Abstract: A catalyst bed assembly, comprising: an outer housing having an open interior, an inlet leading to the open interior, and an outlet from the open interior; a catalyst bed in the open interior; and a gap between the outer housing and the catalyst bed. The open interior receives material from the inlet. A portion of the material enters the catalyst bed to expose said material to a catalyst so that the material and the catalyst react and create heat within the catalyst bed assembly. A remainder of the material enters the gap between the outer housing and the catalyst bed to cool the catalyst bed assembly. The catalyst bed assembly could be part of a turbopump assembly. The turbopump assembly would further include a nozzle downstream of the outlet; a turbine downstream of the nozzle; and a pump driven by said turbine.

Abstract: A rocket propulsion system, comprising: a rocket engine; and a turbopump supplying fuel or oxidizer to the rocket engine. The turbopump can supply an adjustable flow of the fuel or oxidizer to throttle the rocket engine. The turbopump includes a catalyst bed for decomposing a material to produce a discharge; a mixer section downstream of the catalyst bed for introducing an additional amount of the material to the discharge to produce an exhaust stream having a mass flow; a nozzle downstream of the mixer section; a turbine downstream of the nozzle; and a pump driven by the turbine. The additional amount of said material is selected to produce a desired amount of mass flow.

Abstract: A gas generator, comprising a catalyst section and a mixer section. The catalyst section has openings through which a material enters and reacts with a catalyst to decompose and to produce heat. The mixer section, located downstream of the catalyst section, introduces an additional amount of the material. The heat produced in the catalyst section is sufficient to decompose the additional amount of material introduce into the mixer section without exposure to said catalyst. The catalyst bed assembly could be part of a turbopump assembly. A nozzle downstream of the mixer section directs the exhaust towards a turbine. The turbine drives a pump. The amount of material added to the mixer is selected to produce a desired amount mass flow through the nozzle. Adding material to the mixer causes the turbopump to supply more fuel or oxidizer to a rocket engine. Thus, the rocket engine is throttleable.

Abstract: The occurrence of cavitation is reduced in a rotary pump while the pump performance is maintained. A low-temperature source present in the pump system is utilized, and heat exchange with the fluid flowing to the pump is performed with this low-temperature source, thereby lowering the temperature of the fluid flowing into the pump and lowering the saturated vapor pressure of this fluid, which increases the allowable margin for a decrease in the pressure of the fluid and reduces the occurrence of cavitation. In the case of a liquid rocket engine, a coolant or another propellant B whose temperature is lower than that of the primary propellant A can be employed as this low-temperature source. The propellant B is passed through a heat exchanger 2 by a pump 4, which results in heat exchange with the propellant A from a tank 3, and lowers the fluid temperature. Since the fluid velocity of the propellant A is low at the time of this inflow, there is only slight pressure loss inside the heat exchanger 2.

Abstract: A method is provided for using high concentrations of hydrogen peroxide to drive a turbine (20′) in a turbopump fed rocket engine (12′). The method includes the steps of: (a) receiving fuel into a fuel rich pre-burner (50); (b) receiving high concentrations of hydrogen peroxide into the pre-burner (50); (c) converting the fuel and hydrogen peroxide into a fuel rich gas; and (d) passing the fuel rich gas through a turbine (20′), thereby using high concentrations of hydrogen peroxide to drive the turbine. Thus, by utilizing a fuel rich pre-burner (50) that operates at a very low mixture ratio, the drive gas for a turbine (20′) can be maintained at moderate temperature levels.

Abstract: A gas generator, comprising a catalyst section and a mixer section. The catalyst section has openings through which a material enters and reacts with a catalyst to decompose and to produce heat. The mixer section, located downstream of the catalyst section, introduces an additional amount of the material. The heat produced in the catalyst section is sufficient to decompose the additional amount of material introduce into the mixer section without exposure to said catalyst. The catalyst bed assembly could be part of a turbopump assembly. A nozzle downstream of the mixer section directs the exhaust towards a turbine. The turbine drives a pump. The amount of material added to the mixer is selected to produce a desired amount mass flow through the nozzle. Adding material to the mixer causes the turbopump to supply more fuel or oxidizer to a rocket engine. Thus, the rocket engine is throttleable.

