6988results about "Efficient propulsion technologies" patented technology

A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The fan comprises a plurality of fan blades that define a fan diameter and a BEAL. The fan case comprises an inlet and an inlet length between the inlet and the fan. The acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the BEAL determines an acoustic spacing ratio of the gas turbine engine.

A heat dissipation structure includes a housing configured to accommodate a heating element of an aerial vehicle. The housing includes a first air vent configured to guide an airflow into the housing and a second air vent configured to guide the airflow out of the housing. The airflow includes a propeller-generated airflow generated by a propeller of the aerial vehicle during rotation. The housing is located at an end of an arm of the aerial vehicle, away from the propeller in a longitudinal direction of the arm. A projection of the housing on a plane on which an area of rotation of the propeller lies at least partially overlaps the area of rotation of the propeller.

An electric tilt rotor aircraft and a control system thereof. The aircraft comprises a fuselage (1), wings (2), horizontal stabilizers (3) and a landing gear system (4), wherein the fuselage (1) comprises a fuselage body and a rudder (10) for controlling and adjusting a flight direction of the aircraft; the wings (2) each comprise an integral wing box section (21), an aileron (22), a wing left / right nacelle (23) and a supporting rod left / right nacelle (24); left and right supporting rods are integrated on front sides of the integral wing box sections (21); the wing left / right nacelle (23) is connected to the integral wing box section (21) by means of a tilting shaft (25), and the wing left / right nacelle (23) rotates around the tilting shaft (25); the horizontal stabilizers (3) each comprise a horizontal stabilizer box section (31) and a horizontal stabilizer left / right nacelle (32); the horizontal stabilizer left / right nacelle (32) is connected to the horizontal stabilizer box section (31) by means of the tilting shaft (25), and the horizontal stabilizer left / right nacelle (32) rotates around the tilting shaft (25); the landing gear system (4) uses a tricycle landing gear system, and comprises a nose landing gear (41) and two main landing gears (42); brake systems are integrated in the main landing gears (42); and the nose landing gear (41) plays a role in balancing and controlling direction. By using the electric tilt rotor aircraft, urban air traffic can be realized.

A VTOL aircraft includes a lift propeller configured to provide lift during takeoff, landing and hover operations. The lift propeller is configured to be stowed close to a boom surface when not in use during, e.g., cruise flight, to minimize drag around propeller surfaces. The lift propeller may be displaced vertically with respect to the boom when switching to a lift configuration to provide greater separation from the boom to improve lift efficiency and reduce unwanted vibrations. The lift propeller may have a fail-safe configuration that allows the propeller to move passively between lift and stowed configurations, and to fully rotate even when fully retracted.

An aircraft thermal management system comprising a high temperature loop and a low temperature loop. The high temperature loop is configured to manage a temperature of a fuel cell stack in a nacelle using a nacelle heat exchanger. The low temperature loop is configured to cool a number of heat loads in a nacelle using a heat capacity of liquid hydrogen.

A method of operating a gas turbine engine comprising an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system arranged to supply oil to the gearbox; and a heat exchange system comprising an air-oil heat exchanger through which the oil flows, a fuel-oil heat exchanger through which the oil and the fuel flow; and a valve arranged to allow a proportion of the oil sent via at least one of the heat exchangers to be varied, is disclosed, the method comprising controlling the valve such that, under cruise conditions, an oil flow ratio of:rate⁢ of⁢ oil⁢ flow⁢ into⁢ air-oil⁢ heat⁢ exchangerrate⁢ of⁢ oil⁢ flow⁢ into⁢ fuel-oil⁢ heat⁢ exchangeris in the range from 0 to 0.59.

An aircraft includes a battery pack mounted inside the aircraft, a vent coupled between the battery pack and a surface of the aircraft to at least partly define a vent path between the battery pack and the surface of the aircraft, and a burst membrane located in the vent path. The vent may be coupled to a rear upper portion of a wing or to an outboard side of a nacelle. The aircraft may also include a flexible coupling between the vent and the surface of the aircraft. The aircraft may also include a fairing over a vent outlet to provide a smooth surface for the vent outlet. The battery pack may include battery modules and an enclosure, the battery modules and the enclosure defining paths along which discharge from a thermal event can flow towards the vent.

A turbine shroud assembly adapted for use with a gas turbine engine includes a shroud segment. The shroud segment includes a heat shield, an attachment flange, and a multi-layer coating. The heat shield extends circumferentially partway around the axis to define a portion of gas path for the gas turbine engine. The attachment feature extends radially outward from the heat shield. The multi-layer coating is applied to different surfaces of the heat shield and the attachment feature of the shroud segment.

Disclosed are devices, systems and methods for mission-adaptable aerial vehicle. In some aspects, a mission-adaptable aerial vehicle includes a configuration having swappable, manipulatable, and interchangeable sections and components connectable by a connection and fastening system able to be modified by an end-user in the field. In some embodiments, a mission-adaptable aerial vehicle can be configured to include a main center body extending along a longitudinal direction, a wing with a lateral cross-sectional airfoil shape, and / or stabilizer and control surface structures with corresponding cross-sectional airfoil shapes.

