coupled to the fore pylon and extending in the aft direction from the first annular portion. The gas turbine engine also includes a second annular portion, arranged aft of the first portion and coupled to the rail. The second annular portion is movable along an engine core centerline between a closed position and at least one open position. The second annular portion is configured to engage the first annular portion in the closed position, thereby providing access to the engine core. The gas turbine engine further comprises a thrust reverser arranged in the second annular portion.

Abstract: The present invention relates to an RF system (1) able to generate a voltage for accelerating charged particles in a synchrocyclotron, the RF system (1) including a resonant cavity (2) comprising a conducting enclosure (5) within which are placed a conducting pillar (3) of which a first end is linked to an accelerating electrode (4) able to accelerate the charged particles, a rotary variable capacitor (10) coupled between a second end opposite from the first end of the pillar (3) and the conducting enclosure (5), the said capacitor (10) comprising fixed electrodes (11) and a rotor (13) comprising mobile electrodes (12), the fixed electrodes (11) and the mobile electrodes (12) forming a variable capacitance able to vary a resonant frequency of the resonant cavity (2) in a cyclic manner over time, an exterior layer of the rotor (13) having a conductivity of greater than 20,000,000 S/m at 300 K.

Abstract: An exoatmospheric vehicle uses a control system that includes a thrust system to provide thrust to control flight of the vehicle. A regenerative heat system is used to preheat portions of the thrust system, prior to their use in control of the vehicle. The heat for preheating may be generated by consumption of a fuel of the vehicle, such as a monopropellant fuel. The fuel may be used to power a pump (among other possibilities), to pressurize the fuel for use by thrusters of the thrust system. The preheated portions of the thrust system may include one or more catalytic beds of the thrust system, which may be preheated using exhaust gasses from the pump. The preheating may reduce the response time of the thrusters that have their catalytic beds preheated. Other thrusters of the thrust system may not be preheated at all before operation.

Abstract: A propulsion device including a chamber that stores a superfluid, a substrate coupled to a portion of the chamber, a plurality of orifices extending through the substrate, each of the plurality of orifices having a first end and a second end opposite the first end, the first end disposed in an interior of the chamber and the second end disposed outside the chamber; and a pressure source that generates a pressure differential between the first end of each of the plurality of orifices and the second end of each of the plurality of orifices.

Abstract: An acoustic resonance igniter uses gas expanding through a nozzle to form a sonic, or under-expanded supersonic, jet directed against the opening of a blind resonance cavity in a central body, setting up a high-frequency sonic resonance which heats the gas within the cavity. A pintle extends coaxially with the nozzle and injects liquid propellant into the jet. The liquid propellant ignites with the heated gas within the resonance cavity forming combustion gases. The combustion gases flow through openings in a flange which supports the resonance cavity into a combustion chamber in the same direction as the gas jet flows. The liquid propellant is injected from within the support flange in the direction of combustion gas flow to film cool the combustion chamber wall and the flange and the central body supported by the flange. The acoustic resonance igniter may form a rocket engine ignition torch or a RCS thruster.

Abstract: An engine apparatus is provided for a flying object, the engine apparatus including a combustion chamber having a throat region and a nozzle region, the nozzle region having a nozzle wall, wherein the nozzle region expands from the throat region towards an exit end relative to a combustion chamber axis, wherein the nozzle region has associated therewith a skirt having a skirt wall, the skirt being positioned downstream relative to the exit end and surrounding the exit end of the nozzle region, and wherein the skirt wall is at an acute angle away from the combustion chamber axis with respect to the nozzle wall, at least at the exit end of the nozzle region.

Abstract: An engine propulsion system is configured to utilize bursts of media in order to create mechanical energy. The engine propulsion system includes at least one cannon, wherein each cannon is configured to displace the media and further includes a firing pin casing configured to accommodate a firing pin. The firing pin is configured to create the mechanical energy when moved thus allowing the media to exit the cannon.

Abstract: A vessel, such as a rocket motor tube, comprised of an insert, and a joint that allows the insert to decouple from the vessel to relieve pressure when subjected to excessive temperatures and/or a predetermined temperature or temperature range.

