Abstract: This invention is a packaged propellant air-induced variable thrust rocket engine that has a vast number of uses and applications for this invention. The primary purpose of the device described here is to provide a light weight, torque and vibration free thrust generator for the propulsion of aircraft. This device will facilitate the fabrication of very light weight aircraft because of the lack these forces. This device can also be used anywhere high velocity air flow and or the resulting thrust is needed. The invention uses aerodynamic principles to compress and accelerate the incoming air, prior to it being heated and accelerated by a short duration burst of thermal and kinetic energy from discrete packets of a mixture of oxidizable fuels. The heated and accelerated air then expands as it travels thru the device providing thrust.

Abstract: An apparatus for igniting a larger rocket motor is provided. The apparatus may be a smaller rocket motor that can be ignited hypergolically, when a pressurized oxidizer contacts hypergolic fuel grains of the smaller rocket motor. The hypergolic ignition causes the larger rocket motor to be ignited. The hypergolic ignition of the smaller rocket motor may be stopped after the larger rocket motor is ignited, and the remaining hypergolic fuel grains and the pressurized oxidizer can be reserved for reigniting the larger rocket motor at a later time.

Abstract: The invention relates to the field of aerospace propulsion, and more particularly to a propulsion assembly having at least one channel defined by an inner wall and an outer wall, the channel presenting an inlet opening and an outlet opening, the assembly including a plurality of rocket engines oriented axially in the channel and forming ejectors for accelerating a flow of air in the channel to supersonic speed. Downstream from the outlet opening from the channel, the inner wall forms a single expansion ramp nozzle. The invention also provides an aerospace craft and a method of propulsion using such a propulsion assembly.

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: Specific impulse and rocket engine efficiency can be improved by injecting reactants, e.g., a propellant combination or a monopropellant and a catalyst, into a plasma flow of a rocket engine. In some aspects, a catalyst or a propellant is carried by plasma formed by passing a flow of a feed gas through an electrical arc. In some aspects, reactants are combusted in supersonic plasma flow to generate combustion ionization in the plasma flow.

Abstract: An ignition system and method of igniting the ignition system includes a main catalyst section in a staged relationship with a pilot-catalyst section to stage a decomposition though the pilot-catalyst section which preheats the main catalyst section.

Abstract: An injection device including at least one injection plate adjacent a combustion space of a combustion chamber, and at least one first injection nozzle including a first entry bore having a first discharge into the combustion chamber, and a first orifice bore, having a cross-sectional dimension less than or equal to the first entry bore, coaxially arranged with the first entry bore and remote from the first discharge. At least one second injection nozzle includes a second entry bore having a second discharge into the combustion chamber, and a second orifice bore, having a cross-sectional dimension less than or equal to the second entry bore, coaxially arranged with the second entry bore and remote from the second discharge. The instant abstract is neither intended to define the invention disclosed in this specification nor intended to limit the scope of the invention in any way.

Abstract: An injection device including at least one injection plate adjacent a combustion space of a combustion chamber, and at least one first injection nozzle including a first entry bore having a first discharge into the combustion chamber, and a first orifice bore, having a cross-sectional dimension less than or equal to the first entry bore, coaxially arranged with the first entry bore and remote from the first discharge. At least one second injection nozzle includes a second entry bore having a second discharge into the combustion chamber, and a second orifice bore, having a cross-sectional dimension less than or equal to the second entry bore, coaxially arranged with the second entry bore and remote from the second discharge. The instant abstract is neither intended to define the invention disclosed in this specification nor intended to limit the scope of the invention in any way.

Abstract: A rocket propulsion system having a liquid fuel tank and a liquid oxidizer tank, where a liquid inert gas tank supplies liquid inert gas to a liquid inert gas pump driven by a turbine of the engine to pressurize the liquid inert gas, which is passed through a heat exchanger around the nozzle to vaporize the inert gas, that is then used to pressurize the liquid oxidizer tank so that a liquid oxidizer pump is not needed, or to pressurize both the liquid oxidizer tank and the liquid fuel tank so that a liquid oxidizer pump and a liquid fuel pump are not needed in the rocket engine.