Abstract: The present invention provides for a pressurizer for pressurizing a fluid, comprising a pressurant entrance for the introduction of a pressurant, a fluid entrance for the fluid, a fluid exit for the fluid, and a transfer chamber movable in a cycle with respect to the fluid exit, where for a portion of a cycle the pressurant exerts a force on the fluid inside the transfer chamber. In a preferred aspect of the present invention, the pressurizer further comprises a spindle housing more than one transfer chamber, rotatable about an axis between the fluid entrance and the fluid exit. In another preferred aspect, the transfer chamber comprises either a flexible membrane or a movable piston to separate the pressurant and the fluid.

Abstract: The present invention provides apparatus and methods for integrating structural members inside the body of a propulsion vehicle with tankage used to store fluid propellant and the like. Propulsion vehicles may be made lighter, more compact, cheaper, and easier to manufacture by using pressurized membranes of the tankage to accomplish other structural purposes. More specifically, tanks may be integrated with thrust structures to transfer thrust loads from the engine to the main body of the vehicle. Alternatively, the tanks may be integrated with the vehicle engine. Also, one tank may be integrated with one or more other tanks to form a single pressure vessel with multiple interior chambers. Tankage may additionally be combined with more than one of the foregoing to save additional weight and space. Methods of manufacturing a metallic integrated tank assembly include weld fabrication, machining, spinning, hydroforming, casting, forging, plating, metal deposition, or some combination thereof.

Abstract: The present invention provides for a pressurizer for pressurizing a fluid, comprising a pressurant entrance for the introduction of a pressurant, a fluid entrance for the fluid, a fluid exit for the fluid, and a transfer chamber movable in a cycle with respect to the fluid exit, where for a portion of a cycle the pressurant exerts a force on the fluid inside the transfer chamber. In a preferred aspect of the present invention, the pressurizer further comprises a spindle housing more than one transfer chamber, rotatable about an axis between the fluid entrance and the fluid exit. In another preferred aspect, the transfer chamber comprises either a flexible membrane or a movable piston to separate the pressurant and the fluid.

Abstract: A liquid fuel rocket engine with a closed direct auxiliary flow cycle has a combustion chamber connected to each liquid fuel source through a respective main liquid fuel supply line. A turbo pump is provided in each main fuel supply line. The turbo pump is driven by a gas generator which is connected to a first bypass fuel supply line for supplying fuel to the gas generator. An exhaust gas duct (8, 9, 10) leads from the gas generator into at least one main fuel supply line for feeding gas generator exhaust gas into the main fuel supply line upstream of the turbo pump.

Abstract: A method is provided for using high concentrations of hydrogen peroxide to drive a turbine (20′) in a turbopump fed rocket engine (12′). The method includes the steps of: (a) receiving fuel into a fuel rich pre-burner (50); (b) receiving high concentrations of hydrogen peroxide into the pre-burner (50); (c) converting the fuel and hydrogen peroxide into a fuel rich gas; and (d) passing the fuel rich gas through a turbine (20′), thereby using high concentrations of hydrogen peroxide to drive the turbine. Thus, by utiliizing a fuel rich pre-burner (50) that operates at a very low mixture ratio, the drive gas for a turbine (20′) can be maintained at moderate temperature levels.

Abstract: Various types of valves are disclosed, as well as the use of the same in a reaction control system for space travel vehicle applications. Multiple functions are provided by one of these valve designs by providing multiple flowpaths through the valve. An inflow and outflow tube of this multifunctional valve are fluidly interconnected with first and second chambers, respectively, within the valve. These first and second chambers are isolated from each other by a barrier assembly which may be “removed” at the desired type by a barrier rupture assembly to provide one of the noted multiple flowpaths (i.e., the inflow tube, the first chamber, the second chamber, and the outflow tube). A separate service port is fluidly interconnected with each of the first and second chambers, and a separate service valve may be disposed in each of these service ports to provide additional flowpaths for the multifunctional valve.