Abstract: This invention relates to devices for discharging high pressure exhaust, and in particular to pulse detonation engines. More specifically, the invention describes a rotary pulse detonation engine having a rotary valve system. The rotary valve includes a generally-triangular rotor having rotor tips within a rotor chamber having trochoid inner end surfaces and side surfaces. The rotor defines three working chambers defined by the rotor tips contacting the rotor surfaces. In operation, the rotor tips move in a circumferential direction around the rotor chamber as the rotor spins. During operation, each of the working chambers will sequentially pass through an intake interval, compression interval, expansion interval, and an exhaust interval to create and detonate compressed fuel air mixtures for effective release to an exhaust chamber and nozzle thereby creating a pulsed detonation sequence.

Abstract: A propellant tank for storing a liquid propellant A and supplying vapor produced by evaporation of part of the liquid propellant A to an external location comprises a tank body for storing the liquid propellant A, a mesh member arranged inside the tank body to cover a liquid surface of the liquid propellant A to divide an interior of the tank body into a liquid propellant storing area LA and a gas storing area GA by utilizing surface tension of the liquid propellant, and a heater arranged to a gas storing area GA side of the tank body to keep the gas storing area GA at higher temperature than temperature in the liquid propellant storing area LA. The tank body has a propellant inlet open into the liquid propellant storing area LA and a gas outlet open into the gas storing area GA.

Abstract: A system and methods are provided for combining systems of an upper stage space launch vehicle for enhancing the operation of the space vehicle. Hydrogen and oxygen already on board as propellant for the upper stage rockets is also used for other upper stage functions to include propellant tank pressurization, attitude control, vehicle settling, and electrical requirements. Specifically, gases from the propellant tanks, instead of being dumped overboard, are used as fuel and oxidizer to power an internal combustion engine that produces mechanical power for driving other elements including a starter/generator for generation of electrical current, mechanical power for fluid pumps, and other uses. The exhaust gas from the internal combustion engine is also used directly in one or more vehicle settling thrusters. Accumulators which store the waste ullage gases are pressurized and provide pressurization control for the propellant tanks.

Abstract: An electrospray thruster and methods of manufacturing such thrusters are provided. The micro-electrospray thruster increases the thrust density of conventional electrospray thrusters by miniaturizing the individual components of the thruster thereby allowing for the increase in the number and density of the charged particle emitters.

Abstract: A propulsion system for a supersonic aircraft includes an engine, a compression surface upstream of the engine, a shroud surrounding the engine configured to direct airflow passing over the compression surface towards the engine, and a plurality of vortex generators positioned upstream of the engine. The vortex generators have a height such that when the supersonic aircraft is flown at a predetermined speed, the plurality of vortex generators create a plurality of vortices that propagate partially outside of a boundary layer formed proximate a surface of a supersonic inlet. The vortices cause a high-velocity portion of the airflow to move towards a portion of the engine having a higher sensitivity to changes in stagnation pressure and a low-velocity portion of the airflow to move away from the portion of the engine having the higher sensitivity to changes in stagnation pressure prior to the airflow reaching a face of the engine.

Abstract: Invention discloses an Impulse Drive useful to propel all types of flying, water or land vehicles or devices, without need of the lift the fluid air or water could provide. Propulsion is accomplished by the resultant force of a rotating impulse drive submerged in a fluid that is adjacent to a pressurized gas chamber, separated by a flexible membrane. Multiple impulse drives can be coupled to eliminate rotational movement components, resulting in unidirectional impulse force. Properly installed in a vehicle, it displaces it in the desired direction. This displacement can be performed equally in vacuum. It can be installed inside or outside the vehicle.

Abstract: A system and method of cooling a rocket motor component includes injecting a high pressure liquid coolant through an injector nozzle into a cooling chamber. The cooling chamber having a pressure lower than the high pressure liquid coolant. The liquid coolant flashes into a saturated liquid-vapor coolant mixture in the cooling chamber. The saturated liquid-vapor coolant mixture is at equilibrium at the lower pressure of the cooling chamber. Heat from the rocket motor component to be cooled is absorbed by the coolant. A portion of the liquid portion of the saturated liquid-vapor coolant mixture is converted into gas phase, the converted portion being less than 100% of the coolant. A portion of the coolant is released from the cooling chamber and the coolant in the cooling chamber is dynamically maintained at less than 100% gas phase of the coolant as the thrust and heat generated by the rocket motor varies.

