4404results about "Blade accessories" patented technology

A wind energy plant with a nacelle, a rotor, which features at least one rotor blade adjustable in its blade pitch angle, a central control device for controlling the wind energy plant and a control unit disposed in the rotor for controlling the blade pitch angle of the at least one rotor blade, wherein the central control device and the control unit in the rotor can exchange data with each other via a data link, which comprises at least one first sending and receiving device at the nacelle side and at least one second sending and receiving device at the rotor side, wherein a wireless network connection with a safety-oriented communication protocol is provided between the at least one first sending and receiving device and the at least one second sending and receiving device, and that one monitoring- and communication device at a time is associated to the first and / or the second sending and receiving device, which can monitor the function of the sending and receiving device and initiate a predetermined action in the case of an error.

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 systematic framework for optimizing multi-stage axial compressor disk designs in gas turbine engines. By combining finite element analysis (FEA), Design of Experiments (DOE), and optimization algorithms of multi-objective genetic algorithm (MOGA), the method balances stress, deformation, and mass to enhance structural performance. The six-step process includes blade modeling, parameterizing disk geometry, structural analysis using FEA, developing functional relationships, applying optimization algorithms, and generating manufacturable 3D disk models. This approach reduces weight, improves fuel efficiency, and adapts to various compressor designs and materials, enhancing the overall performance of gas turbine engines.

A power turbine blade for a gas turbine engine includes at least one of a third-stage blade, a fourth-stage blade, or a fifth-stage blade. The power turbine blade includes a platform and an airfoil that extends in a spanwise direction from the platform to a tip shroud. The airfoil has an external airfoil surface formed in conformance with multiple cross-section airfoil profiles described by a set of Cartesian coordinates.

A gas turbine engine includes: a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath; a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio during operation of the gas turbine engine in a cruise operating mode; and a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan; wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8.

An assembly for an aircraft propulsion system includes an open propulsor rotor and an open guide vane structure. The open propulsor rotor is rotatable about a rotational axis. The open guide vane structure is axially next to and downstream of the open propulsor rotor. The open guide vane structure includes a plurality of open guide vanes arranged circumferentially about the rotational axis. The open guide vanes include a first open guide vane. The first open guide vane includes a camber line, a leading edge, a trailing edge, a leading edge section and a trailing edge section. The first guide vane extends longitudinally along the camber line from the leading edge to the trailing edge. The leading edge section is configured to pivot about a pivot axis. The trailing edge section is configured to move relative to the leading edge section to change a camber of the first open guide vane.

The invention discloses a nickel-based single-crystal high-temperature alloy with high structure stability and a preparation method thereof, and relates to the technical field of high-temperature alloys. The nickel-based single crystal high-temperature alloy provided by the invention is prepared from the following components in percentage by mass: 5 to 7 percent of Al, 7.5 to 10 percent of Co, 5 to 7 percent of Cr, 0.1 to 0.2 percent of Hf, 0 to 2 percent of Mo, 5.5 to 7.5 percent of Ta, 0 to 1 percent of Ti, 8 to 10 percent of W, 0.02 to 0.05 percent of C, 0.003 to 0.005 percent of B and the balance of Ni. According to the invention, by controlling the ratio of elements and avoiding the addition of Re, the precipitation temperature of a harmful TCP phase is obviously reduced to about 840 DEG C on the basis of maintaining a sufficient solid solution strengthening effect, and the long-term stability of the structure is greatly improved. The high-temperature mechanical property of the alloy provided by the invention reaches the level equivalent to that of a traditional Re-containing second-generation single-crystal alloy and is remarkably superior to that of foreign Re-free single-crystal alloys MD2 and Rene N500, and the nickel-based single-crystal high-temperature alloy with high structure stability successfully realizes the unification of low cost and high performance, and is suitable for industrial production. And a reliable material solution is provided for manufacturing the heavy-duty gas turbine blade with higher market competitiveness.

The invention relates to the technical field of engine blades, and discloses an electric heating assembly for ice prevention and removal of an engine blade and a preparation method and application of the electric heating assembly. The electric heating assembly and the engine blade are conformal, the long edge of the short axis direction of the engine blade is parallel to the first direction, and the second direction is perpendicular to the first direction; the electric heating assembly comprises a plurality of electric heating pieces which are sequentially arranged in the second direction and connected through a parallel circuit, each electric heating piece is provided with two second edges oppositely arranged in the second direction, and the second edges are parallel to the first direction; the electric heating piece comprises a conductive fabric, and the conductive fabric comprises a fiber fabric and a graphene electric heating layer conformally wrapping the fiber fabric. The electric heating assembly has good heat distribution uniformity, and the anti-icing and de-icing effects of the engine blade can be improved.