Abstract: The invention relates to the field of power machine building and concerns the improvements in turbines for liquid-propellant rocket engines. A turbine includes a casing enclosing a working chamber and a propellant feed passage communicating with the working chamber. A cooling space, and a cooling fluid feed passage also communicate with the cooling space. The turbine also has a turbine wheel to be cooled by a cooling fluid. An annular chamber communicating with an annular slot can allow the cooling fluid to cool the turbine wheel. Also provided is a bush of material having the thermal conductivity higher than that of the casing material, located coaxially to the turbine wheel. The bush is made to have a length longer than that of the turbine wheel. The bush is mounted so that to bulge out of the turbine wheel towards the propellant feed channel. Surfaces of the annular chamber and the annular slot are formed by a surface of the bulging out part of the bush.

Abstract: Various types of valves are disclosed, as well as the use of the same in a reaction control system for space travel vehicle applications. Multiple functions are provided by one of these valve designs by providing multiple flowpaths through the valve. An inflow and outflow tube of this multifunctional valve are fluidly interconnected with first and second chambers, respectively, within the valve. These first and second chambers are isolated from each other by a barrier assembly which may be “removed” at the desired type by a barrier rupture assembly to provide one of the noted multiple flowpaths (i.e., the inflow tube, the first chamber, the second chamber, and the outflow tube). A separate service port is fluidly interconnected with each of the first and second chambers, and a separate service valve may be disposed in each of these service ports to provide additional flowpaths for the multifunctional valve.

Abstract: Plugs for protecting internal spaces of rocket combustion chambers against moisture, dust, and other environmental factors. Each plug includes a central rod with a bracket fastened to it. The plug also includes a sleeve with radially extending structural ribs. Spring-loaded cams with rollers are securely hinged at the ends of the ribs. In addition, a spring-loaded limiter and a valve are mounted on the central rod so that they can possibly move relative to the central rod. The limiter includes a sleeve and a conical shell that are connected by rod ties. The valve includes a metal stamped conical plate that forms a gas-tight fit with the rod. A sealing ring of elastic material is positioned on a peripheral part of the plate. The set of spring-loaded cams and the limiter secure the self-mounting plug in a working position and make the valve operation independent of the operation of other plug components.

Abstract: A rotor system for a rocket engine comprises a first hollow shaft portion and a surrounding annular duct. The interior of the first hollow shaft portion is adapted to receive a first propellant component, and is in fluid communication with a first rotary orifice. The annular duct is adapted to receive a second propellant component, and is in fluid communication with a second rotary orifice located proximate to the first rotary orifice. The rotor system further comprises a third rotary orifice operatively connected to second hollow shaft portion and in fluid communication with the interior of the first hollow shaft portion for discharging a second portion of the first propellant component at second location. The rotary orifices are operatively connected to respective rotary pressure traps that isolate the pressure at the respective rotary orifices from the respective inlet pressures of the respective propellant components.

Abstract: The engine comprises a chamber, fuel and oxidizer booster pumps, a turbopump unit, a gas generator, an electrically driven thrust regulator, and a fuel throttle. A programmable starting and ignition system is introduced, which includes a fuel tank, connected to the thrust regulator, and a first ampoule with starting fuel, an output of which through a starting/cut-off valve is connected to fuel injectors of the gas generator. The aforesaid tank is also connected through a second ampoule to the starting fuel and by a jet to autonomous ignition injectors located in the space of the combustion chamber of the engine. An inclined Archimedian screw pump with a gas turbine as an actuator and an oxidizing gas as a working medium is used as an oxidizer booster pump. The gas is discharged after its use in a turbine into an outlet manifold of the booster pump.