Abstract: 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.

Abstract: A method and apparatus for launching a space vehicle. Operation of a first portion of a plurality of rocket engines associated with a space vehicle is started such that a flow of a gas is initiated through a duct system under the plurality of rocket engines. The gas is located in an area near the plurality of rocket engines. Operation of a second portion of the plurality of rocket engines associated with the space vehicle is started after the flow of the gas begins traveling through the duct system to launch the space vehicle.

Abstract: Embodiments of a solid propellant rocket motor are provided. In one embodiment, the solid propellant rocket motor includes a pressure vessel having a cavity therein, a solid propellant disposed within the cavity, a nozzle fluidly coupled to the cavity, and a tungsten alloy burst disc positioned proximate the nozzle. The tungsten alloy burst disc is configured to block gas flow through the nozzle when the tungsten alloy burst disc is intact and to fragment at a predetermined burst pressure. Embodiments of a method are further provided for manufacturing a burst disc. In one embodiment, the method comprises the step of forming a burst disc from a tungsten alloy. Embodiments of a burst disc are still further provided. In one embodiment, the burst disc includes an outer annular portion, and a central portion comprising a tungsten alloy.

Abstract: A system for in-flight attitude control and side-force steering includes a thruster body and a plurality of valves capable of generating side thrusts put into communication with the thruster body. The valves are arranged in two sets of valves spaced apart from each other towards the front and towards the rear of the thruster body in substantially symmetrical manner relative to the center of gravity of the vehicle situated on a longitudinal axis (A) of the vehicle. Each set comprises a first pair of valves generating thrust in opposite directions along axes that are not in alignment and are parallel to a first axis, and a second pair of valves generating thrust in opposite directions along axes that are not aligned and are parallel to a second axis. The first and second axes are distinct and perpendicular to the longitudinal axis of the vehicle.

Abstract: An apparatus and method for use in conducting an eruption reaction are disclosed. The apparatus includes a catalytic solids container with a mouth and fluid egress opening and a trigger device or mechanism that allows for the controlled release of a catalytic solid into an eruptible fluid. The catalytic solids container may be adapted to be coupled to a container for an eruptible fluid.

Abstract: A replaceable thrust generating structure for an air vehicle, which includes one or more mounts, which are attachable to the air vehicle, and a thrust generating structure attached to each of the one or more mounts, and wherein the thrust generating structure includes an NMSET element.

Abstract: The present invention relates to improved rocket engine systems. In one embodiment, an improved rocket engine system includes a propellant source, at least one power source, at least one power source motor, a rocket engine, and at least one pump. The improved rocket engine system may further include at least one of the following: at least one controller, at least one propellant valve, and a propellant pressurizing source.

Abstract: A system and a method of generating energy in a power plant using a turbine are provided. The system includes an air separation unit providing an oxygen output; a plasma generator that is capable of generating plasma; and a combustor configured to receive oxygen and to combust a fuel stream in the presence of the plasma, so as to maintain a stable flame, generating an exhaust gas. The system can further include a water condensation system, fluidly-coupled to the combustor, that is capable of producing a high-content carbon dioxide stream that is substantially free of oxygen. The method of generating energy in a power plant includes the steps of operating an air separation unit to separate oxygen from air, combusting a fuel stream in a combustor in the presence of oxygen, and generating an exhaust gas from the combustion. The exhaust gas can be used in a turbine to generate electricity. A plasma is generated inside the combustor, and a stable flame is maintained in the combustor.

Abstract: A clustered, fixed cant, throttleable rocket assembly is used to propel and a steer a vessel in terrestrial or extraterrestrial applications. The fixed cant of each of at least three individual rockets in the cluster provides the steering input to the overall assembly. More specifically, by changing the propellant flow rate to the individual rocket engines relative to one another, the overall thrust vector of the rocket assembly may be selected to provide a desired steering input to the vessel. A measured vessel orientation may be compared with a desired vessel orientation to determine what steering input is required to achieve the desired vessel orientation.

