445 results about "Turbojet" patented technology

The composite turbomachine blade of the present invention comprises a preform made of yarns or fibers woven in three dimensions and a binder maintaining the relative disposition between the yarns of the preform. Said preform is made up of warp yarns and of weft yarns, the direction of the warp yarns forming the longitudinal direction of the preform. In characteristic manner, said preform comprises at least a first portion made using a first weave forming the airfoil of the blade, and a second portion made using a second weave forming the root of the blade, and the first and second portions are united by a transition zone in which the first weave is progressively modified to end up with the second weave, thereby obtaining a reduction at least in the thickness of the blade between the second portion and the first portion. The invention is applicable to making a fan blade for a turbojet.

The bay (12) of a turbofan engine (10) comprises a front structural element (30), whose external surface is continuous and extends over at least 50% of the geometrical chord of the bay. Said element (30) is installed on maintaining and guiding members (44), such as slides, which prevent a significant deformation in flight and allow a sliding to the front of the element (30) for maintenance purposes. A laminar air flow around the front half of the bay (12) is consequently ensured.

twin-flow, double body turbojet includes a fan that is positioned upstream from a gas generator and delimits primary and secondary flows. The gas generator is traversed by the primary flow and includes a low-pressure compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine and a low-pressure turbine. The low-pressure turbine is linked to the low-pressure compressor by a low-pressure rotating shaft, and the high-pressure turbine is linked to the high-pressure compressor by a high-pressure rotating shaft. The turbojet has an electric motor for injecting mechanical power into at least one of the rotating shafts. The turbojet also has a device for removing power from at least one of the rotating shafts and transforming the excess power into electrical energy. An electric storage means is positioned between the device for removing power and the electric motor.

A mixer for a separate-stream turbojet nozzle, the mixer comprising along a longitudinal axis (X-X′) both a fastener shroud (10) for connecting said mixer to the exhaust casing of the nozzle, and a lobe structure (20) presenting a succession of inner and outer lobes (22 and 21) distributed circumferentially around the longitudinal axis (X-X′) of the mixer. The lobe structure (20) is made of a ceramic matrix composite material and further comprises a stiffener ring (30) that forms a connection between at least some of the lobes of said structure. Thus, the major portion of the mixer, i.e. the lobe structure, is made of a ceramic matrix composite material, thereby significantly reducing the weight of the mixer, and consequently reducing the weight of the nozzle.

The invention relates to a high-speed high-temperature multifunctional friction and abrasion tester which is especially applicable to simulating the service conditions of the sealed pairs of turbojet engines and solves the problems that the friction pairs in the prior art can not meet the needs of high speed (over 100m/s) and high temperature (up to 1200DEG C) in the aspects of motion line speed and experiment temperature and can not simulate the high-speed high-temperature friction and abrasion conditions of the gas path sealing of the turbojet engines and the like. The tester is provided with a rotation main shaft, a rotation disc, a numerical control sliding table and heating devices, wherein the rotation disc is arranged on the rotation main shaft, a rotation sample is arranged on therotation disc, the numerical control sliding table is arranged at one side of the rotation main shaft, a translation sample is arranged on the numerical control sliding table and is opposite to the rotation sample to form a friction pair; and the heating devices are arranged at the side of the rotation sample and at the side of the translation sample. By adopting the invention, the high-speed stable rotation of the rotation main shaft of the tester and the rapid heating of surface of the friction pair sample can be realized.

The invention relates to a thrust inverter with grids (5) for a turbojet, that defines at least a portion of the downstream section (103) of a nacelle (100) housing the turbojet, comprising a front frame (105) to be secured to a fixed portion (102) of the nacelle and bearing, on the one side, cylinders (6) for the translational actuation of at least one mobile cowling (2) and, on the other side, grids for inverting the peripheral thrusts, characterised in that the front frame includes a fixed portion for securing the front frame to the fixed portion of the nacelle and bearing the diversion grids, and a portion mounted so as to be capable of translation movement along the longitudinal axis of the nacelle and on which one end of the cowling actuation cylinders is attached, wherein the fixed and mobile portions of the front frame can be removably connected to each other by locking means.

A single-stage hypersonic vehicle is comprised of a low-speed and a high-speed propulsion system. The low-speed propulsion system propels the single-stage vehicle to a threshold velocity, after which the high-speed propulsion system then takes over. The low-speed propulsion system includes a combined-cycle engine featuring a swirl generator that is integrated into a turbojet engine to provide a compact turbojet and swirl afterburner-ramjet propulsion system. The high-speed propulsion system includes a hypersonic engine that is operable at the threshold takeover velocity and beyond. In various embodiments, the high-speed propulsion system comprises a scramjet, rocket, or scramjet/rocket engine depending requirements. Benefits of the swirl generator design include its ability to rapidly and efficiently atomize, vaporize, mix and burn the fuel and oxidizer for the low speed propulsion system, significantly reduce length, weight, cooling requirements and complexity for both propulsion systems, while maintaining high propulsion performance and reducing propulsion and launch costs.

The modular aircraft system includes a single fuselage having a permanently installed empennage and plural sets of wing modules and engine modules, with each wing and engine module optimized for different flight conditions and missions. The fuselage and each of the modules are configured for rapid removal and installation of the modules to minimize downtime for the aircraft. Short wings having relatively low aspect ratio are provided for relatively high speed flight when great endurance and/or weight carrying capacity are not of great concern. Long wings having high aspect ratio are provided for longer range and endurance flights where speed is not absolutely vital. A medium span wing module is also provided. Turboprop, single turbojet, and dual turbojet engine modules are provided for installation depending upon mission requirements for any given flight. The aircraft is primarily adapted for use as an autonomously operated or remotely operated unmanned aerial vehicle.