Abstract: A rocket engine comprises first and second combustion chambers with respective combustion chamber liners bounding respective annular passages, wherein the first combustion chamber discharges into the second, and the respective annular passages are in fluid communication with one another. A portion of the effluent from the first combustion chamber flows from the first combustion chamber to the second combustion chamber through the respective annular passages via orifices in the respective combustion chamber liners, so as to provide for effusion cooling of a surface of the second combustion chamber. The first combustion chamber preferably operates fuel rich, reducing the temperature of the effusion cooling gases, which may be further cooled by a portion of unburned fuel. A flow restriction such as a turbine between the first and second combustion chambers provides a pressure differential therebetween that induces flow of effusion cooling gases.

Abstract: A gas generator and ignitor assembly for powering a turbine of a turbopump assembly for a rocket engine comprises an injector and a combustion chamber, the injector having a body member including a fuel inlet and an oxidizer inlet spaced one from the other and communicating with respective radially spaced apart annular members in the body member. Three annuli communicate with the fuel inlet and two annuli communicate with the oxidizer inlet, the annuli which communicates with the oxidizer being positioned between pairs of the other annuli. The body member is enclosed by a plate having an array of bores arranged in two series with three radially spaced apart groups of circular rows in each series. The outer series has 28 groups of triplet bores while the inner series has 14 groups of triplet bores. The annuli which communicate with oxidizer feed bores of each series that are between the other bores of a triplet, the latter bores communicating with annuli that communicate with fuel.

Abstract: A rocket according to the invention includes a motor having a fuel chamber and an oxidizer tank, a nose portion provided forward of the motor, and a nozzle surrounded by fins provided aft of the motor. The fuel chamber is provided with a relatively central solid propellant, and a hybrid fuel surrounds the solid propellant. The oxidizer tank is filled with a reactant and coupled to the forward end of the fuel chamber. A pathway is provided between the tank and the fuel chamber for the passage of reactant therethrough. At least one of a valve and a barrier is coupled in the pathway to prevent passage of the reactant into the fuel chamber until after the solid propellant is at least partially consumed. After the solid propellant is at least partially consumed, the valve is opened and/or the barrier is removed to permit the passage of reactant into the fuel chamber.

Abstract: A hybrid heater vaporization system (280, 380) including hybrid heaters (250, 350) can be used to vaporize liquid oxygen as it enters a motor (200, 300) to improve the combustion characteristics of a hybrid rocket (11). The motor (200, 300) preferably includes hybrid fuel both in a substantially cylindrical portion (216, 316) in the substantially cylindrical portion of the motor and a substantially multi-toroidal shaped portion (217, 316) in the forward end of the motor (200, 300). The vaporization system (280, 380) of the present invention finally makes hybrid rockets (11) practical for aerospace applications.

Abstract: A HAN TEAN mixing gas generator pressurization system is proposed for use on pressure-fed rockets. HAN (hydroxyl ammonium nitrate) and TEAN (triethanol ammonium nitrate) are salts dissolved in water to form a single liquid propellant. Combustion of HAN TEAN results in water, carbon dioxide, and nitrogen at high temperatures and pressures. A cryogenic primary working fluid is mixed the HAN TEAN exhaust products to provide warm high-pressure gas to pressurize the rocket's main propellant tanks.

Abstract: A method and system for delivering pressurized propellants to a rocket engine, that has significant advantages over the current state-of-the-art. One of the propellants, the "pressurizing propellant", is at least partially vaporized and the vapor is in pressure communicating relationship with other propellants on board the rocket-propelled vehicle. This vapor pressure pressurizes the propellants to a sufficient degree that they can be charged directly to the rocket engine, or the pressure may be boosted through pumps, if required. Moreover, the pressurized vapor may be used in other applications on board the vehicle, such as orbital adjustment, attitude control, station keeping, and the like. In several embodiments, the propellants are contained in variable volume reservoirs, exemplified by bladders and diaphragms. These variable volume reservoirs are preferably not subjected to tensile stresses when expanded, and are preferably designed for controlled volume reduction, when being drained of propellant.