Abstract: An injection head including an annular distribution cavity for distributing a propellant upstream from an injection plate supporting injectors. The cavity includes a multiply-perforated distribution grid distributing the propellant, which grid is coaxial with the distribution cavity and of a concave shape that projects into the distribution cavity, the grid being fastened to a dome of the cavity.

Abstract: A constant pressure aerospike thruster has been developed. The aerospike thruster includes a hollow combustion chamber with a circular shaped throat area, a spring loaded pintle located inside the combustion chamber and a spike extended out from the pintle through the throat area.

Abstract: A vessel, such as a rocket motor tube, comprised of an insert, and a joint that allows the insert to decouple from the vessel to relieve pressure when subjected to excessive temperatures and/or a predetermined temperature or temperature range.

Abstract: A rocket stage for a spacecraft includes an engine, a tank for storing a fuel and an oxidizer for combustion in the engine, a rocket stage primary structure, and an engine thrust frame that connects and transmits forces between the engine and the primary structure. The engine thrust frame includes at least an internal part thereof arranged internally within the tank. This internal part of the engine thrust frame forms an imperforate partition that divides an interior space of the tank into at least two chambers for storing the fuel and the oxidizer separately from one another. Propellant management devices including liquid guide vanes and refillable liquid reservoirs may be provided in connection with the partition in the tank interior space.

Abstract: A space launcher propulsion system includes: an engine fed by at least two propellant component tanks; a measurement mechanism to measure a quantity of propellant component actually consumed in each of the tanks, by sensors becoming uncovered; an estimation mechanism to estimate instantaneous flow rates of each of the propellant components on the basis of at least one operating parameter of the engine; and a correction mechanism correcting the way the estimation mechanism estimates on the basis of the measured consumed quantities and of the estimated instantaneous flow rates.

Abstract: 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.

Abstract: A midframe portion (313) of a gas turbine engine (310) is presented and includes a compressor section with a last stage blade to orient an air flow (311) at a first angle (372). The midframe portion (313) further includes a turbine section with a first stage blade to receive the air flow (311) oriented at a second angle (374). The midframe portion (313) further includes a manifold (314) to directly couple the air flow (311) from the compressor section to a combustor head (318) upstream of the turbine section. The combustor head (318) introduces an offset angle in the air flow (311) from the first angle (372) to the second angle (374) to discharge the air flow (311) from the combustor head (318) at the second angle (374). While introducing the offset angle, the combustor head (318) at least maintains or augments the first angle (372).

Abstract: 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.

Abstract: A system for the recovery and management of atmospheric gas is disclosed, such as for use as a vehicle propellant in a vehicle propulsion system. The system can include a compressor configured to compress atmospheric gas and first and second storage tanks configured to store liquefied atmospheric gas from the compressor. The second storage tank can have a heater operable to heat liquefied atmospheric gas therein to convert it to a high pressure gas. The second storage tank includes an outlet duct fluidly coupled to the first storage tank for supplying high pressure gas to the first storage tank.

Abstract: A method, and a related material, for utilizing high performance solid rocket propellants, which are molding powders. A propellant molding powder are selected to have a design burning rate and a tailored compaction profile. A morphology of a center-port of a rocket is selected for the design burn rate and a spin-rate. The molding powder is compacted isostatically around a core through application of triaxial pressure therein forming a solid rocket propellant charge with the selected center-port shape. The solid rocket propellant charge is placed in a cartridge or a case. The cartridge is selected from various types of cartridges and specialty charges. The solid rocket propellant molding powders are highly filled with metallic fuels, and have a binder in the range of 4% to 18%, which at least partially coats the surface of the molding powder.

Abstract: Whereby the spiral channels of the bladeless rotor and stator of the Bezentropic Bladeless Turbine are attached to the Laval nozzle, as well as to any preferred modification of the same nozzle, act as an extension of nozzle's divergent end, transforming it into a Bezentropic Bladeless Turbine, all the while retaining the nozzle's efficiency and efficacy, achieved as a result of the maintenance and sustenance of the mono-directional rectified molecular flow of gas, steam or its combination thereof, emitted by the nozzles into the Bezentropic Bladeless Turbines spiral channels to produce mechanical work or thrust.

