Abstract: A propulsion system is provided for an aircraft. This propulsion system includes a turbine engine, and the turbine engine includes a propulsor rotor, an engine core and a heat exchanger. The propulsor rotor is configured to rotate about an axis and direct a first air stream along a primary flowpath that bypasses the engine core. The propulsor rotor is configured to direct a second air stream along a secondary flowpath with a geometry extending at least eighty degrees about the axis. The turbine engine is configured such that the first air stream enters the primary flowpath at a first pressure and the second air stream enters the secondary flowpath at a second pressure that is different than the first pressure. The engine core is configured to drive rotation of the propulsor rotor. The heat exchanger is disposed within the secondary flowpath.

Abstract: An aircraft turbojet engine extending along an X axis and comprising a blower configured to provide a reverse thrust and a nacelle comprising an air intake which comprises at least one deflection member movably mounted between a deployed position in which the deflection member projects from the inner wall or from the lip of the air intake in a radially inward direction of deployment facing the X axis or in a longitudinal direction of deployment with respect to the X axis, in order to allow a release of the reverse air flow from the inner wall to support the reverse thrust phase, and a retracted position in which the air intake has an aerodynamic profile so as to guide the internal air flow along the inner wall in order to support the thrust phase.

Abstract: Disclosed is a method for using an aircraft turbojet engine comprising an air inlet comprising a plurality of rectifier vanes, each rectifier vane being mounted such that it can move between a retracted position to assist the thrust phase and a deployed position in which the rectifier vane protrudes from the inner wall in a radially inward direction in order to rectify the reverse air flow of the inner wall to assist a thrust-reverse phase, in which method at least one rectifier vane is in the retracted position during a turbojet engine thrust phase, the method comprising, during a thrust-reverse phase of the turbojet engine, a step of moving the rectifier vane to the deployed position.

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: A drive aggregate is operated independently from the availability of external air, by using a fuel being composed of a blend of hydrocarbons and metal, e.g. lithium hydrides in a pasty, pumpable, i.e., fluidized, condition, and a metal chlorate, e.g., lithium chlorate or metal perchlorate is used as an oxidizer, having been rendered fluid in a high concentration by the addition of water so that the oxidizer, when decomposing, yields oxygen which combines with hydrogen released from the fuel. Reaction products remain, or are pumped into the tanks or into the ambient sea.

Abstract: A multi component propulsion system utilizing metallic fuel particles and a carrier gas as a propellant medium. The metallic particles and some of the carrier gas are carried from a fluidization chamber to the combustion chamber to be combusted with a liquid oxidizer. A major portion of the carrier gas is passed from the fluidization chamber to the combustion chamber to mix with the products of combustion of the oxidizer and metallic particles.

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 fuel for rocket motors is provided in the form of a slurry of particulate metal or metalloid particles having a particle size of about 0.1 to 10 microns suspended in an alcohol having from 1 to 3 carbon atoms in the presence of a non-thixotropic amount of high viscosity grade hydroxypropyl cellulose.

Abstract: High-pressure gas generating apparatus includes a combustion chamber and a water jacket surrounding the combustion chamber. A metal fuel injecting nozzle is disposed in the center of one end of the combustion chamber. The water in the jacket is injected into the combustion chamber in the form of a swirling stream around the fuel injected from the nozzle.