31results about How to "Increase bypass ratio" patented technology

A turbofan engine of the invention is provided with a fan first-stage moving blade for taking an air therein, and a spinner rotationally driving the fan first-stage moving blade, the spinner has a spiral blade extending spirally to an outer side in a radial direction, sucking the air from a front face of the spinner and supplying the air to the fan first-stage moving blade. Further, the fan first-stage moving blade and the spinner are integrally coupled, and the spiral blade and the fan first-stage moving blade are formed such that blade surfaces are smoothly connected.

A hub-side leading edge part of a fan first-stage rotating blade 10 for taking an air thereinto more extends in a forward direction of an engine than a tip-side leading edge part and a mid-span leading edge part. The hub side of the fan first-stage rotating blade 10 is integrally connected as one with the tip side and the mid span while having a longer chord length than those of the tip side and the mid span. A radius at a root of the hub-side leading edge part is set in a boss ratio of 0 to 0.4.

A propulsive wing of an aircraft including at least two structural spars, upstream and downstream that extend in a wing span direction of the wing. A propulsion assembly includes at least two fans that are each driven by at least one gas generator. The at least one gas generator extends along an axis that is remote from the axis of rotation of at least one fan. At least one of the spars includes a first part and a second part between which the propulsion assembly is arranged. The first and second parts are interconnected by a rigid structure that is shaped to extend at least in part around the fans. The propulsion assembly is supported by the rigid structure.

An aircraft propulsion assembly comprising at least a main turbine mounted along a longitudinal axis, at least one main fan arranged upstream of the main turbine along the longitudinal axis and driven in rotation by the said main turbine, the said main fan being ducted by a main fan casing, an auxiliary turbine mounted along the longitudinal axis, the auxiliary turbine being independent of the main turbine, an auxiliary fan of axis offset with respect to the longitudinal axis and driven by the auxiliary turbine, the auxiliary fan being ducted by an auxiliary fan casing, the main casing being separate and distinct from the auxiliary casing so as respectively to generate a main secondary flow and an auxiliary secondary flow which remain independent of one another until they are discharged into the atmosphere.

A nacelle for a gas turbine engine having a longitudinal centre line includes an intake lip disposed at an upstream end of the nacelle. The intake lip includes a crown and a keel. The crown includes a crown leading edge and the keel includes a keel leading edge. The crown leading edge and the keel leading edge define a scarf line therebetween. The scarf line forms a scarf angle (θscarf) relative to a reference line perpendicular to the longitudinal centre line. A fan casing is disposed downstream of the intake lip and includes a casing leading edge. The casing leading edge defines a droop line normal to the casing leading edge. The droop line forms a droop angle (θdroop) relative to the longitudinal centre line. A relationship between the droop angle (θdroop) and the scarf angle (θscarf) is given by: θdroop=θscarf / 1.5±1 degree.

A platform of a fan outlet guide blade has a first rib and a second rib in an opposite surface of a flow passage surface thereof. The first rib and the second rib are formed in portions excluding a predetermined area separated to an upstream side and a downstream side by 8 to 30% of an inter-flange distance between a first flange and a second flange with respect to an axial center of the platform, the portions being located on a pressure surface side in the opposite surface.

The invention discloses a fan aerodynamic layout structure and method of an inlet turbofan engine with a non-full-blade high-pressure rotor, which belongs to the technical field of aero-engine fans. The blades of the rotor are pressed to pressurize the airflow near the fan inlet disc. By rationally designing the non-full-blade high-pressure rotor at the fan inlet, the circulation capacity of the fan can be effectively improved, and the internal pressure ratio of the fan can be effectively improved at the same time. The invention can be directly applied to a high-performance turbofan engine, and can reduce the weight, fuel consumption rate and noise level of the turbofan engine without increasing the external dimension of the engine.

The invention relates to a self-driven fan large-bypass-ratio turbofan engine with an inner loop air turbine and a design method. The self-driven fan large-bypass-ratio turbofan engine is characterized in that an outer bypass part is provided with a self-driven fan rotor with a larger bypass ratio and with an inner loop air turbine, the rotor comprises an auxiliary bypass fan rotor (1), an air-driving turbine rotor (3) and a rotating medium casing (5); an auxiliary bypass fan stator (2) is arranged on the downstream of the auxiliary bypass fan rotor (1); an air-driving turbine stator (4) is arranged on the downstream of the air-driving turbine rotor (3). The self-driven fan large-bypass-ratio turbofan engine has advantages: (1) while the bypass ratio is increased, the problem of the large-bypass-ratio turbofan that the rotation speed of the fan is not matched with that of the low voltage rotor can be avoided; (2) compared with the scheme of a gear turbofan engine (GTF) and a three-rotor turbofan engine for solving the mismatch problem of the rotation speed, the structure is simplified; (3) the self-driven fan large-bypass-ratio turbofan engine can be improved on the basis of the existing two-rotor turbofan engine, so that the technical risk and research cost can be reduced, and the research period can be shortened.