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: Supply of a liquid component in a combustion chamber of a rocket engine is controlled by a feed valve provided with an obturator mobile between a pen position and a closed position of at least one supply pipe, which has an inlet that communicates with a tank for containing the liquid component and an outlet that communicates with the combustion chamber; the displacement of the obturator from its closed position to its open position being triggered by a pressurized fluid supplied to the outlet of the supply pipe.

Abstract: An improved coaxial injector for injecting first and second propellants into a combustion chamber of a rocket engine is provided. The injector includes a first plate, a second plate, a plurality of channels formed in the first plate, and a plurality of tubes extending through the second plate and into the plurality of channels. The tubes inject a first propellant, such as an oxidizer, into the combustion chamber, and the channels inject a second propellant, such as a fuel, around the tubes and into the combustion chamber.

Abstract: An injection device including at least one injection plate adjacent a combustion space of a combustion chamber, and at least one first injection nozzle including a first entry bore having a first discharge into the combustion chamber, and a first orifice bore, having a cross-sectional dimension less than or equal to the first entry bore, coaxially arranged with the first entry bore and remote from the first discharge. At least one second injection nozzle includes a second entry bore having a second discharge into the combustion chamber, and a second orifice bore, having a cross-sectional dimension less than or equal to the second entry bore, coaxially arranged with the second entry bore and remote from the second discharge. The instant abstract is neither intended to define the invention disclosed in this specification nor intended to limit the scope of the invention in any way.

Abstract: There is provided an injector assembly having two or more oxidizer manifolds and/or two or more fuel manifolds for delivery of liquid propellants to a combustion chamber such that combustion instability is reduced or eliminated during throttling. Delivery of the oxidizer to the oxidizer manifolds is controlled by an oxidizer valve, which may comprise an integral valve. The oxidizer passes from the oxidizer manifolds into the oxidizer element and then into the combustion chamber. The multiple oxidizer manifolds allow the oxidizer to be provided through selective openings of the oxidizer element thus reducing the change in pressure drop across the oxidizer element to thereby reduce or eliminate combustion instability and other problems. Additionally, the injector assembly may also include a lift-off seal or a filler fluid source to fill any temporarily unused oxidizer manifolds with an oxidizer or filler fluid.

Abstract: A vehicle includes at least one polyoxymethylene structural support member. The polyoxymethylene structural support member includes a polyoxymethylene component that is a propellant that provides thrust to the vehicle upon pyrolysis or combustion of the polyoxymethylene component of the product of pyrolysis of the polyoxymethylene component. The vehicle can be a satellite or other type of spacecraft.

Abstract: The present invention is a constant volume rocket motor that uses a non-detonating constant-volume, bipropellant combustion process in pulse-mode operation. Opening and closing of the combustion chamber exhaust outlet is controlled by an actuated reciprocating thrust valve (RTV). Fuel enters the combustion chamber at low pressure with the RTV closed. The valve opens after or during combustion when combustion chamber pressure is at or near maximum. The motor has applications in reaction control systems and attitude control systems in spacecraft.

Abstract: A supersonic combustion apparatus and method for using the same including a fixed geometric nozzle having a converging area, throat, and a diverging area, at least one fuel injection means and at least one flame stabilization means located in the divergent area, and an exit plane adjacent and downstream to the diverging area, where an initial first injection/flame stabilization means is located in the diverging area and the exit plane Mach is varied by heat addition in the diverging area by at least one more fuel injection means.