A method of operating a gas turbine engine including: an engine core with a turbine, compressor, a fuel combustor, and a core shaft connecting the turbine to the compressor; a fan; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system arranged to supply oil to the gearbox; and a heat exchange system including: an air-oil heat exchanger through which the oil flows; a fuel-oil heat exchanger through which the oil and the fuel flow; a temperature sensor downstream of the fuel-oil heat exchanger; and a valve arranged to allow oil or air flow rate through at least one heat exchanger to be varied, the method including: determining if the fuel temperature has increased above a set threshold at cruise conditions; and in response, controlling the valve so as to change the flow rate through the heat exchanger.

The invention discloses a terahertz-based curved surface thermal barrier coating nondestructive thickness measurement system and method, and the method comprises the steps: firstly carrying out the coordinate calibration and reverse reconstruction of a sample piece and a mechanical arm through an optical scanning and trinocular vision measurement system, and building a virtual space; then matching the virtual space with the actual space, extracting a normal vector of the curved surface of the sample piece, performing path planning on the mechanical arms to generate a control script, and controlling the double mechanical arms to carry a terahertz measuring head to scan the sample piece in a manner of being vertical to the surface of the sample piece to obtain a terahertz signal; and finally, constructing a terahertz signal analytical model, optimizing terahertz simulation signal parameters through a self-adaptive teaching and learning optimization algorithm, and fitting a detection signal until an optimal solution is obtained, namely the thickness of the thermal barrier coating. According to the invention, through cooperation of the two mechanical arms and path planning, rapid detection of a complex curved surface structure is adapted; by establishing an analytical model and combining with an optimization algorithm, the detection precision is improved, and high-precision nondestructive detection suitable for a complex curved surface structure is realized.

A propulsion system for an aircraft as disclosed herein may include a nacelle, a shaft positioned centrally within a cylindrical passageway of the nacelle, a fan coupled to one end of the shaft, a turbine coupled to an opposite end of the shaft, an electric motor coupled to the shaft, a compressor positioned within the cylindrical passageway, and a solid oxide fuel cell positioned with a hollow ring-shaped interior of the nacelle. The hollow ring-shaped interior may surround and be isolated from the cylindrical passageway. The turbine may be configured to provide primary torque to the shaft while the electric motor may be configured to provide additional torque to the shaft. The electric motor may be powered an electric output of the solid oxide fuel cell while the turbine may be powered at least in part by output gases from the solid oxide fuel cell.

A turbine engine for an aircraft. The turbine engine includes a combustor fluidly coupled to a fuel delivery assembly to receive fuel from the fuel delivery assembly. The fuel is injected into the combustor and combusted in the combustor to generate combustion gases. A condenser is located downstream of the combustor to receive the combustion gases and to condense water. A fuel heat exchanger is thermally coupled to the condenser by at least one heat transfer loop to receive heat from the condenser.

Disclosed are devices, systems and methods for mission-adaptable aerial vehicle. In some aspects, a mission-adaptable aerial vehicle includes a configuration having swappable, manipulatable, and interchangeable sections and components connectable by a connection and fastening system able to be modified by an end-user in the field. In some embodiments, a mission-adaptable aerial vehicle can be configured to include a main center body extending along a longitudinal direction, a wing with a lateral cross-sectional airfoil shape, and / or stabilizer and control surface structures with corresponding cross-sectional airfoil shapes.

A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.

An airfoil is provided that includes a first end, a second end, a leading edge, a trailing edge, a pressure side and a suction side. The leading edge and the trailing edge are joined by the pressure side and the suction side to provide an exterior airfoil surface extending in a spanwise direction from the first end of the airfoil to the second end of the airfoil. The exterior airfoil surface are formed in conformance with a plurality of cross-section profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading edge and the trailing edge at the span location.

A turbomachine includes an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline. The fan includes fan blades that extend radially outwardly toward the annular casing. The fan has an average chord fan width according to a first performance factor. The fan has a quantity of fan blades according to a second performance factor.

A turbomachine includes an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline. The fan includes fan blades that extend radially outwardly toward the annular casing. The fan has an average chord fan width according to a first performance factor. The fan has a quantity of fan blades according to a second performance factor.

A turbomachine includes an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline. The fan includes fan blades that extend radially outwardly toward the annular casing. The fan has an average chord fan width according to a first performance factor. The fan has a quantity of fan blades according to a second performance factor.

An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.

An aircraft air management system comprises an air heat exchanger; an intercooler; and a conduit system connected to the air heat exchanger, the intercooler, and a fuel cell stack. Heated air flows through the conduit system to the air heat exchanger. The air heat exchanger is configured to cool the heated air to form cooler air, wherein the cooler air flows from the air heat exchanger through the conduit system to the intercooler; and the intercooler is configured to cool the cooler air to form cooled air, wherein the cooled air flows from the intercooler through the conduit system to the fuel cell stack.

