2483results 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 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.

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.

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.

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.

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 turbine engine rotor has a central shaft and a disk stack having a plurality of disks encircling the shaft. A seal has a first member having, 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. The seal has a second member having, in central axial cross-section: a forward-facing surface contacting an aft-facing surface of the shaft; and a rear-facing surface contacting the forward surface of said portion extending radially inward from the sleeve.

A method of overhaul is provided. During this overhaul method, a substrate is scanned using structured light to provide substrate scan data. The substrate is from a component previously installed within an engine. The substrate scan data is compared to substrate reference data to provide additive manufacturing data. Material is deposited with the substrate using an additive manufacturing device based on the substrate scan data to provide a first object. The first object is scanned using the structured light to provide first object scan data. The first object scan data is compared to first object reference data to provide machining data. The first object is machined using the machining data.

The utility model relates to a novel slotting structure for improving the air film cooling efficiency of a turbine blade. The structure comprises a wall surface, an air film hole and a transverse slot. The air film holes are formed in the upstream area of the wall face, the air film holes are in an air film hole shape with the inlet expanded and the outlet flow direction gradually shrunk and expanded in the spanwise direction, the inlet expansion angle is 5-15 degrees, the outlet flow direction gradually shrunk angle is 5-10 degrees, the outlet is in a rounded rectangle shape, and the length of the outlet ranges from 1.5 D to 2D. The transverse groove is formed in the outlet position of the air film hole, the structure is arranged to be in a double-step shape on the downstream groove wall, the step close to the groove bottom is called as the first step, the appearance structure is rectangular, the height of the first step is 0.5 D to 1D, the width of the first step in the main flow direction is 0.5 D to 1D, the bottom of the first step is perpendicular to the groove bottom, the step close to the wall surface is called as the second step, and the outer surface is an arc surface; the tangent line of the highest point coincides with the wall face, the tangent line of the lowest point is perpendicular to the first step surface, and the radius of the arc is 0.5 D-1D. The flow channel design and the step-shaped structure in the cooling structure can effectively improve the air film cooling efficiency.

A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:(TOUTADTExit)2*EGTAHPCExit*1⁢0-1⁢1;wherein CSP is greater than 0.0001194×EGT2−0.103×EGT+22.14 and less than 0.0003294×EGT2−0.306×EGT+77.91; and wherein EGT is greater than 525 degrees Celsius and less than 1250 degrees Celsius.

A method of operating an automated fiber placement system includes supplying a plurality of tows to a tow deposition assembly of an automated fiber placement head, cutting a first tow into a first upstream and downstream tow portions, continuing to supply the first tow through a compaction roller to apply the first tow to a substrate such that a trailing edge of the first upstream tow portion substantially aligns with a leading edge of the first downstream tow portion at a first interface, cutting a second tow into a second upstream and downstream tow portions, and continuing to supply the second tow through the compaction roller to apply the second tow the substrate such that a trailing edge of the second upstream tow portion substantially aligns with a leading edge of the second downstream tow portion at a second interface that is offset from the first interface.

An airfoil assembly and engine component for a turbine engine having an engine core extending along an engine centerline, the airfoil assembly having an airfoil extending between a root and a tip in a radial direction, away from the engine centerline, to define a span length. The airfoil comprising a composite core and a set of skins defining at least a portion of an exterior surface of the airfoil. A shank extending in the radial direction from the root toward the engine centerline and having a flared cross-section.

The disclosure describes a ceramic matrix composite (CMC) component having a cooling cavity. The CMC component has a base having a bottom inner surface and a top outer surface, in which the base is made from a plurality of ceramic fiber plies within a matrix. A cooling cavity is provided within the base in which the cooling cavity has an opening in the top outer surface of the base. A cover plate made from one or more ceramic fiber plies within a matrix is provided to close the opening of the cooling cavity. The cover plate is bonded to the base by simultaneous densification of the base and cooling cavity.

A pressure side rib wall is arranged on the side, close to the tip end face, of the pressure face of the high-pressure turbine blade, a suction side tail-cut winglet is arranged on the side, close to the tip end face, of the suction face, and the pressure side rib wall and the suction side tail-cut winglet form a groove located in a blade tip. The inclined rib wall on the pressure side is used for preventing fluid from entering the gap, meanwhile, the suction side tail cut winglet of an improved structure is adopted, formation and development of leakage vortexes can be controlled, a high-heat load area at the tail edge of the suction surface can be improved, cooling holes are formed in the bottom of the groove to effectively cool the blade top, and therefore the purposes of reducing leakage loss and improving the service life of the blade are achieved. And the heat exchange performance of the blade tip is improved.