Abstract: In space exploration, an attitude control engine or a thruster may have an extended period, such as years, of inactivity prior to being required to complete a desired mission. The fuel for this mission is contained in a fuel tank for the same extended period of time and tends to corrode elastomeric components. To isolate the fuel from elastomeric components during the storage period, the fuel tank has a metallic shear seal. When actuated, a piston ruptures the shear seal and delivers the fuel to a power plant. A shear seal that is unitary with the inner walls of the fuel tank eliminates the need for an interfacial weld that may constitute a shear seal failure site.

Abstract: For a liquid rocket engine including the SSME the gaseous heat exchanger that is in the liquid oxygen turbopump exit flow path that serves to gasify the liquid oxygen for pressurizing the liquid oxygen tank and the POGO is modified by changing the flow circuit to replace the liquid oxygen with high pressure hydrogen tapped off of the low pressure fuel turbopump so that hydrogen/hydrogen at a favorable pressure differential is in indirect heat relation for providing a safe environment and avoiding what may result in a catastrophic failure in the event the heat exchanger fails. An external heat exchanger in communication with the heat exchanger in the liquid oxygen turbopump is provided to gasify the liquid oxygen for the pressurizing purposes.

Abstract: Disclosed is a rocket engine which uses pressurized, vaporized propellant from the engine cycle to drive the boost pumps that provide propellants to the primary pumps. Additionally, an accumulator tank is provided which stores some of the pressurized, vaporized propellant and selectively releases the propellant to drive the boost pumps at subsequent engine start-up.

Abstract: A valve assembly for transferring fluid from a fluid supply tank to a fluid holding tank. The valve assembly includes a housing having a first port for communication with the fluid supply tank and a second port for communication with the fluid holding tank. A drum is rotatably mounted in the housing and has a passageway therein. The passageway has an opening located at each end thereof positioned on an outer surface of the drum for alignment with one of the ports in the housing. The drum is rotatable between a flow position in which the openings are aligned with the ports to permit fluid flow from the fluid supply tank to the fluid holding tank and a shut-off position in which one of the openings is not aligned with the ports to prevent flow from the supply tank to the holding tank. The valve assembly further comprises an actuator for rotating the drum within the housing between the fluid flow and shut-off positions.

Abstract: One of the propellants used in a bipropellant rocket engine has a vapor pressure sufficiently high under prevailing condition to generate vapor at a pressure sufficiently high to pressurize both the volatile component and the second rocket propellant and provide substantially the sole means for forcing the two rocket propellants from their storage to the rocket engine. In the preferred arrangement, both the components will be stored within the same pressure vessel, one preferably being contained within an expandable/collapsible bladder.

Abstract: A hybrid heater vaporization system (280, 380) including hybrid heaters (250, 350) can be used to vaporize liquid oxygen as it enters a motor (200, 300) to improve the combustion characteristics of a hybrid rocket (11). The motor (200, 300) preferably includes hybrid fuel both in a substantially cylindrical portion (216, 316) in the substantially cylindrical portion of the motor and a substantially multi-toroidal shaped portion (217, 316) in the forward end of the motor (200, 300). The vaporization system (280, 380) of the present invention finally makes hybrid rockets (11) practical for aerospace applications.

Abstract: A regenerative liquid propellant rocket motor including a body having an integrally formed central tube. A primary combustion chamber is disposed in the body. A secondary combustion chamber is formed by the central tube of the body. A reservoir of liquid propellant is disposed in the body. A regenerative piston is slidably disposed in the body for delivering the liquid propellant to the primary and secondary combustion chambers. The piston includes a plurality of injection ports for injecting the liquid propellant. An ignitor is disposed in the body for initiating combustion of the liquid propellant. A nozzle is formed at a free end of the central tube for producing thrust, wherein upon ignition of the ignitor combustion occurs in the primary combustion chamber driving the piston into the liquid propellant reservoir and the liquid propellant is injected into the primary combustion chamber to produce combustion products which are channeled into the secondary combustion chamber and expanded through the nozzle.