Abstract: An injector includes an injector body that has at least an injection side surface. A plurality of passages extends within the injector body for injecting combustion fluids at the injection side surface. The plurality of passages include a first set of passages that has a first arrangement defining a first impingement point in space with regard to the injection side surface and a second, different set of passages that has a second arrangement defining a second impingement point in space with regard to the injection side surface.

Abstract: An apparatus for propulsion includes a main body, a fluid flow channel for fluid mass to circulate within a self-sealed apparatus, an air flow channel for air to circulate within a self-sealed apparatus and/or to the open environment, a fluid mass, a rotator having compartments to move the fluid mass centrifugally, a cover providing sealed containment for the apparatus and providing the means to remove the rotator from the main body, and a rotator shaft providing rotation to the rotator by motor. A preferred embodiment includes an apparatus with two or more fluid flow channels for fluid mass to circulate within a self-sealed apparatus. Another preferred embodiment includes an apparatus with two or more air flow channels for air to circulate within a self-sealed apparatus and/or to the open environment. And another preferred embodiment includes the main body that can be manufactured in one or more parts.

Abstract: A liquid propellant rocket engine with a propulsion chamber shutter, the rocket engine comprises a combustion chamber, a propellant injector device placed at an upstream first end of the combustion chamber, a nozzle throat disposed at a downstream second end of the combustion chamber remote from the upstream first end, and a diverging portion disposed downstream from the nozzle throat. A selective shutter device for the nozzle throat comprises an axially-symmetrical shutter member placed downstream from the nozzle throat, and axial rod for controlling the shutter member, a first short centering member for the control rod situated at the upstream first end of the combustion chamber level with the propellant injector device, a second short centering member for the control rod situated in the combustion chamber in the vicinity of the nozzle throat and upstream therefrom, and a system for returning the control rod of the selective shutter device for the nozzle throat to the closed position.

Abstract: A nozzle arrangement for use in a gas thruster is presented. At least one heater micro structure (20) is arranged in a stagnation chamber (12) of the gas thruster. The heater microstructure (20) comprises a core of silicon or a silicon compound coated by a surface metal or metal compound coating. The heater microstructure (20) is manufactured in silicon or a silicon compound and covered by a surface metal coating. The heater microstructure (20) is mounted in the stagnation chamber (12) before or after the coverage of the surface metal or metal compound coating. The coverage is performed by heating the heater microstructure and flowing a gas comprising low quantities of a metal compound. The compound decomposes at the heated heater microstructure (20), forming the surface metal or metal compound coating. The same principles of coating can be used for repairing the heater microstructure (20) in situ.

Abstract: 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.

Abstract: A gas generator assembly includes a propellant chamber housing an amine based propellant. A reaction chamber is coupled with the propellant chamber. The reaction chamber includes a reaction chamber housing, and a porous reaction matrix within the reaction chamber housing. The reaction matrix includes a catalyzing agent, and the catalyzing agent is configured to non-combustibly catalyze the amine based propellant into one or more pressurized gases. An injector is in communication with the propellant chamber. The injector is configured to deliver the amine based propellant to the porous reaction matrix. A discharge nozzle is coupled with the reaction chamber and is configured to accelerate and discharge the one or more pressurized gases. In one example, the gas generator is coupled with one or more of an impulse turbine assembly and an electric generator to form a micro power unit.

Abstract: A propulsion system may be operated by determining pressure in a cryogenic liquid tank storing a fluid and cooling the cryogenic liquid tank in response to determining that the pressure is greater than a predetermined value. The cryogenic liquid tank may be pressurized by admitting a gaseous form of the fluid into the cryogenic liquid tank in response to determining that the pressure in the cryogenic liquid tank is less than a predetermined value.

Abstract: The purpose is to develop, design a very inexpensive rocket to “clean up” the expendable debris orbiting the earth. This liter is a serious pollution of space and a threat to working satellites and the space station. Develop an inexpensive rocket engine with optimum efficiency conversion: propellants-chemicals to acquire the maximum thrust generation offering an inexpensive means to acquire low orbit to space. Parameters selected to minimize fuel consumption: by trying different hydrogen-oxygen mixture ratios. A critical design requirement for the Clean Up space engine is the ease to start and operate reliably in the cold vacuum conditions of space. The space engine will have a reliable pressure feed system, simplicity and dependability multiple starts.