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

An impeller is directly driven by an output shaft of a core engine. The airflow produced by the impeller rotates an air turbine and a fan disposed integrally with the air turbine. The impeller and the air turbine form a fluid coupling which serves also as a speed reducing mechanism. The rotational speed of the fan can be reduced to be lower than that of the output shaft while retaining efficiency of the core engine. The outer diameter of the fan can be increased, raising a bypass ratio.

To provide a turbofan jet engine which is capable of increasing the bypass ratio without increasing the fan diameter, and of reducing air resistance acting on the engine, a front fan duct that discharges air compressed by a front fan to the atmosphere and an aft fan duct that introduces air into an aft fan are disposed such as to change cross-sectional shapes thereof by rotating around a core engine in a counterclockwise direction, so that the cross-sectional geometric relationship between the front fan duct and the aft fan duct at a position immediately posterior to the front fan and a cross-sectional geometric relationship between the front fan duct and the aft fan duct at a position immediately anterior to the aft fan are inverted.

An aircraft gas turbine engine (110) comprises first and second non-coaxial propulsors (113a, 113b), each propulsor (113a, 113b) being driven by a common gas turbine engine core (176) comprising a propulsor drive turbine (143) arranged to drive the first and second propulsors (113a, 113b) via a propulsor drive coupling (127). The core (176) further comprises a first core module (190) comprising a first compressor (129) and a first turbine (131) interconnected by a first shaft (177), and a second core module (191) comprising a second compressor (128) and the propulsor drive turbine (143) interconnected by a second shaft (127), the first and second core modules (190, 191) being axially spaced.

A variable cycle propulsion system for a supersonic airplane, the system comprising at least one engine having means for producing exhaust gas and a gas exhaust nozzle for generating thrust for supersonic flight speeds, and at least one separate auxiliary propulsion assembly dissociated from said engine, having no gas generator, and capable of generating thrust for takeoff, landing, and subsonic flight speeds. The system further comprises gas flow tapping means movable between a position in which they tap off at least a fraction of the exhaust gas produced by said engine and feed it to said propulsion assembly to enable it to generate thrust for takeoff, landing, and subsonic cruising flight, and a position in which the gas produced by the engine is directed solely to the engine nozzle for supersonic cruising flight.

The invention relates to a variable cycle engine with a high bypass ratio. The variable cycle engine comprises a front fan (1), a booster stage (2), a high-pressure compressor (3), a combustor (4), a high pressure turbine (5), a low-pressure turbine (6) and a tail pipe nozzle (9), wherein the high-pressure compressor (3) and the booster stage (2) are respectively driven by the high pressure turbine (5), the low-pressure turbine (6) and the exhaust nozzle (9) through a high pressure shaft (8) and a low pressure shaft (9); a rear fan (10) is arranged between the booster stage (2) and the high-pressure compressor (3), the revolving speed of the rear fan (10) is between the revolving speed of the booster stage (2) and the revolving speed of the high-pressure compressor (3), and a plurality of first bypass valves (12) and corresponding bypass channels are arranged on the outlet of the rear fan (10). The air current matching of engine parts can be improved by additionally setting the rear fan between the booster stage and the high-pressure compressor, and bypass valves capable of being opened or closed are arranged on the outlet of the rear fan, so that the bypass ratio of the engine can be changed so as to be applied to the requirements of different power states.

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.

The invention provides a bypass ratio ultra-wide adjustable double-shaft turbofan engine structure based on a variable booster stage. The defects and shortage that the booster stage and a fan in an existing double-shaft turbofan engine are coaxial and have the same rotating speed, as the cruising height of a platform is higher and higher, the characteristic Reynolds number of the booster stage is continuously reduced due to the fact that the characteristic speed of the booster stage is low and the size of the booster stage is small, thus a boundary layer of the booster stage is separated, and the performance is sharply decayed are overcome; the booster stage is in driving connection with a high-pressure shaft through a variable reduction ratio gearbox, when an engine cruises at a low speed, the variable reduction ratio gearbox is set to be in a high speed reduction ratio, so that the rotating speed of the booster stage is relatively low, the total temperature and the total pressure of inlet airflow of a high-pressure compressor are relatively low, and the bypass ratio of the engine is relatively large; and when the engine cruises at a high speed, the variable reduction ratio gearbox is set to be at a low reduction ratio, so that the rotating speed of the booster stage is relatively high, the mass flow sucked by the high-pressure compressor is large, and the bypass ratio of the engine is relatively small.