Abstract: A regenerative liquid propellant engine system includes a body having opposed ends. A combustion chamber is disposed in the body. A reservoir of liquid propellant is also disposed in the body. A regenerative piston is slidably disposed in the body for delivering the liquid propellant to the combustion chamber. When combustion occurs in the combustion chamber the piston is driven into the liquid propellant reservoir and the liquid propellant is injected into the combustion chamber to produce combustion products. A storage tank stores a quantity of liquid propellant and a pump delivers the liquid propellant from the storage tank to the liquid propellant reservoir. The regenerative piston produces the necessary pumping pressure to operate the pump, thus, the mechanical complexity and inefficiencies of other pumping techniques is avoided.

Abstract: There is provided a system for the delivery of a gaseous product to a reaction chamber. The system includes a fuel tank containing a hypergolic liquid displaced from the fuel tank by a pressurized gas whose flow is controlled by a thruster control valve and directed to a gas generator. Pressurant gas entrained in the hypergolic liquid is not decomposed in the gas generator and causes pressure oscillations. The pressure oscillations are minimized by reducing the diameter of the gas bubbles or by disposing a gas fluid resistor between the gas generator and the reaction chamber.

Abstract: A system for pressurizing a liquid oxygen tank (20) in a pressure-fed hybrid rocket (100) provides pressure at a first end of the tank (20) so that liquid oxygen can exit the second end. A hybrid heater (50) creates the pressure by mixing gaseous oxygen with the solid fuel (51) in the hybrid heater (50) to create an extremely high temperature. Helium is then mixed with the combustion products via an inlet (60) and this hot, inert mixture enters the liquid oxygen tank (20) via a diffuser (70). A second hybrid heater (90) pressurizes the gaseous helium sphere (40) to reduce the pressure decay caused by the withdrawal of the helium. Two of the hybrid heaters (50 and 90) are fed gaseous oxygen from a sphere (30). A novel ignition system is used to ignite the hybrid heaters (50, 90, 250). One of the hybrid heaters (250) is used to ignite the hybrid motor (100) in the rocket (11).

Abstract: A device for pressurizing at least one propellant tank feeding the combustion members of a space vehicle, the tank(s) being connected to one or more pipes for feeding said combustion members, and at least one of the propellants being catalytically or thermally decomposable, the device comprising: tapping means placed at the outlet from the decomposable propellant tank to tap a portion of said propellant, drive gas generation means for delivering drive gases, suction means formed by a converging-diverging nozzle connected firstly to the drive gas generation means and secondly to the tapping means to suck decomposable propellant from the tapping means by lowering the pressure of the drive gases to a value that is less than that required for pressurizing the tank(s), and decomposition means placed at the outlet from the suction means for decomposing the sucked propellant into pressurization gas for pressurizing the tank(s).

Abstract: Thrust-generating device with rocket engines arranged in pairs opposite each other for multiple ignition as well as individual and group activation and with an amount of fuel sufficient for the duration of the mission. The thrust-generating device is designed as a closed, storable, modular unit. The rocket engines are operated with liquid fuel and liquid oxidizing agent. At least one tank is provided for each fuel component. The tanks are in the filled, pressureless state before the activation of the thrust-generating device. The rocket engines and a fuel distributor of block design are integrated within a central bulkhead. The thrust-generating device is surrounded by a sealed outer jacket and is divided by the bulkhead into two separate chambers. The outer jacket is lined on the inside with a "self-sealing" film. The oxidizing agent is stored in one chamber, and the fuel in the other chamber. The empty spaces in the chambers contain an adsorbent filling.

Abstract: There is provided a system for expelling a liquid fuel component from a tank to a rocket engine utilizing a pressurized gas to displace the propellant. A pressurized high vapor pressure liquid is combined with the pressurized gas reducing the volume of pressurizing agent required to expel a desired volume of the liquid fuel component. Since the volume of pressurizing gas is less, the weight of the propelled craft may be reduced or additional fuel carried. In another embodiment of the invention, the vapor phase of the high vapor pressure liquid is used as a reference pressure to control other systems.