Abstract: A method and apparatus for augmenting thrust in a rocket traveling through atmospheric gas. Rocket motor designs are provided where a throat(s) from one or more rocket motors eject high-speed primary exhaust gas in a configuration which peripherally surrounds an outlet for induced, secondary gas. The secondary gas is mixed with the jet of primary exhaust gas to add momentum, and therefore thrust. Either expansion deflection or plug type rocket discharge nozzles can be utilized. In one embodiment, a thrust augmentation of over one hundred percent is achieved. In another embodiment, a plurality of rocket motor assemblies each containing a thrust augmenting rocket motor design is affixed to a rocket body. Such rocket motors enhance rocket thrust performance, and enables more efficient payload to rocket motor selection, or, alternatively, allows higher loads to be carried with the same amount of thrust.

Abstract: A bi-propellant injector (66) includes a first injector element (68) and a second injector element (70) injecting a first propellant (69) and a second propellant (71), respectively, into a combustion chamber (53). A flame-holding zone igniter (74) is adjacent to and ignites recirculation of at least a portion of the first propellant (69) and at least a portion of the second propellant (71) within a flame-holding zone (76).

Abstract: In a supersonic nozzle incorporating injectors and a combustion chamber as part of an expander cycle rocket engine, the oxidizer is injected in two streams. One of the streams, preferably a small fraction of the total, is injected into an upstream or preburner section of the combustion chamber and the other to a downstream or main section of the chamber. The preburner combustion gas is cooled in a substantially uniform manner to a moderate temperature by cooling the bulk of the gas rather than cooling only the gas in a boundary layer adjacent to the chamber wall. The combustion gas produced in the downstream section is hotter, and heat from that gas is drawn through the chamber wall into a jacket. The limited combustion in the preburner permits the use of a cooling element with highly intimate heat exchange construction, extracting a high level of energy from the preburner gas without damage to the cooling element and an overall improvement in the regenerative cooling.

Abstract: A hypergolic fuel system comprises hydrogen peroxide, silane and a liquid fuel. The hypergolic fuel system is employed in a method for producing thrust, for example in a rocket, by contacting hydrogen peroxide with silane to obtain initial ignition and thereafter feeding a liquid fuel for combustion.

Abstract: The reduced toxicity hypergolic bipropellant fuels of the present invention contain a hypergolic fuel and rocket grade hydrogen peroxide oxidizer, preferably HTP. The hypergolic fuel may be a reactive fuel or a catalytic fuel. The rocket grade hydrogen peroxide oxidizer consists of about 90 weight % to about 99 weight % H2O2, more preferably about 94 weight % to 99 weight % H2O2. However, hypergolic ignition may be attained with a H2O2 content as low as 70 weight % for some reactive fuels. The reactive fuel comprises about 6 weight % to 10 weight % reducing agent in a solvent. The catalytic fuel comprises about 6 weight % to 10 weight % catalytic agent in a solvent. The reactive fuels contain ingredients that are intrinsically reactive with rocket grade hydrogen peroxide. Upon contact with rocket grade hydrogen peroxide oxidizer, reactive fuels react vigorously with the hydrogen peroxide resulting in ignition. The catalytic fuels are produced by dissolving a catalytic agent in a solvent.

Abstract: A marine propulsion system is disclosed that includes at least one nozzle having a combustion chamber therein. The nozzle has an exit end that is shaped so as to flare outwardly. A fuel reformer separates gasoline or other liquid hydrocarbon fuel into hydrogen gas and carbon monoxide and a fuel line feeds the hydrogen gas into the combustion chamber of the nozzle. Combustion air is also fed into the combustion chamber. In addition, measured quantities of raw sea water are also delivered into the combustion chamber. An electrical igniter within the combustion chamber ignites the hydrogen gas causing the water to flash into steam. The steam exhausts through the exit end of the nozzle resulting in a forward thrust.

Abstract: A system configured to deliver dry pressurized air, derived from ambient air, to underground or outdoor cables, conduit, waveguides, plenums or other air spaces for the purpose of preventing the ingress of moisture or contamination. The system includes a housing and a modular membrane pressurization unit removably received within an interior space of the housing. The housing has a base configured for being at least partially buried below ground level and an air intake above ground level. The modular membrane pressurization unit intakes humid ambient air from the environment surrounding the housing and generates a continuous supply of dry pressurized air housing that is routed to the underground air space.