The invention concerns a method for making a soundproofing panel with at least one double resonator used for example on pods enclosing aircraft turbojets. Said method is characterized in that the septum is formed by assembling a plurality of parts against one of the honeycomb cores, said parts being cut out in foil, said parts being defined to enable the shape of the septum to be more or less adapted to extensible surfaces, said parts enabling significant reduction of the cost for perforating the septum.

Propulsion from combustion of solid propellant pellet-projectiles for providing a useful propulsion that has the advantages of both solid and liquid propulsion engines, and also can make use of either solid chemical propellants or fissionable nuclear material as the fuel. Preferred methods and systems can include a storage chamber for storing solid propellant pellets, a feeding system having a pellet feeding channel and a pellet feeding mechanism connected to the storage chamber and to a gun assembly, which is positioned along a longitudinal axis to eject the pellets at a certain velocity. A triggering system positioned between gun assembly and thrust chamber can initiate the propellant pellet-projectile, and a thrust chamber having a combustion chamber for combustion of propellant pellet-projectile with an exhaust nozzle. Additionally, an auxiliary power system can be used to power the components and various electrical and electronic systems that may be present in the invention for controlling the engine components. The gun assembly may include an ejector mechanism for ejecting the propellant pellets through at least one barrel. The triggering system can produce a medium creating an ambience for the initiation of propellant pellet-projectiles. Methods and systems can be used for space and rocket crafts, turbojets and ramjets.

Propulsion system with integrated pylon Propulsion system comprising a bypass turbojet and a nacelle (10) a downstream cylindrical part (20) of which is rigid and is attached at its upstream end to an intermediate case of the engine, this downstream cylindrical part (20) also comprising a longitudinal beam (26) for attaching rods (28) for securing the engine to part (22) of an airplane.

A device for supporting and housing auxiliaries in a bypass turbojet, comprising two coaxial case rings arranged one inside the other and connected by tubular radial arms through which fluid ducts and electric cables pass, at least one of the radial arms comprising on a lateral face a detachable panel whose removal allows access to turbojet equipment situated radially inside the internal case ring in alignment with the radial arm.

The invention relates to a motor fuel suitable for diesel, gas-turbine and jet engines and standard engines, comprising a mixture of organic compounds having oxygen-containing functional groups, and optionally a hydrocarbon fraction. The fuel is typically a stable homogeneous liquid at atmospheric pressure and normal ambient temperature and achieves a reduction of harmful pollutants in the exhaust emissions of the engines. A total of at least four different oxygen-containing functional groups are present in at least two different oxygen-containing organic compounds. The total concentration of organic compounds containing bound oxygen in the fuel composition generally varies from 5% and to 100% of the total volume of the fuel composition, and the concentration of the hydrocarbon compounds varies, correspondingly, from 95% to 0% of the total volume of the fuel composition.

The invention relates to a mixer (100) for a separate-stream nozzle of a turbojet, the mixer being designed to mix a hot inner stream from the combustion chamber of the turbojet with a cold outer stream from the fan of the turbojet. The mixer comprises an inner shroud (120) defining a flow channel for said hot inner stream, an outer shroud (110) disposed around the inner shroud (120) and co-operating therewith to define a flow channel for said cold outer stream, and a lobed structure (130) having lobes (1321, 1322) extending longitudinally from the trailing edges (110a, 120a) of said shrouds. The lobed structure (130) is made of a ceramic matrix composite material and is attached to the outer shroud (110) by flexible fastener tabs (140).

The invention provides a method of ventilating equipment of a turbojet, the turbojet (10) having a cold zone (ZF) including a fan (14) upstream from a hot zone (ZC), the turbojet equipment (32) for ventilating being arranged in a ventilation space (30) available around the hot zone between a primary air stream passage (26) and a secondary air stream passage (28) of the turbojet, the method comprising taking ventilation air from one of the air stream passages in order to convey it to the ventilation space, and providing forced flow of the ventilation air by means of at least one electric fan (36) positioned inside the ventilation space and powered electrically by at least one thermoelectric generator (38) suitable for inducing a potential difference (ΔV) and positioned in a wall (40) between the ventilation space and the secondary air stream passage.

The invention discloses a control device for a microminiature turbojet engine and a starting control method thereof. The device comprises a rotating speed sensor for collecting the rotating speed of an engine, a temperature sensor for collecting exhaust temperature, a switch controller, a throttle controller, an electronic control unit (ECU) and an executing mechanism, wherein the executing mechanism comprises a starting motor, a hot piston, a propane gas valve, a fuel protective valve and a fuel gear pump. The control method comprises the following steps: the ECU monitors the working state of the engine through the sensor, and outputs a pulse-width modulation signal to a power drive circuit to control the starting motor, the hot piston, the propane gas valve, the fuel protective valve and the fuel gear pump so as to complete starting control to the engine; functions of establishing fuel pressure and adjusting fuel quantity can be realized by the fuel gear pump after starting of the engine is completed; and fuel can be accurately supplied to ensure that the engine operates in an expected working state. The device can be used for a microminiature delivery vehicle, and the executing mechanism of the device has the advantages of simple structure, small volume and light weight, and easily realization of reliable and efficient control to the engine.

A system of attaching an injection system to a turbojet combustion chamber base. It comprises a deflector welded onto the chamber base. The deflector comprises an annular portion having an edge forming a retaining shoulder directed toward the front of the turbojet and the injection system comprises a collar on which is formed a retaining shoulder directed toward the rear of the turbojet and pressing against the retaining shoulder of the deflector.