Abstract: A hybrid propulsion system includes a housing, at least two electrodes, a solid-grain fuel material, a combustion chamber, an oxidizer port, and a nozzle. The housing has a first end and a second end and defines a cavity. The electrodes extend into the cavity. The fuel material is free of an oxidizer and is positioned in the cavity. The fuel material has a combustion surface and is exposed to the electrodes. The combustion chamber is defined between the combustion surface and the second end. The oxidizer port provides a flow of oxidizer to the combustion chamber. The nozzle is positioned at the second end. Combustion of the fuel material in the combustion chamber may be dominated by radiative heat transfer. Combustion of the fuel material in the combustion chamber may generate thrust of no more than 5 N at an oxidizer flow rate of no more than 5 g/s.

Abstract: The present invention provides a gas generator including a cylindrical housing in which an ignition device is attached at a first end portion thereof and a diffuser portion provided with a gas discharge port is attached at a second end portion thereof, an annular end surface being formed on the side of the second end portion, a cylindrical flow path forming member being disposed in the cylindrical housing, an inner space of the cylindrical flow path forming member being partitioned by a partition wall into a first chamber on the side of the first end portion and a second chamber on the side of the second end portion, a space, which includes the first chamber, being a combustion chamber filled with a gas generating agent, and the second chamber forming part of a gas discharge path, in the cylindrical flow path forming member, a first cylindrical space, which is formed between a first wall surface of the cylindrical flow path forming member on the side of the first chamber and the inner wall surface of the cyli

Abstract: Methods and systems for improving operation of an engine at higher speeds and loads are disclosed. In one example, fuel may be injected to an exhaust system of the engine so that temperatures of exhaust system components may be reduced when the engine is operated at higher speeds and loads.

Abstract: A spacecraft including at least one main propellant tank, a main engine fed with propellant by the main tank, and a de-orbiting device. The de-orbiting device includes a detonation engine fed with propellant by the main tank.

Abstract: The present invention relates to a cooling device (2) for cooling combustion gases from a rocket motor (1) in an antitank weapon. The cooling device (2) comprises a coolant reservoir (3, 21) containing a coolant (4). The coolant reservoir (3, 21) is arranged such that the coolant (4) is transferred from the coolant reservoir (3, 21) to the combustion gases in the gas outlets (6) of the rocket motor (1) in response to a pressure increase in the rocket motor (1). The coolant reservoir (3, 21) constitutes an integral part of the rocket motor (1) and is connected to the combustion chamber (7) by at least two gas inlets (8) for pressurization of the coolant (4). The coolant reservoir is furthermore connected to the rocket motor nozzle (9) by at least two coolant outlets (10) for transfer of coolant (4) from the coolant reservoir (3, 21) to the rocket motor nozzle (9).

Abstract: A pulse rocket motor has a hollow cylinder-like first grain situated in a rear section of a pressure vessel; a first igniter for igniting the first grain; a hollow cylinder-like second grain situated in a front section of the pressure vessel; a second igniter for igniting the second grain; and a dividing sheet member that covers the second grain within the pressure vessel. The dividing sheet member includes a dividing sheet expanding at least along the inner circumferential surface of the second grain and holders formed integrally with the dividing sheet at both ends of the dividing sheet by cure adhesion and attached to the pressure vessel. The dividing sheet includes a brittle portion that expands along the inner circumferential surface of the second grain through a rear end face of the second grain and is more brittle than other portions.

Abstract: An active composition for a decoy which radiates spectrally as the active composition burns, contains an additive which is distributed in the active composition and which increases the ratio of the intensity of radiation emitted during combustion of the active composition in the wavelength range from 3.7 to 5.1 ?m to the intensity of radiation emitted during combustion of the active composition in the wavelength range from 1.9 to 2.3 ?m. The active composition contains a fuel having carbon atoms and hydrogen atoms, and an oxidizer for the fuel, containing oxygen atoms. The amount of the oxidizer being such that it is not sufficient for complete oxidation of the carbon, and the additive being a catalyst present in the form of particles that catalyzes a redox reaction.

