257results about How to "Increase thrust-to-weight ratio" patented technology

The present invention is a rotary device that may be adapted for use as a propeller assembly and electrical generator for aerial vehicles or other vehicles intended for fluid media. In one example, the device includes a ring assembly having a plurality of centrally linked blades coupled to a rotatable common hub. Rotary motion of the ring assembly is facilitated by coupling it to an opposed cylinder, opposed piston, internal combustion. The ring assembly includes components of an electrical power generating system so that electrical power is produced from the rotation of the ring assembly.

Supersonic Magnetic Advanced Generation Jet Electric Turbine (S-MAGJET) described herein, and a subsonic derivative, MAGJET, integrate a gas power turbine, superconducting electric power and propulsion generation, and magnetic power flux field systems along with an ion plasma annular injection combustor which utilizes alternative petroleum-based fuel and combustion cycles to create a hybrid turbine turbomachine for aerospace propulsion. The propulsion unit is able to achieve a dramatic increase in horsepower, combustion and propulsion efficiency, and weight reduction. In addition, the turbomachinery structures may be disposed within an exo-skeleton architecture that achieves an increase in thrust to weight ratio with a concomitant increase in fuel efficiency and power generation over traditional gas turbine technology today. The engine continuously adjusts the temperature, pressure and mass airflow requirements using an electromagnetic power management system architecture. Engine performance may be controlled across the entire desired flight envelope, whether subsonic, transonic or supersonic flight conditions. With bypass fan(s), compressor stages and turbine segments controlled electrically in a shaftless turbine design, the S-MAGJET maximizes the propulsion efficiencies over a broader range of operating conditions compared to current art of turbine technology and at much higher thermodynamic and aerodynamic efficiencies.

The invention discloses a skew line tooth surface gear transmission pair and a tooth width geometric design method. The skew line tooth surface gear transmission pair adopts an alternating axis transmission form composed of involute spur gears and skew line tooth surface gears. Each skew line tooth surface gear is formed by spreading straight tooth involute slotting cutters in a staggered mode; the skew line tooth surface gears can only mesh with the straight tooth cylindrical gears and can not mesh with the skewed tooth cylindrical gears with helix angle; and the tooth trace is approximate to a skew line and forms an oblique angle with the radius direction, thereby being very suitable for the demands of design diversification for aviation compact space. The method illuminates the gear shaping principle of the skew line gear and skew line tooth surface gear, and deduces the tooth surface equation of the skew line tooth surface gear. The margin line of the involute slotting cutters is utilized to calculate the position of the tangent point of the internal end tooth root. The condition that the outer end tooth top tooth thickness is equal to zero is utilized to obtain the sharpening condition. The condition of avoiding secondary cutting of the tooth root is combined with determine the cutter top fillet radius, thereby finally obtaining the tooth width of the skew line tooth surface gear.

The invention discloses a solid rocket ramjet based on detonation combustion. The solid rocket ramjet comprises an outer shell, an inner core, an air inlet channel, an isolation section channel, a detonation combustion channel, a gas generator, a flow stabilizing ring cavity, a main fuel injection system, a rotary detonation combustion chamber, a tail spray pipe and other structures, wherein the air inlet channel, the isolation section channel and the detonation combustion channel are formed between the inner wall of the outer shell and the outer wall of the inner core; air captured by the air inlet channel is decelerated and pressurized in an isolation section and then enters the rotary detonation combustion chamber; the front portion of the inner core is of a hollow structure, the gas generator is arranged in the inner core, an oxygen-deficient solid propellant is combusted in the gas generator and forms fuel-rich gas, and the fuel-rich gas enters an expansion section of the isolation section channel through a main fuel injector, is fully mixed with fresh air and then is sprayed into the ring cavity of the rotary detonation combustion chamber; and the rotary detonation combustion chamber is provided with an ignition device, the tail spray pipe is installed at an outlet, chemical energy is converted into kinetic energy, and thrust is generated. The solid rocket ramjet has higher combustion efficiency and work performance.

The invention discloses an aero-engine aerodynamic stability active composite control method based on stability estimation and prediction. The active composite control method comprises (1) a gas compressor pneumatic instability mechanism and an active control method, (2) an aero-engine stability margin estimation and active stability control method, (3) an aero-engine instability prediction and active anti-surge control method, and (4) an aero-engine active stability composite control method. Aiming at the aero-engine aerodynamic stability problem, the aero-engine aerodynamic stability is improved from three different aspects of gas compressor part stability extension, residue stability margin utilization and active anti-surge, an adjustment and control strategy and algorithm with the surge margin dynamically adjusted and controlled along with the application requirements and suitable for the aero-engine active stability control feature and engineering application background is obtained, and the actual use requirements are met.

The invention discloses a contra-rotating outer framework water-air dual-purpose engine. The engine mainly structurally comprises a supporting frame, an oil electric pipeline, a front thrust bearing, a peripheral bearing, a radome, an inner duct, an outer duct, a gas compressor, a ring-shaped combustion chamber, a worm wheel, a rear thrust bearing, a convergent tail spray tube, a central duct, an inner rotor, an outer rotor, a blade tongue-and-groove, a fan, an engine cover, a power transmission circuit, a conductive slide ring, an oil conveying pipeline, a convergent channel, a rotary joint, a metal fuel cavity, a metal-water reaction combustion chamber and a fan stator blade. The gas compressor and the worm wheel of the engine are not provided with a stator blade, and an outer frame rotor and an inner rotor are in counter rotating in an air working state, so that the gas compressor and the worm wheel can realize counter rotating, the absolute rotation speed of the rotor is lowered when the engine efficiency is improved, the axial dimension of the engine is shortened and the structure weight is reduced. When the engine works in water, the ring-shaped combustion chamber does not work, the inner and outer rotors are static, and power is produced by the metal-water reaction combustion chamber, so that water-spraying propulsion is realized.