Abstract: There is provided a two-stage system for expelling a liquid fuel component from a main fuel tank to a rocket engine. A first vessel contains a first gas pressurized to the desired operating pressure at a minimum operating temperature. Since the first gas is subject to pressure increase due to environmental heating, opening a first valve exposes the pressurized first gas to a pressure relief valve, returning the pressure to the desired operating pressure. Subsequent opening of a second valve causes the pressurized first gas to displace a partition, such as a piston, expelling a desired quantity of the fluid. A gas generator converts the fluid to a second gas that expels a desired quantity of fuel from a main fuel tank. The two-stage system is at an equilibrium pressure. The only active control required for the system is a controller to meter the flow of fuel from the main fuel tank to the engine.

Abstract: A fluid storage tank utilizing an isolation seal in which a leak proof isolation seal covers the entire head of the piston of the storage tank requiring only a single fluid tight seam to seal off the fluid storage chamber from the remainder of the storage tank. The isolation seal has a circumferential notch located proximate to its outer periphery which permits the isolation seal to rupture along a predetermined path. Elastomeric annular seals on the piston provide a dynamic seal to prevent the fluid from leaking while the piston is moving to eject the stored fluid. The isolation seal is welded or otherwise fastened to the interior surface of the tank at a location behind the elastomeric seals so that the latter need not pass over the weld during movement of the piston. The interior of the tank has a first or larger diameter section, a second or smaller diameter section, and a tapered transition section which is disposed therebetween just past the rupture zone of the isolation seal.

Abstract: A liquid hydrogen-liquid oxygen (LH.sub.2 -LOX) rocket engine system wherein a conventional oxidizer turbo pump is replaced with a jet pump--a jet oxidizer pump. Beneficial features of the jet oxidizer pump system, over a conventional oxidizer turbo pump system, include: 1) it does not require a complicated seal system, 2) it has no moving parts, and 3) it is lighter. All these features improve the operational reliability of the rocket engine while decreasing its manufacturing cost.

Abstract: The present invention relates to a system for feeding the engines of a space vehicle with at least one liquid propellant, the system comprising for the purpose of storing said liquid propellant under a pressure greater than the feed pressure, firstly a main tank having no liquid expulsion device, and secondly an auxiliary tank that is smaller than the main tank and that is provided with a device for separating the liquid phase from the gas phase, switching means enabling the auxiliary tank to be selectively refilled with liquid propellant from the main tank; such that liquid propellant that is free from pressurization gas bubbles can always be delivered regardless of the flight conditions of the vehicle, and in particular when it is not possible to take said liquid propellant from the main tank.

Abstract: There is provided a two-stage system for expelling a liquid fuel component from a main fuel tank to a rocket engine. A pressurized high vapor pressure liquid combined with the vapor of that liquid expel a desired volume of a gas forming liquid from a first stage fuel tank. A gas generator converts this liquid to a gas that expels a desired quantity of fuel from the main fuel tank. The two-stage system is at an equilibrium pressure. The only active control required for the system is a controller to meter the flow of fuel from the main fuel tank to the engine. The pressure drop in the system resulting from the discharge of fuel causes the high vapor pressure liquid to partially evaporate charging more gas into the system until equilibrium is achieved again.

Abstract: There is provided a system for expelling a liquid fuel component from a tank to a rocket engine utilizing a pressurized gas to displace the propellant. A pressurized high vapor pressure liquid is combined with the pressurized gas reducing the volume of pressurizing agent required to expel a desired volume of the liquid fuel component. Since the volume of pressurizing gas is less, the weight of the propelled craft may be reduced or additional fuel carried. In another embodiment of the invention, the vapor phase of the high vapor pressure liquid is used as a reference pressure to control other systems.