The invention provides a thermal barrier coating with an interface three-dimensional bionic structure and a preparation method and application of the thermal barrier coating, and belongs to the technical field of thermal barrier coatings. By means of 3D printing, preparation of a complex, special and arrangement-distribution-controllable three-dimensional texture component with a complex geometrical shape and an internal structure can be achieved, and therefore the coating interface residual stress, the contact area and the pinning effect are improved; 3D printing can avoid internal oxidation of the bonding layer, and meanwhile it is ensured that the structure of the multi-level interface is uniform and stable; and then the ceramic layer is prepared by combining the spraying technology, preparation of the interface three-dimensional bionic structure thermal barrier coating is achieved, interface mechanical engagement of the bonding layer and the ceramic layer can be increased through the structure, an interlocking mechanism is achieved, the coating interface structure and residual stress are improved, and therefore the service life of the coating is greatly prolonged, and the bonding strength of the coating is greatly improved.

The invention provides a turbine blade with a cooling structure for a gas turbine, and relates to the technical field of gas turbine blade cooling, the turbine blade comprises an inner layer wall and an outer layer wall, the inner layer wall is a hollow pipe body extending out from the normal direction of a turbine rotating shaft, and the hollow part is used as a cold air flow channel and receives cooling air flow. The outer layer wall is arranged on the outer side of the inner layer wall in a sleeving mode at intervals, an annular runner is defined by the outer layer wall and the inner layer wall, a plurality of impact holes are formed in the inner layer wall, a plurality of air film holes are formed in the outer layer wall, and cooling air flow in the cold air flow channel enters the annular runner through the impact holes to impact and cool the outer layer wall and is attached to the outer layer wall after being jetted through the air film holes. And a protective layer is formed between the outer layer wall and high-temperature fuel gas to realize gas film cooling. The opposite surfaces of the inner layer wall and the outer layer wall are flow guide curved surfaces, the projection of the flow guide curved surfaces in the normal direction of the rotating shaft is a first curve, the projection of the flow guide curved surfaces in the tangential direction of the annular flow channel is a second curve, and the flow guide curved surfaces are modeled to restrain secondary contra-rotating vortexes generated when cooling airflow flows in the annular flow channel.

A gas turbine engine includes, among other things, a propulsor section including a rotor and blades, a geared architecture including a gear train, a compressor section and a turbine section. A static structure includes a first case and a second case. The first case includes a first engine mount location. The second case includes a second engine mount location. The first engine mount location is axially near the geared architecture.

The turbocharger impeller comprises a first impeller disc and a limiting shell, a plurality of concave sliding grooves are formed in the upper surface of the first impeller disc, the interiors of the concave sliding grooves are connected with convex sliding blocks in a sliding mode, the upper surfaces of the convex sliding blocks are fixedly connected with impeller blades, and the lower surfaces of the convex sliding blocks are fixedly connected with the limiting shell. Inserting grooves are formed in the two sides of the upper surface of the limiting shell, bolts matched with the inserting grooves are fixedly connected to the two sides of the upper surface of the first impeller disc, and the surfaces of the bolts are sleeved with fixing nuts in a threaded mode. According to the device, the concave sliding grooves and the convex sliding blocks are arranged, so that the device can simply and conveniently disassemble and assemble the impeller blades, when a certain impeller is damaged in the using process of the turbocharger, the damaged impeller can be directly disassembled, and the limiting shell is arranged, so that the impeller can be conveniently disassembled and assembled. And the device can limit the impeller blades into the first impeller disc through the limiting shell, and dislocation of the impeller blades is avoided when the turbocharger works.

A turbine wheel may include an inner hub having a central axis, an outer ring concentric with the inner hub, and an intermediate ring concentrically disposed between the inner hub and the outer ring, and a plurality of turbine blades extending between the intermediate ring and the outer ring. The turbine blades may be oriented to be driven by a fluid stream flowing through the turbine wheel to rotate the turbine wheel in a first direction about the central axis. In addition, the turbine wheel may further include a plurality of compressor blades extending between the inner hub and the intermediate ring. Also, the compressor blades may be oriented to propel a fluid in a downstream direction when the turbine wheel is rotated in the first direction.