An electrical propulsion system for a vertical take-off and landing (VTOL) aircraft comprises an electrical motor assembly and an inverter assembly. The inverter assembly comprises a housing, a capacitor assembly, at least one printed circuit board assembly (PCBA), and a plurality of positioning pins. The capacitor assembly comprises a center hole, at least one capacitor, a capacitor housing having at least one busbar, and a plurality of through holes in the capacitor housing. The capacitor assembly and the at least one PCBA are positioned inside the housing. The plurality of positioning pins pass through the through the plurality of through holes of the capacitor housing and the at least one PCBA and are connected to the housing.

A turbomachine includes an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline. The fan includes fan blades that extend radially outwardly toward the annular casing. The fan has an average chord fan width according to a first performance factor. The fan has a quantity of fan blades according to a second performance factor.

A method of operating a gas turbine engine includes an engine core with a turbine, compressor, fuel combustor, and core shaft; a fan upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system supplying oil to the gearbox; and a heat exchange system including: air-oil and fuel-oil heat exchangers, and wherein the oil loop system branches such that part of the oil flows along each branch and the air-oil and fuel-oil heat exchangers are parallel on different branches; and a modulation valve allows the oil sent via each branch to be varied, the method including controlling the heat exchange system wherein, under cruise conditions, a heat transfer ratio of:rate⁢ of⁢ heat⁢ transfer⁢ from⁢ oil⁢ to⁢ airrate⁢ of⁢ heat⁢ transfer⁢ from⁢ oil⁢ to⁢ fuelis in the range from 0 to 0.67.

A propulsion system for an aircraft as disclosed herein may include a nacelle, a shaft positioned centrally within a cylindrical passageway of the nacelle, a fan coupled to one end of the shaft, a turbine coupled to an opposite end of the shaft, an electric motor coupled to the shaft, a compressor positioned within the cylindrical passageway, and a solid oxide fuel cell positioned with a hollow ring-shaped interior of the nacelle. The hollow ring-shaped interior may surround and be isolated from the cylindrical passageway. The turbine may be configured to provide primary torque to the shaft while the electric motor may be configured to provide additional torque to the shaft. The electric motor may be powered an electric output of the solid oxide fuel cell while the turbine may be powered at least in part by output gases from the solid oxide fuel cell.

An isothermal near-net forming method of a TC4 titanium alloy aeroengine blisk comprises the following steps that (1) the surface of a TC4 titanium alloy bar is coated with a glass lubricant, and the TC4 titanium alloy bar is heated to T beta-(30-80) DEG C; 2) coating a glass lubricant on the pre-forging die and heating the pre-forging die to T beta-(30-80) DEG C; (3) the bar is loaded into a pre-forging die to be forged into an intermediate blank, at least one heating number of forging is carried out, the strain rate is smaller than or equal to 0.01 s <-1 >, and the deformation amount is 60-70%; after forging is completed, the intermediate blank is water-cooled to the room temperature; (4) the intermediate blank is heated to T beta-(30-80) DEG C; 5) coating a glass lubricant on a finish forging forming die and heating to T beta-(30-80) DEG C; the method comprises the following steps of (1) carrying out primary forging on an intermediate blank, (2) carrying out secondary forging on the intermediate blank, (3) carrying out secondary forging on the intermediate blank, (4) carrying out secondary forging on the intermediate blank, (5) carrying out secondary forging on the intermediate blank, (6) loading the intermediate blank into a final forging forming die, pressing the intermediate blank into a blisk forge piece, carrying out at least one-heating-number forging, carrying out water cooling on the blisk forge piece to room temperature, (7) carrying out solution treatment on the blisk forge piece and carrying out water cooling to room temperature, and (8) carrying out aging treatment on the blisk forge piece and carrying out air cooling to room temperature.

The invention provides a thermal protection coating as well as a preparation method and application thereof. The thermal protection coating comprises a NiCoCrAlY first bonding layer, a NiCoCrAlY second bonding layer, a CeYSZ first ceramic layer and a GdYbYSZ second ceramic layer which are sequentially arranged on the surface of a base material, and the second NiCoCrAlY bonding layer is made of CeO2 (CeO2) and Al2O3 (Al2O3). According to the invention, a thermal protection coating structure with double bonding layers and double ceramic layers is constructed on the surface of the base material, and CeO2 and Al2O3 nano-particle reinforced phases are introduced, so that the use temperature of the thermal protection coating is increased, and the thermal strain tolerance, the crack resistance and the oxidation resistance are greatly improved.

The invention relates to an aircraft turbomachine subassembly comprising an annular duct (36) extending concentrically about a longitudinal axis and in which an air stream flows from upstream to downstream, the annular duct comprising a duct portion (28) defining a swan neck having an upstream inlet section and a downstream outlet section. A plurality of radial arms (B1, B2) extend circumferentially in the duct portion (28). One or several intermediate walls (P1, P2) extend between two circumferentially adjacent arms (B1, B2) which form a pair of circumferentially adjacent arms among the plurality of circumferentially adjacent arms so as to locally separate, between the two arms of the pair of arms, the annular duct portion (28) into a radially inner portion (Ci) and a radially outer portion (Ce) further away from the longitudinal axis than the radially inner portion.