A system and a method for assembling a core engine and a fan case of a gas turbine engine. The system includes a core engine support that detachably supports the core engine. The core engine support includes at least one core engine alignment sensor. The system further includes a fan case support that detachably supports the fan case. The fan case support includes at least one fan case alignment sensor. The system further includes a computer-controllable unit configured to align the core engine support and the fan case support based on signals from the at least one core engine alignment sensor and the at least one fan case alignment sensor. The computer-controllable unit is further configured to guide a movement of at least one of the core engine support and the fan case support to couple the core engine with the fan case.

The invention relates to a propulsive wing (1, 7) of an aircraft (10), comprising at least two structural spars, upstream and downstream (11, 12), that extend in a wing span direction of the wing (E-E), and a propulsion assembly (7) comprising at least two fans (9, 10) that are each driven by at least one gas generator (8).

According to the invention, at least one gas generator (8) extends along an axis that is remote from the axis of rotation of at least one fan (9, 10), and at least one of the spars (11, 12) comprises a first part and a second part (11a, 11b, 12a, 12b) between which the propulsion assembly (7) is arranged, the first and second parts being interconnected by a rigid structure (13) that is shaped so as to extend at least in part around said fans, the propulsion assembly (7) being supported by the rigid structure.

The invention provides a technical scheme of an intermittent inflating turbine engine. The type of engine mainly comprises a ring-body-shaped combustion chamber stator, a front functional end cover (rotor) and a rear functional end cover (rotor), wherein the ring-body-shaped combustion chamber stator is formed by a plurality of integral combustion chambers which are arranged into a ring and are same in shape and dimension, and the two functional end covers (rotors) rotate synchronously. During the operation process, when the front and rear parts of the combustion chamber are open, the combustion chamber is pressurized and ventilated, when oil mixing and ignition are carried out in the combustion chambers, the front end of the combustion chamber is sealed by the front end cover, thus a turbine blade behind the combustion chamber is pushed to rotate and do wok by utilizing high-pressure fuel gas efficiently; from the point of view of the whole engine, the engine pushes the turbine blade to rotate and do work by utilizing the high-pressure fuel gas continuously; from the point of view of certain combustion chamber, the combustion chamber has the characteristic of intermittent and cycle operation; the technical scheme of the intermittent inflating turbine engine is suitable for all occasions with easy flowing fuel materials which are combusted by virtue of air, and is widely applied to all mechanical motive power devices, is simple in structure, is low in operating temperature of the engine body, is short in engine body, is light in weight, is high in fuel oil power efficiency, and obtains high bypass ratio and thrust-weight ratio.

The invention discloses an external parallel fan turbine engine and application thereof in the field of flying wing aircrafts. The external parallel fan turbine engine comprises a main engine assembly, bypass fan assemblies and a bevel gear transmission assembly; the main engine assembly is a turboshaft engine/turboprop engine; the two or more sets of bypass fan assemblies are symmetrically and externally arranged on the two sides of the main engine assembly; a power turbine shaft of the main engine assembly and fan rotary shafts of the bypass fan assemblies are arranged in parallel; the transmission assembly comprises a main engine built-in transmission assembly and bypass fan built-in transmission assemblies; and the main engine assembly is connected with the bypass fan assemblies through the transmission assembly. Compared with conventional turbofan engines, the high bypass ratio design of the external parallel fan turbine engine can be achieved under a small engine height dimension, fuel consumption can be reduced, and the external parallel fan turbine engine is suitable for providing propulsion power for the flying wing aircrafts as blended wing bodies and is a power device with great potential for the new generation of passenger aircrafts with blended wing bodies.

An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition M_rel, wherein M_rel is between 0.4 and 0.93, and L/D is between 0.2 and 0.45.

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 (θ_scarf) 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 (θ_droop) relative to the longitudinal centre line. A relationship between the droop angle (θ_droop) and the scarf angle (θ_scarf) is given by: θ_droop=θ_scarf/1.5±1 degree.