A surface polishing method for micro inner flow channel, a micro inner flow channel part and a polishing medium. The diameter of the micro inner flow channel is less than or equal to 3 mm and length-diameter ratio is greater than or equal to 50:1. A polishing medium including a liquid phase and a solid phase, and viscosity of the liquid phase <1000 cP, the solid phase includes abrasive particles; a predetermined pressure being set on the polishing medium, making the polishing medium flow in the micro inner flow channel at a flow velocity more than 5 m / s, and flow rate of the polish polishing medium flowing into the micro inner flow channel at one end reaches a saturation flow rate allowed by the bore of the micro inner flow channel, making the hydraulic pressure inside the inner flow channel be in a pressure holding state.

A cooled brush seal for use with a CMC component of a gas turbine engine having a hot side exposed to a hot gas path and an opposing side having a cavity for feeding a cooling flow to at least one film cooling hole. The cooled brush seal includes: an OD backing plate; an ID backing plate; a plurality of brush seal bristles sandwiched between the OD backing plate and the ID backing plate; a first hole extending through the OD backing plate, the brush seal bristles, and the ID backing plate; and a cover plate attached to an inner surface of the ID backing plate and including a second hole circumferentially offset from the first hole and forming a second passage into the cavity, wherein the ID backing plate and the cover plate form a circumferential cooling passage extending between the first hole and the second hole.

The invention discloses a composite environmental barrier coating as well as a preparation method and application thereof, and relates to the technical field of high-temperature protective coatings. The composite environmental barrier coating comprises a bonding layer, a barrier layer and a ceramic layer, wherein the bonding layer comprises Si and HfO2 in a molar ratio of (1: 10)-(1: 5); the ceramic layer is (Lux, Yby) 2SiO5, and 0.1 lt; xlt; 0, 0.4, 0.6 lt; yt; Yt; 0.9, 0.9; the thickness of the composite environmental barrier coating is not greater than 200 [mu] m. The composite environmental barrier coating can resist CMAS corrosion for up to 100 h in the atmospheric environment of 1300-1500 DEG C, the coating does not peel off after being corroded by water and oxygen for 300 h, and the composite environmental barrier coating has excellent heat insulation performance, bonding strength and thermal cycle performance and is expected to be applied to preparation of aero-engine hot end components.

A ring fastener comprising a discontinuous band having first and second axial faces on opposite sides, a first end with a first end face and a groove extending from the first end face and opening to the first axial face; a second end with a second end face and a tongue projecting from the second end face; wherein the groove is u-shaped and the second axial face defines the base of the groove, and the tongue is dimensioned to fit in the groove; such that, when the tongue is received in the groove, the first end is located relative to the second end and a continuous ring is formed. A method of assembling the ring fastener into a circumferential groove in a component comprises the steps of feeding the ring fastener into the groove so that the ends are adjacent to each other; distorting the second end axially out of the plane of the groove; bringing the ends towards each other until the tongue overlies the groove; and inserting the tongue into the groove so that the ends are within groove.

A gas turbine engine component, comprises a densified three-dimensional woven ceramic matrix composite comprising the shape of the gas turbine engine component, the densified three-dimensional woven ceramic matrix composite further comprises a three-dimensional woven structure comprising a densified ceramic matrix disposed on, within and throughout a three-dimensional weave comprising at least one inorganic fiber comprising at least one multilayer coating material structure disposed thereupon, wherein the at least one multilayer coating material structure comprises at least one first interface coating material layer comprising a boron nitride; at least one coating material layer comprising a silicon nitride; at least one second interface coating material layer comprising the boron nitride; and at least one third interface coating material layer comprising a Si3N4-BN multilayer coating having one or more alternating layers comprising at least one layer of silicon nitride at a first thickness; and at least one layer of boron nitride at a second thickness.

The disclosed fan blade is composed at least in part of a structure (50) made of composite material comprising a fiber reinforcement and a matrix. The fiber reinforcement includes a plurality of plies that are each formed of warp yarns and weft yarns. The warp yarns and weft yarns include incomplete yarns that each have a terminal end on the pressure side (40) or the suction side. The pressure side (40) and the suction side each have at least one ply drop line (80, 81) connecting the terminal ends of the incomplete yarns of the same ply. At least one ply drop line (81) has an end portion (96) that extends from the blade tip up to 90% of the blade height. Said end portion (96) extends generally along a direction of extent (E) forming an angle between 5° and 85° with the local tangent of the warp yarns.

A turbine engine rotor has a central shaft and a disk stack having disks encircling the shaft. A seal has in central axial cross-section: a rearwardly-open channel receiving a portion of one of the disks; a sleeve extending rearward from the channel and between the disk and the shaft; and a portion extending radially inward from the sleeve and having a forward surface restrained against forward movement relative to the shaft.