A fan section of an aeronautical propulsion system includes a fan rotor including seventeen blades to twenty blades and having a solidity strictly less than 1.0. The solidity is equal to a ratio between a chord at the blade tip and an inter-blade pitch at the blade tip. The fan section has a hub-tip ratio greater than or equal to 0.22 and less than or equal to 0.32, a pressure ratio greater than or equal to 1.05 and less than or equal to 1.5, and a peripheral speed at the blade tip greater than or equal to 260 m / s and less than or equal to 400 m / s. The pressure ratio and the peripheral speed are measured at cruising speed.

A blade fiber preform for a turbomachine has a three-dimensional or multi-layer weave including in one piece a root preform part, a shank preform part, first and second platform preform parts, upstream and downstream wall preform parts, upstream and downstream spoiler preform parts, and an airfoil part. The preform further includes unidirectional fabric portions folded down onto the shank preform part and third and fourth unidirectional fabric portions folded down onto the airfoil part.

The invention discloses a double-ceramic-layer thermal barrier coating of a differentiated columnar crystal structure and a preparation method, and belongs to the technical field of thermal barrier coating corrosion and protection. The invention innovatively discovers and utilizes the double influence of the columnar crystal size on the performance of the coating; the YSZ layer adopts a fine grain structure of 1-1.5 microns, so that the bonding strength and thermal shock resistance of the coating can be obviously improved; and the modified gadolinium zirconate layer adopts a 2-3 [mu] m coarse grain structure, so that the heat conductivity can be reduced, the permeation path of CMAS along a grain boundary can be effectively inhibited, and meanwhile, the modified gadolinium zirconate layer reacts with CMAS to form a compact barrier layer to prevent further permeation of CMAS. Through accurate regulation and control of electron beam physical vapor deposition (EB-PVD) process parameters, differential control of the sizes of the two layers of ceramic columnar crystals is realized, and under the synergistic effect of the two columnar crystal structures with different sizes, the thermal cycle life and CMAS corrosion resistance of the whole coating system in an ultra-high-temperature environment are remarkably prolonged and improved.

A method of servicing aviation equipment including at least partially-autonomously inspecting components of the equipment using an environmental capture device, wherein at least partially-autonomous inspection includes: capturing information associated with the equipment using the environmental capture device; identifying the components of the equipment using one or more computing devices; determining, by the one or more computing devices, whether each identified component is properly positioned relative to the equipment using reference information; and for components not properly positioned relative to the equipment, determining if there is damage to the component or equipment.

The invention provides a nickel-based superalloy with high cobalt and gamma'content and a preparation method and application thereof, and belongs to the technical field of superalloy additive manufacturing. The invention provides spherical nickel-based high-temperature alloy powder which comprises the following components in percentage by mass: 20 to 30 percent of Co, 8 to 10 percent of Cr, 4 to 7 percent of Al + Ti, 14 to 16 percent of W + Mo + Ta, 0 to 3.5 percent of Nb, 0.001 to 0.015 percent of B, 0.05 to 0.1 percent of C and the balance of Ni. And the spherical nickel-based superalloy powder serves as a printing raw material, laser additive manufacturing is conducted, then heat treatment is conducted, and the nickel-based superalloy with the high cobalt content and the high gamma'content is obtained. The nickel-based superalloy with high cobalt and gamma'content prepared by the method has excellent high-temperature mechanical properties and excellent printability, and is suitable for direct forming of turbine blades or additive manufacturing repair of damaged turbine blades.

The application belongs to the field of turbine blade thermal barrier coating of an aero-engine. Specifically, it is a structure design, a preparation method and a prepared coating of a double-layer radiation-resistant thermal barrier coating with a high-reflection layer and a high-absorption layer. The coating comprises a high-absorption layer and a high-reflection layer. The high-absorption layer can absorb part of the energy of photon radiation and re-emit it into the environment, while relying on a cooling gas film to quickly remove heat. Part of the energy of the photon radiation is reflected by the high-reflection layer again to the high-absorption layer after passing through the high-absorption layer, and the process of absorption-emission-cooling gas film heat dissipation is repeated. Within a set temperature range, the absorption rate of the high-absorption layer is determined based on the temperature difference requirement between the inner and outer surfaces of the thermal barrier coating, and the high-absorption layer material is determined based on the absorption rate. The application greatly reduces the radiation heat flow penetrating the coating, avoids the direct heating of the high-temperature alloy by infrared radiation, and significantly improves the heat insulation capacity of the thermal barrier coating and the service life of the turbine blade.

A method for manufacturing a tubular fiber preform for a composite material part. A method for manufacturing a tubular preform comprises: - the creation by three-dimensional weaving of two fiber skins, - the shaping of the first fiber skin so as to form a first part (310) of a tubular fiber preform comprising a first portion (314) of the outer shell of a cylinder (C) of predetermined diameter (D), - the shaping of the second fiber skin so as to form a second part (320) of a tubular fiber preform comprising a second portion (324) of the outer shell of the cylinder of predetermined diameter, - the fitting of the second part (320) into the first part (310) of the tubular fiber preform so as to form a tubular fiber preform (300) having an external surface corresponding to the joining of the first and second portions (314,324) of the outer shell of the cylinder (C) of predetermined diameter (D), the remaining portions (3211, 3212) of the second skin (200) being folded inside the tubular fibrous preform. Figure for the abstract: Fig. 9.