Abstract: A reduced toxicity fuel satellite propulsion system including a reduced toxicity propellant supply (10) for consumption in an axial class thruster (14) and an ACS class thruster (16). The system includes suitable valves and conduits (22) for supplying the reduced toxicity propellant to the ACS decomposing element (26) of an ACS thruster. The ACS decomposing element is operative to decompose the reduced toxicity propellant into hot propulsive gases. In addition the system includes suitable valves and conduits (18) for supplying the reduced toxicity propellant to an axial decomposing element (24) of the axial thruster. The axial decomposing element is operative to decompose the reduced toxicity propellant into hot gases. The system further includes suitable valves and conduits (20) for supplying a second propellant (12) to a combustion chamber (28) of the axial thruster, whereby the hot gases and the second propellant auto-ignite and begin the combustion process for producing thrust.

Abstract: A fuel containing methylacetylene-propadiene, commonly referred to as MAPP gas, produces thrust in a flight vehicle having a pulse detonation engine. The MAPP gas fuel of this invention may be used alone or combined with other conventional fuels such as hydrogen, JP-4, JP-5, JP-10, kerosene or any other suitable hydrocarbon fuel. Such hydrocarbon containing fuel includes, but is not limited to, acetylene, methane, ethylene, propane, butane or liquified petroleum gas. MAPP gas fuel is mixed with an oxidant containing oxygen or air and ignited. The detonation wave created produces thrust for the flight vehicle. A method of powering a flight vehicle having a pulse detonation engine with MAPP gas fuel is also disclosed.

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 system for propelling a watercraft using hydrogen. The system comprises a combustion chamber, an accumulator system, an ignition system, and a propulsion control system. The combustion chamber defines an upper portion and a lower portion. The accumulator system stores pressurized fluid. A first check valve is arranged to allow water to flow from the exterior of the watercraft into the lower portion of the combustion chamber. A second check valve is arranged to allow water to flow from the lower portion of the combustion chamber to the accumulator system. A propulsion control valve is arranged to control the flow of water from the accumulator system to the exterior of the watercraft. A mixture of hydrogen and oxygen is introduced into the upper portion of the combustion chamber.

Abstract: The non toxic bipropellent of the present invention contains a non-toxic hypergolic miscible fuel (hereinafter referred to as “NHMF”) and a rocket grade hydrogen peroxide. This NHMF has rapid ignition capability and minimizes the formation of precipitate, even when exposed to extreme heat or water. The NHMF of this invention contains 5 species. Namely, a polar organic species miscible with hydrogen peroxide; a propagator, which may be substituted or unsubstituted amines, amides or diamines; an inorganic metal salt, which reacts to form a catalyst in solution or as a colloid; acetic acid; and alkali acetate. The inorganic metal salt is miscible with the polar organic species and the propagator in solution.

Abstract: A method for using an injector to combust a first fluid and a second fluid comprises a directing the first fluid step and a shrouding step. The directing step comprises directing the first fluid in a central coaxial flow generally perpendicularly through a first fluid passage that extends through an injector faceplate, and into a combustion chamber. In the shrouding step, a length of the central coaxial flow adjacent to the injector faceplate is shrouded with an outer coaxial flow. The shroud is formed by directing a portion of the second fluid through an annular/conical passage in the injector faceplate that surrounds the first fluid passage. The annular/conical passage is arranged such that the outer coaxial flow is directed radially inward and impinges on the central coaxial flow a distance after the central coaxial flow has exited the first fluid passage and after the outer coaxial flow has exited the annular/conical passage.