Abstract: An active composition for a decoy which radiates spectrally as the active composition burns up, including radiation emitted during combustion of the active composition in the wavelength range from 3.7 to 5.1 ?m which is stronger by a factor of at least 15 than radiation emitted during combustion of the active composition in the wavelength range from 1.9 to 2.3 ?m. The active composition contains a nitrate ester and/or one nitrosamine as a fuel having carbon atoms, hydrogen atoms, and ammonium perchlorate as oxidizer. The amount of the ammonium perchlorate is such that it is not sufficient for complete oxidation of the fuel. The active composition contains either the nitrate ester in the form of a polymeric solid, or a binder, not more than 5 carbon atoms in the fuel being joined to one another by direct bonding, and the active composition containing substantially no carbon source containing elemental carbon.

Abstract: A multi-pulse gas generator includes a pressure vessel, an outer propellant arranged in the pressure vessel and which has a tubular shape, an intermediate propellant arranged inside the outer propellant and which has a tubular shape, an inner propellant arranged inside the intermediate propellant and which has a tubular shape, an internal structure arranged inside the inner propellant and fixed to the pressure vessel, a first barrier membrane arranged between the outer propellant and the intermediate propellant so as to isolate the outer propellant and the intermediate propellant from each other, and a second barrier membrane arranged between the intermediate propellant and the inner propellant so as to isolate the intermediate propellant and the inner propellant from each other. The outer propellant is supported on its outer surface by the pressure vessel. The inner propellant is supported on its inner surface by the internal structure.

Abstract: The method serves to calculate a start sequence or a stop sequence for an engine, said sequence comprising a plurality of cues at which predetermined actions should be performed in the control of said engine. The method comprises: a step of obtaining at least one thermodynamic condition of said engine; and a step of calculating said sequence as a function of: said at least one thermodynamic condition; dimensional parameters of said engine; and criteria for proper operation of said engine.

Abstract: A pulse detonation device for dividing a pulse detonation shock wave into an primary and control portion to reduce the strength of a propagating shock wave and/or change its direction. The device contains a flow separator which directs a portion of the shock wave into itself, thus reducing the shock wave's strength. In one configuration, the control region converges in cross-sectional area so as to accelerate the flow in the control region, while the primary region diverges to slow the flow in the primary region. The flow in the control region is directed, at an angle, into the flow of the primary region to impede and/or redirect the flow of the primary region.

Abstract: A supersonic combustion apparatus and method of using the same including a side wall cavity having an enhanced mixing system with ground-based oxygen injection for hypersonic material and engine testing.

Abstract: An apparatus and method are provided for decomposition of a propellant. The propellant includes an ionic salt and an additional fuel. Means are provided for decomposing a major portion of the ionic salt. Means are provided for combusting the additional fuel and decomposition products of the ionic salt.

Abstract: A pulse combustion device has a number of combustors with upstream bodies and downstream nozzles. Coupling conduits provide communication between the combustors. For each given combustor this includes a first communication between a first location upstream of the nozzle thereof and a first location along the nozzle of another. There is second communication between a second location upstream of the nozzle and a second communication between a second location upstream of the nozzle of a second other combustor and a second nozzle location along the nozzle of the given combustor.

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 hybrid rocket motor is provided with a precombustion chamber supplied with propellant from separate fuel and oxidizer sources. The propellant can be in the form of gas or liquid and injected substantially tangentially into the head end of the hybrid motor adjacent the oxidizer injector to form a propellant swirl. As the hybrid motor oxidizer is injected into the swirl, it is heated and gasified, and assumes a swirling motion which increases the oxidizer path length and thereby increases the dwell time of the oxidizer. The increased dwell time increases combustion efficiency and permits multiple restarts of the hybrid motor. The propellant may also be a combination of solid and fluid reactants. In one embodiment, the oxidizer injector is extended into the combustion chamber to form a toroidal precombustion chamber which has an annular nozzle adjacent a face of the oxidizer injector.

Abstract: A hybrid rocket motor is provided with a precombustion chamber supplied with propellant from separate fuel and oxidizer sources. The propellant can be in the form of gas or liquid and injected substantially tangentially into the head end of the hybrid motor adjacent the oxidizer injector to form a propellant swirl. As the hybrid motor oxidizer is injected into the swirl, it is heated and gasified, and assumes a swirling motion which increases the oxidizer path length and thereby increases the dwell time of the oxidizer. The increased dwell time increases combustion efficiency and permits multiple restarts of the hybrid motor. The propellant may also be a combination of solid and fluid reactants. In one embodiment, the oxidizer injector is extended into the combustion chamber to form a toroidal precombustion chamber which has an annular nozzle adjacent a face of the oxidizer injector.