The invention discloses a secondary-flow evaporation tube type combustion chamber structure. A combustion chamber body and a baffle plate are adopted, wherein the front end of the combustion chamber body is shaped like a cylinder, and the upper end of the baffle plate is provided with an annular groove; the baffle plate is additionally arranged on an outer wall surface of a combustion chamber, andair film holes are formed in the annular cylindrical barrel wall and the groove wall of the annular groove, so that the reverse air flow of a secondary flow in a casing is heated and heated again after being injected by a primary air flow at the head of the combustion chamber and mixed; and then the combustion flame temperature field is more uniform, the heat transfer of the fin reinforced wall surface enables the temperature of the wall surface to be effectively decreased, the uniformity of the flow field is optimized, the combustion efficiency is improved, and the emission of pollutants such as NOx is greatly reduced while fuel is saved. The annular cylindrical outer wall and the inner wall of the annular groove are connected through fins to form an integrated forming structure, the overall weight of the device is reduced, the outline size of the combustion chamber is decreased, the thrust ratio of an engine is effectively increased, and the maneuverability of the engine is increased; and the using number of connecting parts between structural members is decreased at the same time, and the overall performance is improved.

A miniature turboject engine with 10-30 kg of propulsive force is composed of air inlet channel, compressor, combustion chamber and exhaust system. The front and back nuts, rotor of centrifugal compressor, oil thrower and turbo rotor are installed on the shaft of engine. It has also the flame barrel, tail jet nozzle, shaft couple and DC motor. Its advantages are simple structure, small size and light weight.

The invention relates to the technical field of precision casting, in particular to a method for making a guide vane inner ring by opening a process window. The method comprises the following steps of: designing a computer model of the guide vane inner ring by a UG solid modeling method; firing with a laser rapid forming machine; performing combination, sticking and wax-welding on a fixture with a fixing tenon; performing size detection and local trimming on the wax piece to finish the manufacturing of a wax mould; assembling the wax mould into a pouring system; smearing a mixed coating consisting of silica sol and fused corundum on the surface of the wax mould; dewaxing and sintering to obtain a shell of the pouring system; pouring on a vacuum furnace to obtain an integral mother casting of the guide vane inner ring; opening 14 process windows in 15 mm*20 mm on the annular thin wall of the integral mother casting of the guide vane inner ring; welding 15 through holes of the outer ring and 35 nozzle holes of the inner ring on the integral mother casting of the inner ring; and welding a 15 mm*20 mm blocker on the process window to obtain the final product.

The invention provides an afterburner and a turbine engine. The afterburner comprises an outer cylinder assembly, a center body assembly, an oil way assembly, an annular combustion chamber and an igniter; the center body assembly comprises a front body, a middle body and a rear body, wherein the front body is located at the axial front end of the afterburner and forms a gas inlet, accepting entry of oxygen-containing gas, of the annular combustion chamber with the outer cylinder assembly, the middle body is located behind the front body along the axis and is connected to the front body and integrated with the front body, and the rear body is located behind the middle body in the axial direction, is arranged in the middle body in a slid and sleeved mode and forms a gas outlet of the annular combustion chamber with the outer cylinder assembly; the igniter is arranged at the corresponding part, forming the annular combustion chamber with the outer cylinder assembly, of the center body assembly and is used for igniting fuel gas formed by fuel and the oxygen-containing gas entering the annular combustion chamber, then rotation and knocking combustion of the fuel gas are realized, and therefore the problem that combustion of the afterburner is instable is solved, the heat efficiency of combustion is improved, and the power capability of working media is improved.

Based on an existing double-layer shell type impact/gas film/turbulent flow composite cooling structure, the invention discloses a gas-liquid coupling turbine blade cooling unit. The gas-liquid coupling turbine blade cooling unit comprises a plurality of mutually independent cooling subunits. Each cooling subunit comprises a cold gas wall, a hot gas wall, a plurality of impact holes formed in the cold gas wall, a plurality of gas film holes formed in the hot gas wall, a plurality of turbulent flow columns located between the cold gas wall and the hot gas wall, and a collecting table connected with the cold gas wall and the turbulent flow columns, wherein the cold gas wall and the hot gas wall are arranged oppositely, the upper surface of the collecting table is connected with the inner surface of the cold gas wall, the lower surface of the collecting table is connected with the turbulent flow columns, and at least one cold source pipeline penetrating through the collecting table longitudinally is arranged in the collecting table. Further, in the cooling subunits, the side walls of the collecting tables are smoothly connected with the inner surfaces of the cold gas walls and the side walls of the turbulent flow columns to form a smooth hook face structure together. Through the gas-liquid coupling turbine blade cooling unit, a liquid cooling path is added, and the overall temperature of a turbine is further lowered.

The invention relates to a method for inhibiting flow separation in a transition section of a compression system of an engine of a low-bypass aerial turbofan by using a guide vane. The method comprises the following steps: placing the guide vane at a first curved conduit on the inner wall of a transition section of an air compressor, and forming a vane force with gradient opposite to that of the prior pressure at the curved conduit of the transition section through changing a peripheral included angle of a laminated line and a radial line of the guide vane, so as to reduce pressure gradient of the flow direction at the curved conduit, and inhibit the flow separation in the transition section. Because the method can inhibit the flow separation in the transition section, the method can increase bending degree of the transition section, so as to reduce the length of the transition section. At the same time, because the guide vane is placed in the transition section, the method saves space occupied by the prior guide vane, and is favorable for reducing the length of the whole air compressor. The length of the transition section is estimated to be shortened by about 20 percent by adopting the method and the method contributes to the thrust mass ratio of the engine by about 1 percent.