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

Abstract: A simplified rocket engine a rocket engine has a rotor assembly with an ultracentrifugal (over 200 meters/second tangential velocity) liquid pump arranged around a combustion chamber so as to provide both forced convective and Coriolis-effect and centripetal-acceleration enhanced free convective regenerative cooling to the combustion chamber while pressurizing the liquid propellant. Both fuel and oxidizer may be stored a cryogenic temperatures and pumped around the combustion chamber for regenerative cooling. The combustion chamber may be structured to rotate together with the pump to provide Coriolis-effect and centripetal-acceleration enhanced combustion. The rotor assembly is driven directly by a tangential component of the primary thrust vector by means of tilted nozzles or by vanes, flutes, or other reaction surfaces.

Abstract: A liquid gas generator system supplies gas pressure only when it is needed. Hydrazine and hydrazine blends have been considered for liquid gas generators because of their ability to decompose at ambient conditions on an iridium catalyst to form warm (1000° F. to 1500° F.) gases. Hydrazine is undesirable because of its toxicity and high melting point (34° F.). The tertiary amine azides, which are defined hereinabove and below, are non-carcinogenic alternatives to hydrazine in liquid or gel gas generator systems. These tertiary amines azides are non-carcinogenic alternatives for use with a thermal reactor bed where exothermic reaction releases enough heat to sustain decomposition for furnishing gases for gas generator systems employed. A tertiary amine typically has three hydrocarbons moieties attached to the nitrogen atom. The tertiary amine azides of this invention can have no more than seven carbon atoms in the molecules.

Abstract: In order to improve a method of injecting a first and a second fuel component into a combustion chamber, particularly of a rocket engine, wherein an injection element guides the first fuel component in an inner cylindrical stream and the injection element guides the second fuel component in an annular stream surrounding the inner cylindrical stream, so that by comparison with known methods improved mixing and more homogeneous preparation of the fuel takes place on injection into the combustion chamber, it is proposed that in the injection element the first and/or the second fuel component is guided by an element which produces a pressure drop and is disposed and constructed in such a way that the energy released during the pressure drop is at least partially converted into turbulence energy of the stream of the first and/or the second fuel component in order to achieve a good intermixing of the two fuel components in the mixing zone.

Abstract: Inhibited Red Fuming Nitric Acid (IRFNA) type IIIB and monomethyl hydrazine (MMH) ignite when contacted with each other because of a hypergolic chemical reaction and are the preferred oxidizer and fuel for bipropellant rocket propulsion systems. These propellants can deliver a specific impulse of 284 lbf sec/Ibm and density impulse of 13.36 lbf sec/cubic inch when the engine operating pressure is 2000 psi. Special precautions must be used when handling because of its toxic properties. A fuel gel propellant fuel that would be a suitable replacement for MMH must be less toxic and have a competitive density impulse for the same engine operating conditions. Three compounds meeting the specified requirements have been synthesized and their physical and ballistic properties are evaluated herein as shown in Table 1. The chemical names for these compounds are dimethylaminoethylazide (DMAZ), pyrollidinylethylazide (PYAZ), and bis (ethyl azide)methylamine (BAZ).

Abstract: The non-toxic bipropellent of the present invention contains a non-toxic ergolic miscible fuel (NHMF) and a rocket grade hydrogen peroxide. This non-toxic hypergolic miscible fuel (NHMF) has rapid ignition capability. The non-toxic hypergolic miscible fuel (NHMF) contains 3 species. Namely, a polar organic species miscible with hydrogen peroxide, a propagator, which may be substituted or unsubstituted amines, amides or diamines, and an inorganic metal salt, which reacts to form a catalyst in solution or as a colloid. The inorganic metal salt is miscible with the polar organic species and the propagator in solution. The catalyst has a faster rate of reaction with said rocket grade hydrogen peroxide than the propagator, the propagator has a faster rate of reaction with the rocket grade hydrogen peroxide than the polar organic species, and the polar organic species, propagator and catalyst are mutually soluble.