Abstract: A fuel spraybar assembly for spraying fuel within a gas turbine engine. The spraybar assembly includes radial and lateral members that distribute fuel within the flowpath. In one embodiment two lateral members are located at the radially inward end of a radial member and generally form a “T” shape. Circumferentially spaced adjacent spraybars subdivide the flowpath into a plurality of circumferential combustion zone segments. In one embodiment the junction of the radial and lateral members provides a flameholding feature that stabilizes the combustion flame. In another embodiment, fuel is introduced non-uniformly within the afterburner resulting in thermal vectoring of the engine thrust.

Abstract: Batch-mode and continuous-mode decomposition of nitrous oxide is used to provide multiple ignitions of a solid-propellant gas generator and subsequently control its output gas temperature and flow rate, respectively. To reignite the solid-propellant gas generator, a controlled mass of a reactive oxidizer, such as hot nitrous oxide decomposition products, is injected into the gas generator chamber.

Abstract: The present invention relates to a combined cycle system of enhanced efficiency. The system comprises a top stage, such as a fuel cell, a partial oxidation reactor or a heat engine, and an oxygen-enriching device, such as a temperature swing adsorption device or a chemical reactor bed device, as its bottom stage. The bottom stage uses waste heat produced by the top stage to enrich the oxygen content of air that is inputted to the bottom stage, thereby producing an oxygen-enriched gas mixture as the bottom stage output. This output mixture constitutes a superior oxidant which is fed back as an input for the top stage, thus enhancing the energy conversion efficiency, cheapness, and compactness of the combined cycle system as compared to that of ordinary fuel cells, partial oxidation reactors and heat engines that use unenriched air as their oxidant input.

Abstract: An apparatus for the dry treatment of a fabric having two enclosures for stocks of fabric, and means for the gradual in-feed and drawing out of the fabric from the enclosures. There are pneumatic means blowing air into a passage which comprises a first portion communicating one enclosure with a space and a second portion communicating another enclosure with the same space. A cylindrical drum is located in this space and adapted to rotate alternately in both directions, such that the fabric entering the space through one of the portions leaves it through the other portion, after partially wrapping the drum.

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: A method, device and composition for achieving ignition and sustained combustion using a solid consumable catalytic bed (CCB) is described. The CCB accomplishes this by decomposition of an injected fluid. Initially this reaction is catalytic or hypergolic in nature, eventually becoming entirely thermal as the CCB is consumed in the reaction generated. The CCB may be placed in a hybrid motor system such that the decomposed injected fluid will undergo combustion with the solid grain. When the injected fluid is high concentration hydrogen peroxide the catalytic materials of construction for the CCB are selected from the metallic hydride family, the manganese oxide family and the cuprocyanide family.

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 reduction in the emission of infrared radiation by the engine of a jet aircraft can be achieved by employing a low molecular weight monohydric alcohol, having a luminometer number of at least about 200 as at least a portion of the fuel. Additional reduction in emission of infrared radiation can be achieved by passing a low molecular weight monohydric alcohol in indirect heat exchange relationship with the hot exhaust nozzle prior to its introduction to the combustion chamber of the engine.

Abstract: The present invention relates to a general method for increasing the energy conversion from metal-containing rocket and ramjet propellants. According to the invention, this is achieved by means of combining the combustion of the fuel with exothermic intermetallic alloying reactions between components incorporated in the fuel.The invention also concerns rocket and ramjet propellants formulated in accordance with the abovementioned method.

Abstract: A ramjet solid fuel composed of Hydroxyl terminated polybutadiene aluminum, agnesium, and boron carbide is described. The high volumetric heating value fuel of the present invention significantly increases the distance range of missiles.

Abstract: A composite solid propellant with a stable burning rate comprising an oxidizer, a binder system of telomeric polybutadiene or copolymers of butadiene and acrylonitrile with terminal functional groups or functional groups distributed statistically along the chain, which are hardened by means of corresponding hardeners into rubber-elastic binders, finely pulverized, readily oxidizable metals and optionally plasticizers and burning rate moderators, wherein one or more of said metals is admixed with inorganic fluorides and incorporated into the composite solid propellant, preferably said metals and said inorganic fluorides are admixed in the form of an agglomerate having a particle size of between 100 .mu.m and 2,000 .mu.m.