The invention relates to a bladed turbomachine assembly extending around an axis of rotation and comprising blades (11a, 11b), each blade (11a, 11b) extending in a radial direction with respect to the axis of rotation up to a heel (14a, 14b), characterized in that it comprises: – an electrically conductive rod (26) which passes through the heels (14a, 14b) of the blades (11a, 11b), this rod (26) being subjected to shear stress if a heel (14a, 14b) shifts along the direction of the axis of rotation relative to another circumferentially adjacent heel (14a, 14b); – detection means connected to the ends of the rod (26) to identify that this rod (26) is broken when it is no longer electrically conductive. Figure for the summary: Figure 8

This stator blade comprises a blade body having a cross-sectional shape that forms a blade profile. The blade body includes: a trailing end space and an adjacent space that are located between a leading edge and a trailing edge; and a plurality of trailing end cooling passages that penetrate from the trailing end space to the trailing edge. The adjacent space is located further on the leading edge side than the trailing end space. The edge of the trailing end space on the leading edge side is open and communicates with the trailing end space. A plurality of divider ribs, which are aligned in the blade height direction and divide the trailing end space in the blade height direction, and a plurality of pins are arranged in the trailing end space. The plurality of pins all both joined to a positive pressure-side inner wall surface and a negative pressure-side inner wall surface that define the trailing end space. Among the plurality of divider ribs, at least one divider rib forms an extending divider rib which extends to the interior of the adjacent space and which is joined to an inner wall surface demarcating the adjacent space.

Disclosed is a turbine airfoil including an airfoil body including an outer wall configured to define an internal space, in which a cooling hole is formed in the outer wall and allows the internal space S and an external space of the airfoil body to communicate with each other. The inner surface of the outer wall, which defines the cooling hole, includes a first inner surface, and a second inner surface configured to define an outer hole region of the cooling hole H that communicates with the external space, in which the outer hole region is provided as a plurality of outer hole regions spaced apart from one another, and in which the outer wall includes a partition section provided between the two adjacent outer hole regions.

The present disclosure provides a kind of blade for inhibiting the blade of leakage vortex break, blade assembly, impeller machinery and blade shaping method.The blade profile is defined according to the cross section profile of the height direction accumulation, the cross section profile has the ratio of the height of the cross section profile and the total height of the blade, the dimensionless blade height H=1 is the blade tip, the dimensionless blade height H=0 is the blade root;The cross section profile also has the back bending angle δ;At the dimensionless blade height H=1, the back bending angle δ=13 °, and has the tolerance range of the numerical ratio ±20%.

The invention relates to a stator assembly (10) for a turbine engine compressor, comprising two successive rows of upstream stator vanes (121) and downstream stator vanes (124) defining a tandem configuration, the vanes (121, 124) extending radially between inner and outer support members, wherein, within the tandem configuration, the chord (26) of the upstream vane (121), extending between the leading edge and the trailing edge of the vane, increases with the height of the upstream vane: at 0% of the height of the upstream vane (121), the chord (26) of the upstream vane is at least 1.9 times smaller than the chord of the upstream vane at 100% of the height of the upstream vane and, at 70% of the height of the upstream vane (121), the chord (26) of the upstream vane is at least 1.2 times smaller than the chord of the upstream vane at 100% of the height of the upstream vane.

A fibrous reinforcing texture (10) for a part made of composite material has three-dimensional weaving between juxtaposed layers of warp yarns (C1-C8) on several wales (CC1-CC7) of warp yarns and layers of weft yarns (T1-T3). Each warp yarn has a specific initial fiber volume ratio TVFi and a count k corresponding to the number of thousands of filaments present in a warp yarn. The warp yarns are twisted yarns. The width (ICCH) of each wale (Cc1-Cc7) of warp yarns is between 1.8 times and 2.2 times a value a, the value a being calculated using the following formula: (I), where d is the diameter of the filaments of the warp yarns or strands, and TVFc is the desired fiber volume ratio of the warp yarns or strands.

A support device (1) for a stator blade (22) of an aircraft turbomachine with longitudinal axis (X) comprising an external wall (21) from which the stator blade (22) projects, the support device (1) comprising: - a housing (2) configured to radially receive a mounting foot (23) of the stator blade (22); - an upstream support foot (5) and a downstream support foot (6) connected respectively to an upstream end (3) and a downstream end (4) of the housing (2) and configured to be mounted on the external wall (21) of the aircraft turbomachine; the downstream support foot (6) comprising a cavity in which is mounted a shock indicator (10) configured to change its visual appearance irreversibly in the event of deformation of a contour (9) of predetermined shape of the cavity, so as to visually detect an impact exerted on the stator blade (22). Figure from the summary: Figure 2