A fan of a turbine engine includes a disc provided with blades at the periphery thereof, the blades being mounted so as to pivot on the disc about a pivot axis, and a mechanism for changing the pitch of the blades. The mechanism is configured to adjust an angular position of each blade around the pivot axis. The angular position is in an angular setting range no greater than 20°.

The invention relates to an aero-engine core machine, a control method and an aero-engine, wherein the aero-engine core machine comprises a fan, a low-pressure compressor, a high-pressure compressor, a nozzle, a first movable part and a second movable part, and the low-pressure compressor It is arranged downstream of the fan; the first movable part is arranged between the fan and the low-pressure compressor, and is used to adjust the bypass ratio of the core engine of the aero-engine through its own motion; the high-pressure compressor is arranged downstream of the low-pressure compressor; the nozzle pipe It is arranged between the low-pressure compressor and the high-pressure compressor, and is used for exporting the outlet air flow of the low-pressure compressor; the second movable part is arranged at the inlet of the nozzle, and is used to open or close the inlet of the nozzle. The aero-engine includes the above-mentioned core engine. The present invention can realize the adaptability to different operating conditions of the engine through the cooperation of the first movable element and the second movable element.

The invention discloses an air-driven ducted fan jet propulsion power system, comprising a core turbine, a ducted fan connected to the core turbine through an air induction pipe, and a blade tip turbine coaxially installed with the ducted fan, the core turbine exhausts gas It is transmitted to the blade tip turbine, and the blade tip turbine works on the expansion of the gas, driving the bypass fan to rotate and inhale the air to generate thrust, which breaks through the design limitation of conventional turbofan engines with large bypass ratio, greatly increases the bypass ratio of the engine, and further exerts its large capacity. The basic principle of bypass ratio turbofan engine to improve propulsion efficiency, lower fuel consumption and higher power output, and the function of the induction pipe connected between the energy source core turbine and the bypass fan is only the airflow channel without complicated machinery The connection can be arranged inside the fuselage and wings according to the design needs of the aircraft, giving the aircraft a more flexible layout.

The invention discloses a small-sized turbofan engine. The small-sized turbofan engine comprises an air compressor, a combustion chamber and a power turbine, wherein the air compressor comprises an air compression housing and centrifugal impellers; the power turbine comprises a turbine body and a turbine housing; a connector is arranged in the combustion chamber; the connector is provided with an air compression end cap and a turbine end cap; communication holes are formed in the air compression end cap; a plurality of detonation flame tubes are arranged in the combustion chamber; an air inlet and an air outlet are formed in each detonation flame tube; the air outlets are communicated with the air inlet end of the turbine housing; a one-way valve is arranged at the air inlet of each detonation flame tube; a fuel nozzle and an ignition plug are arranged in each detonation flame tube; and the fuel nozzles are connected through a general fuel tube. According to the small-sized turbofan engine disclosed by the invention, through the sequential detonation of the detonation flame tubes, the turbine body is continuously pushed by utilizing four kinds of pushing force, and detonation force is accumulated by the detonation flame tubes; and under the action of the one-way valves, air successively enters the engine, so that the bypass ratio and the compression ratio of the engine are increased.

The invention provides a large-bypass-ratio turbofan engine of a vane tip jet driven fan and a designing method. A large-bypass-ratio fan rotor driven by vane tip jet is arranged on an external duct portion and comprises an additional ducted fan rotor, a vane tip driving portion and a rotary additional duct rotor hub; an additional duct stator is arranged on a downstream portion of the additional ducted fan rotor and a downstream portion of a vane tip driving portion; and air is led to vane tips of the additional ducted fan rotor from an external duct of the original double-rotor turbofan engine by an air leading channel. The large-bypass-ratio turbofan engine of the vane tip jet driven fan has the advantages that the bypass ratio is increased, torque transmission is implemented by an aerodynamic force, and the problem that the rotation speed of the fan is not matched with that of the driving component; the peripheral speed is greatly improved, and the improvement potential of the bypass ratio is high; compared with the prior art, the large-bypass-ratio turbofan engine is simple in structure; and the large-bypass-ratio turbofan engine is improved on the basis of the existing double-rotor turbofan engine, development risk and the development cost of the large-bypass-ratio turbofan engine of the vane tip jet driven fan are reduced, and the development cycle is shortened.

The coating composition which coats the biologically active substance includes at least one protective material selected from the group consisting of a hardened animal fat, a hardened vegetable oil, a linear or branched, saturated or unsaturated aliphatic monocarboxylic acid having 12 to 22 carbon atoms, a fatty acid ester, and a wax group; lecithin; and at least one preservative selected from a propionic acid or its salt, a sorbic acid or its salt, a benzoic acid or its salt, a dehydroacetic acid or its salt, parahydroxybenzoic acid esters, an imazalil, a thiabendazole, an orthophenyl phenol, an orthophenyl phenol natrium, and a diphenyl.