A gas turbine engine is provided that includes a combustor, a rotational spool, a spool balancing system, and at least one access port. The rotational spool includes turbine and compressor sections, and a shaft that connects the turbine and compressor sections. The spool balancing system includes a circumferentially extending track component and a plurality of balance weights that are engaged with the track component. The balance weights are disposable in secured and unsecured configurations. The track component and the balance weights are disposed to rotate with the rotational spool. The access port is disposed within a non-rotational member of the gas turbine engine. The access port is configured to receive a tool for engagement with the balance weights.

A device and related method that provides, but is not limited thereto, a two-phase heat transfer device with unique combination of enhanced evaporation and increased cooling capacity. An advantage associated with the device and method includes, but is not limited thereto, increased cooling capacity per unit area, controlled and optimized evaporation, prevention of boiling, and prevention of drying of the evaporator. An aspect associated with an approach may include, but is not limited thereto, using a non-wetting coating or structure to keep working fluid away from the spaces between elongated members of an evaporator and using a wetting coating or structure to form thin films of working fluid around the distal region of the elongated members. For example it can be used to cool a computer chip, a skin of a hypersonic flying object, parabolic solar collector, turbine or engine blade, or any other heat source that requires high heat flux.

The invention discloses a blade hub sweepback type self-rectification axial flow air turbine runner, and belongs to the field of wave energy power generation devices. The impeller comprises a hub, the hub is rotationally arranged in a runner chamber, the hub is arranged between a front guide vane and a rear guide vane, a plurality of blades are arranged on the hub, the blades are evenly arranged in the circumferential direction with the axis of the hub as the center, and each blade comprises an air inlet edge adjacent to the front guide vane and an air outlet edge adjacent to the rear guide vane. Axial airflow flows out along the front guide vane, enters the positions among the blades of the hub through the air inlet side, acts to generate torque, flows out from the air outlet side, is rectified by the rear guide vane and flows out in the axial direction of the hub. The invention aims at solving the problems that the local loss is increased and the efficiency is influenced due to the large difference of liquid flow angles from a hub to a rim inlet of the current axial flow air turbine, reducing the local loss of an impeller and improving the turbine efficiency.

The invention discloses a runner rotary vane mechanism for a fluid machine and an engine. The runner rotary vane mechanism comprises a differential control device, a runner, a left impeller set and a right impeller set. A plurality of inner transmission shafts and an outer transmission shaft of the differential control device are concentrically arranged in a sleeved mode. The rotating wheel is installed in a cylinder body of the stator, and a left separation blade and a right separation blade are arranged on the two sides of a wheel disc of the rotating wheel respectively. The cylinder body is divided into a left cylinder and a right cylinder by the rotating wheel, and a plurality of left blades of the left impeller group are arranged in the left cylinder; and a plurality of right blades of the right impeller group are arranged in the right air cylinder. And the rotating wheel rotates under the control of the differential control device so as to carry out two-stage fluid conveying work.

An assembly is provided for a turbine engine. This assembly includes an engine structure, and the engine structure includes a plenum, a reverse flow combustor, a nozzle structure and a cooling circuit. The reverse flow combustor is disposed in the plenum and includes a combustion chamber. The nozzle structure is arranged at an outlet from the combustion chamber. The nozzle structure includes a first platform, a second platform and a plurality of vanes arranged circumferentially about an axis. The vanes extend across the flowpath from the first platform to the second platform. The cooling circuit extends in the first platform and between a circuit inlet into the cooling circuit and a circuit outlet from the cooling circuit. The circuit inlet is fluidly coupled with and disposed along the plenum. The circuit outlet is fluidly coupled with and disposed along the flowpath.

The invention discloses a protrusion and rib combined structure for target surface cooling and a double-wall impingement cooling device, and relates to the technical field of blade heat exchange, the protrusion and rib combined structure comprises a plurality of combined structures, each combined structure comprises a protrusion and a flow guide rib, and each protrusion and each flow guide rib are used for being fixedly connected with the target surface, close to an impingement plate, of a target plate; the impact plate is provided with a plurality of impact holes, the target plate is provided with a plurality of air film hole groups, each air film hole group comprises at least three air film holes, and the air film holes of each air film hole group are arranged at intervals and define a cooling space; each bulge is fixedly connected in the cooling space of one target plate, each bulge is arranged opposite to one impact hole in the impact plate, and each bulge can enable airflow entering from the corresponding impact hole to flow in the direction close to the target surface and far away from the center line of the corresponding bulge; a flow guide rib is arranged between every two adjacent air film holes of each air film hole set, and each flow guide rib extends from one corresponding air film hole to the other corresponding air film hole. The overall heat exchange effect of the blade is remarkably improved.