Abstract: A composite solid propellant with a stable burning rate comprising ammonium perchlorate, a binder system of telomeric polybutadiene or copolymers of butadiene and acrylonitrile with terminal functional groups or functional groups distributed statistically along the chain, which are hardened by means of corresponding hardeners into rubber-elastic binders, finely pulverized, readily oxidizable metals and, optionally, inorganic fluorides, plasticizers and burning rate moderators, wherein ammonium perchlorate together with one or more of said finely pulverized metals, as well as, optionally, said inorganic fluorides are present as an agglomerate of larger particles having a particle size of between 100 .mu.m and 2,000 .mu.m.

Abstract: An autopyrolyzable solid composition for aerobic propulsion consists essentially of an oxidizing agent, a binder and a crosslinking agent. The oxidizing agent is a nitrated explosive material wherein the amount, by weight, of oxygen atoms fixed to the nitrogen atoms of the nitrate groups exceeds 40 percent relative to the total weight of the molecule. The binder is a polyalkadiene having functional terminal groups, such as --OH or --COOH groups, reactive with the crosslinking agent which can have isocyanate, epoxy or aziridinyl functions.

Abstract: A projectile is accelerated through a barrel by a high pressure hydrogen gas jet that is derived by exothermically reacting water or a water-hydrogen peroxide liquid mixture with metal or a metal hydride. The temperature of the reaction is controlled by controlling the power in a plasma discharge applied to a reaction chamber containing the liquid and particles of the metal or metal hydride. A non-vaporous metal oxide resulting from the reaction is centrifugally separated from the hydrogen that drives the projectile.

Abstract: A gas-generating agent for use in a ducted rocket engine essentially consists of an aliphatic polyester represented by the formula: ##STR1## or the formula: ##STR2## and a curing agent and optionally incorporating therein additives such as a curing catalyst.

Abstract: A thruster system has a main bipropellant propulsion unit including a propulsion engine (50) and bipropellant supply having a tank (3) for an oxidizing agent and a tank (4) for a fuel. An auxiliary cold propellant thruster (21) uses propellant supplied by bleeding off vapor from the contents of the oxidizing agent tank (3).

Abstract: Non-hazardous, highly soluble hydroxylammonium salt solutions are disclosed. These solutions have utility as liquid oxidizers in solid/liquid hybrid rocket motors.

Abstract: An IR source and method for generating IR radiation whereby a propellant is urned in a first chamber to produce a product gas which is exited through a critical exit, is accelerated to a supersonic velocity by expanding into a second chamber, is passed through a standing shock wave in the chamber to reduce the gas velocity to a subsonic level, is exited through radial orifices into a larger third chamber where the gas is mixed to obtain a substantial uniformity in temperature and specie, is accelerated by expansion to a chosen subsonic velocity, and is exhausted to the atmosphere.

Abstract: A solid ramjet fuel composition comprising a hydroxyl-terminated or carbo-terminated fluorocarbon binder, a polybutadiene prepolymer, a curative cross-linking compound, and a metal selected from the group consisting of aluminum, zirconium, amorphous boron or magnesium.

Abstract: A castable solid propellant composition suitable for use as a fuel generator for a ducted rocket motor is provided. The binder for the composition is a crosslinked polysiloxane. The polysiloxane polymer which is crosslinked to form the binder has a viscosity measured at 77.degree. F. of from about 12 poise to about 50 poise. The metal fuel for the composition is a mixture of spherical and flake magnesium. The composition contains from about 16% to about 20% polysiloxane binder and from 55% to 63% magnesium fuel. The oxidizer is principally ammonium perchlorate.

Abstract: The overall efficiency of a reaction engine having primary and secondary combustion chambers is greatly improved by reducing the primary exhaust velocity to subsonic speed. This is accomplished in a preferred embodiment by using an exhaust tube downstream of the primary nozzle. A tube diameter of about six times that of the primary throat diameter and a tube length of about five to ten times that of the tube diameter is considered optimum.

Abstract: A process for simulating the infrared radiation in wave length and intensity and the spatially extended characteristics of the plumes of turbojet engines comprising burning fuel-rich solid gas generator compositions, comprising an organic binder, an oxidizer, and dispersed inorganic particles, of a character such that venting of the combustion products directly into the atmosphere or venting them into the inadequate plume of a separate gas generator to provide a combined plume, produces a plume having the size and radiating characteristics of the turbojet plume being simulated.