Abstract: A pulse detonation rocket engine, having at least two detonation chambers. The rocket propelled vehicle includes at least one fuel delivery system in fluid communication with each of the at least two detonation chambers, and at least one oxidant delivery system in fluid communication with the detonation chambers, along with fast-acting valves to inject fuel and oxidant controlledly into the chambers. An ignitor in each of the detonation chambers intermittently initiates detonation of a fuel and oxidant mixture in the chamber, in a controlled cycle, to provide motive force. Also provided is a combined cycle engine, able to operate in air breathing mode, oxidant augmented mode, and as a rocket engine. The combined cycle engine includes at least one detonation chamber, and may include a plurality of such chambers. The invention further provides methods of intermittently detonating sequentially created fuel and oxygen mixtures in these engines, and methods of using these engines.

Abstract: The invention is directed to an electrical ignitor for a model rocket, including two insulated lead wires, with each lead wire having an uninsulated end. An element wire interconnects the uninsulated ends from each lead wire. The element wire forms a plurality of turns around one of the insulated lead wires. A pyrogen compound surrounds the element wire.

Abstract: Tungsten is added to fuel gels to increase the density specific impulse. l gels contain monomethylhydrazine or other hypergolic liquids well known in the art. The quantity of tungsten employed can vary from 10%-98% weight percent depending on the specific application. The important parameters to consider during formulation are particle size, concentration, combustion efficiency, physical properties, and plume signature. Tungsten particle sizes ranging from 10 microns to 0.5 micron were compared with carbon of 0.24 when burned in air. It is shown that tungsten burns as well as or better than carbon; however, the increased density specific impulse achieved with tungsten as compared with carbon verifies that tungsten as a high energy additive to hypergolic fuel gels is superior.

Abstract: The multifuel combustion engine is a Lenoir cycle (constant volume), pulst combustion engine capable of operating on gasoline, diesel or kerosene based fuels. Although the preferred embodiment is described in terms of Lenoir cycle pulsejet engines, the present invention has application to combustion engines in general. The conventional Lenoir cycle engine has been modified to provide a direct, premixed fuel-air spray to the combustion chamber and means for igniting the fuel-air spray. Said fuel-air spray is separate and distinct from the fuel-air charge which is fed to the combustion chamber from the engine head. Additionally, means are provided for preheating the combustion chamber so that the same fuel-air ratio mixes can be fed to the combustion chamber for cold start or hot restart of the engine. A method of burning different fuels in combustion engines is also claimed. The present invention includes the application of the modified Lenoir cycle engines to smoke generator equipment.

Abstract: Current state-of-the art injection methodologies for bipropellant engines ly on either impingement or turbulence to mix the fuel and oxidizer streams. The multiple impinging stream vortex injector combines both mixing schemes into a single injector. Both first stage mixing at point of impingement and second stage mixing or turbulent vortex mixing is accomplished by impinging momentum balanced, tangentially injected propellant streams onto one another. The impingement angles are calculated to yield a resultant stream vector that consists of only a tangential velocity component. The two stages of mixing results in increased engine performance and combustion efficiency.

Abstract: A hypergolic fuel formulation which is consistently hypergolic with inhibd red fuming nitric acid is comprised of diethylethanolamine from about 44-72 weight percent, triethylamine from about 11-18 weight percent, and carbon from about 45-10 weight percent The formulation can be gelled with: silica, clays, carbons, or swellable polymers. The gellants can be combined with chemical agents that stabilize the gel under the standard 30 minute, 500 g centrifuge stability test. A preferred combination comprising diethylethanolamine in an amount of about 44 weight percent, triethylamine in an amount of about 11 weight percent, and carbon in an amount of about 45 weight percent when tested at an oxidizer/fuel ratio of about 4.25 reveals theoretical performance values of specific impulse (ISP) of about 250 at a chamber pressure of 1000 Psi and a density specific impulse (D* ISP) of about 350.

Abstract: The combustion of a hybrid engine is improved by continuously injecting into a precombustion chamber a hypergolic fluid such as triethyl aluminum which exothermically reacts with and vaporizes the oxidizer such as liquid oxygen. The prevaporized oxidizer evenly combusts the solid propellant grain to develop thrust.