Abstract: A fluid storage and expulsion system having a rigid tank, the tank having two propellant chambers and two independently operable piston assemblies slidable in their propellant chambers for expelling propellant. In a first embodiment, a coaxial fluid storage expulsion system having inner and outer pistons is provided. Accordingly, the inner piston is provided which is slidable in an inner annular space for expelling fluid therein through a second opening. As well, an outer piston is provided which is slidable in an outer annular space for expelling a second fluid through a first opening. In a further embodiment, the pistons and tanks are positioned in series. As such, the rigid tank is provided with forward and aft propellant chambers as well as forward and aft piston assemblies. A piston guide, which extends the length of the rigid tank is positioned concentrically within the tank.

Abstract: Disclosed is a space launch system having a main engine which uses turbopumps to provide fuel and oxidizer to the main engine combustion chamber, and one or more booster engines that use high pressure fuel and/or oxidizer to pressurize the fuel and oxidizer tanks of the booster engine to obtain the pressure differential necessary to drive fuel and oxidizer from the fuel and oxidizer tanks of the booster engine to the combustion chamber of the booster engine.

Abstract: Disclosed is a method and apparatus for producing combustion chamber pressure in excess of 1500 psia at thrust levels in excess of 100,000 pounds in an expander cycle rocket engine.

Abstract: The rocket engine is fitted with an injection device comprising a single injection plate having a central portion that defines the end of the burn zone of the combustion chamber and that contains first propellant injectors having a high mixing ratio adapted to the conditions under which hot gases are produced in the burn zone, and a peripheral portion in the form of a ring that defines the end of an annular cavity of a gas generator and that contain second propellant injectors having a low mixing ratio and adapted to the conditions under which hot gases are produced in the gas generator. The gas generator is thus integral with the combustion chamber and the annular cavity of the gas generator is separated from the burn zone of the combustion chamber by a cylindrical wall that extends axially perpendicularly to the injection plate and that is fixed thereto in sealed manner.

Abstract: A fluid storage tank utilizing a shear seal in which a leak proof static shear seal covers the entire head of the piston of the storage tank requiring only a single fluid tight seam to seal off the fluid storage chamber from the remainder of the storage tank. The shear seal has a circumferential notch located proximate to its outer periphery which permits the shear seal to rupture along a predetermined path. Elastomeric annular seals on the piston provide a dynamic seal to prevent the fluid from leaking while the piston is moving to eject the stored fluid. The head of the piston flattens any protruding portions of the ruptured shear seal against the walls of the storage tank.

Abstract: Disclosed is a rocket engine having a low operating mode in which propellants are provided to the combustion chamber by the combined effect of boost pumps and the pressure differential between the propellant tanks and the combustion chamber, and a high operating mode in which primary pumps further increase the pressure at which propellant is provided to the combustion chamber.

Abstract: This invention relates to an injector in which fluid is injected into a reaction region 30 after pressurization and translation in traveling cells 20 between crests 46 of traveling waves of one or more waveplates 12 made of shear transducer material. Multiple phases of applied electrical stimulation cause electrical segments of the waveplate to form traveling waves by shearing in a predetermined sequence of amplitudes. Inlet wave amplitude taper 14 provides a prescribed state of inlet fluid acceleration. Outlet wave amplitude taper 18 predetermines the degree of mixing of two or more exiting fluids. Self filtering, valve action, high pressure, and high mass flow are provided with few and relatively benign modes of apparatus degradation due to friction. High system efficiency results from absence of rubbing and the recirculation of stored energy. The fluid delivery rate is smooth, continuous, and electrically variable. Balance of the mass flow rate of two or more fluids is electrically controlled.

Abstract: A cold vapor pressurization system for outer space vehicles that is used prior to and during engine burn to pressurize its fuel tank and oxidizer tank to produce the required engines net positive suction pressure (NPSP). The engine start system uses reaction control system (RCS) hydrazine to pass through a gas generator where it is disassociated. The by-products are then introduced into the hot side of the propellant heat exchangers. An oxidizer feed system and a fuel feed system pass through the respective cold sides of the heat exchangers to vaporize the propellants which are then used to produce the required engine NPSP for the oxidizer tank and fuel tank respectively. This system could also be used to provide engine NPSP during engine burn.