Abstract: A bi-propellant self-contained propulsion system is provided for powering rockets. A plurality of turbopump assemblies are provided to take liquid propellants from low pressure storage tanks to a substantially higher pressure thrust chamber. Substantially all of the liquid propellants are pressurized and gasified within the plurality of turbopumps. Substantially all of the gasified propellants are then used to drive the turbopumps that pressurize the liquid propellants. Gasification preferably occurs within a preburner internal to the turbopump assembly that combines a small portion of one of the propellants with a substantial portion of the other. The proportions are selected so that gasification of all of the propellants is ensured yet relatively low preburner temperatures are maintained. A multi-stage pintle assembly may be provided to vary the exit-to-throat area ratio of the nozzle. The total thrust and the mixture ratio may be controlled by shutting down some of the turbopumps.

Abstract: A method and apparatus 10 for self-regulating the mass flow rate of a fluid comprises a housing 12 defining an upper chamber 14 and a lower chamber 16, a sealed and pressurized bellows 30 contained within the upper chamber 14, spring bias means 58 contained within the lower chamber 16 and a poppet 32 with a poppet head 40 for defining a fluid passageway 44.Supply fluid flows through inlet 56, lower chamber 16, fluid passageway 44 and into upper chamber 14. Once in upper chamber 14 the fluid is in heat exchange communication with a fluid in sealed bellows 30. When the mass flow rate of the fluid flowing through upper chamber 14 is effected by either in pressure or a change in temperature, a corresponding change takes place in the fluid in the bellows 30 thus repositioning the poppet head 40 accordingly to redefine the area of passageway 44. Fluid leaves the apparatus through exits 28.

Abstract: A compact, valve assembly (10) is provided having multiple chambers. A forward chamber (14) retains dual T-shaped blades (36) having shaped orifices (38) for controllably metering at least two fluids passing through the valve assembly. The valve assembly is also provided with a servo-valve assembly (70) and a piston indicator (62) assembly in cooperative association with the valve assembly.

Abstract: A new compound tris(2-azidoethyl)amine, N-(CH.sub.2 CH.sub.2 N.sub.3).sub.3, and its method of preparation is disclosed. The subject azido derivative of a tertiary amine has the empirical formula C.sub.6 H.sub.12 N.sub.10.

Abstract: Disclosed is a method for the prevention of the formation of ammonia as a product in the reaction of hydrazine and a halogen oxidizer used in a liquid-propelled rocket engine and in chemical laser pumping systems. The method employs vanadium pentafluoride as an additive up to about 1% to the halogen oxidizer to thereby shift the equilibrium of the reaction between the hydrazine and the halogen oxidizer so that no ammonia is produced.

Abstract: Disclosed is a method of converting non-hypergolic, liquid, rocket propellants into hypergolic propellants. The method to accomplish the conversion relates to the use of ammonium metavanadate as an additive to the liquid oxidant, such as, red fuming nitric acid (RFNA). The RFNA with additive is hypergolic with the usual fuels with which RFNA has been employed where a separate ignition system is normally required. RFNA with additive is also hypergolic with fuels which have been non-hypergolic and which have not been so used in the past with RFNA. These additional fuel blends include a fuel material selected from turpentine, aniline, triethylamine, furfuryl alcohol or blends of these fuel materials.

Abstract: Norbornadiene [bicyclo (2.2.1) heptadiene-2,5] is dimerized to a mixture of mainly the exo-endo stereoisomer of the hexacylic dimer of norbornadiene at both an excellent selectivity and conversion using an effective amount of a two component catalytic system of rhodium acetylacetonate and diethylaluminum chloride or ethylaluminum dichloride or aluminum ethylsesquichloride. The mixture also contains the endo-endo stereoisomer and some trimer. After hydrogenation a suitable mixture of the exo-endo and endo-endo dimers can be used as a component of high energy fuel.