Abstract: A mast section (1) for a wind turbine having a longitudinal central axis (L) extending along a longitudinal direction includes at least two tubular mast elements (14) stacked along the longitudinal direction and arranged edge-to-edge at a joining plane (P). Each mast element (14) includes at least two wall segments (16), connected to one another by segment connectors (26) extending along longitudinal edges of the wall segments (16). The mast section (1) further includes element connectors (37) each extending astride the two mast elements (14), in the extension of the segment connectors (26) in the longitudinal direction. The mast section (1) further includes a plurality of element overconnectors (45), each element overconnector (45) extending astride a segment connector (26) and an adjacent element connector (37) in the longitudinal direction.

Abstract: The present application discloses a power generation device so as to solve the generator set overspeed problem. The power generation device comprises: a stand column; and at least one generator set located on the stand column. The generator set comprises a support, blades connected to the support, and a power generator generating power by means of rotation of the blades, and an adjustment device located on the support and used for adjusting a windage area of the blades by moving or rotating the blade according to a wind speed. The wind power generation device may reduce an effective windage area to zero when the wind force is too high, thereby improving stability and applicability in a changeable environment and prolonging the service life of the apparatus.

Abstract: A rotor blade airfoil of a turbomachine, which rotor blade airfoil has: a leading edge, a trailing edge, and a profile chord length which is dependent on the height of the blade airfoil. In a side view of the blade airfoil, a maximum projected chord length his defined as the axial spacing between the axially foremost point of the leading edge and the axially rearmost point of the trailing edge of the blade airfoil in the side view under consideration. Here, the axial position of the leading edge varies in a manner dependent on the height of the blade airfoil above a front axial region.

Abstract: A planetary gearing includes at least a first planetary stage and a second planetary stage and a double-walled pipe. The double-walled pipe has an inner pipe and an outer pipe and is provided with an outlet point in a central section. A bushing is arranged on the outlet point.

Abstract: A turbomachine engine can include an unducted fan assembly, an unducted vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The unducted fan assembly can include a plurality of fan blades. The unducted vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the unducted fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.

Abstract: A propulsion assembly for an unmanned aerial vehicle (UAV) includes a motor configured to rotate in a first direction, a propeller seat configured to be driven by the motor to rotate in the first direction and to receive a propeller, and a sensor configured to collect sensing data useful for determining whether the propeller is locked to the propeller seat, without requiring operation of the motor.

Abstract: This disclosure relates to a compressor having at least two compression stages. In particular, an exemplary compressor includes a first radial compression stage arranged along an axis, a second radial compression stage arranged along the axis, and a connector fluidly connecting an outlet of the first radial compression stage to an inlet of the second radial compression stage. The connector has a plurality of sections arranged about the axis. The compressor may be a refrigerant compressor used in a heating, ventilation, and air conditioning (HVAC) chiller system.

Abstract: An annular parts assembly for mounting onto a shaft of an aircraft engine is provided. The assembly comprises a first annular body having a surface defining a plurality of pulling features extending from a remainder of the surface, the pulling features circumferentially spaced apart on the surface. The assembly comprises a second annular body defining a balancing ring, the balancing ring concentric with the first annular body, the balancing ring having a plurality of protrusions and circumferential spaces between adjacent ones of the plurality of protrusions, the circumferential spaces accommodating the pulling features such that the pulling features of the first annular body and the protrusions intercalate.

Abstract: A process for cold working of the inner surface of a bore in a centrifugal compressor wheel along only a fractional portion of the bore length (i.e., along less than a full axial length of the bore), thereby creating a zone of compressive residual hoop stress in the metal surrounding the bore where the wheel needs the beneficial residual stress. The process purposefully avoids cold working of the bore at locations adjacent to high-stress areas and features of the wheel, where cold working in such locations could negatively impact the wheel's overall life.

Abstract: A turbomachine engine can include a fan assembly, a pitch change mechanism, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly can include a plurality of fan blades. The pitch change mechanism can be coupled to the fan assembly. The vane assembly can include a plurality of vanes. The core engine can include one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.

Abstract: An oil system for a gas turbine engine and a method of supplying oil to the system. The oil system includes a main oil tank connected by oil lines with a supplementary oil storage tank, which has an actuator, and that are connected to one oil pump for supplying oil to the gas turbine engine. The supplementary oil storage tank is equal in size or larger than a steady state oil gulp of the system. The method includes supplying oil from a main oil tank through a pipe line using an oil pump, detecting the oil level in the oil system and determining if additional oil is required or requires removing using a sensor and an electronic controller, and transmitting a signal to an actuator to supply or remove oil to and from the pipe lines in the oil system from or into a supplementary oil storage tank.

Abstract: A method of manufacturing an impeller includes attaching blades to a hub. The impeller includes the blades, the hub, and a shroud. The blades each include a tip, and the shroud includes an inner surface and at least one ring extending from the inner surface. The method also includes applying a brazing compound to the tips of the blades and to the inner surface of the shroud. The method includes inserting the blades and the hub into the shroud such that the tips of the blades press against the rings. The rings and the tips of the blades form an interference fit between the rings and the tips of the blades that maintains a consistent gap between the shroud and the blades during manufacture of the impeller. A compressor including the impeller is also disclosed.

Abstract: A unique seal for a wave rotor disk engine is disclosed herein. The seal is operable for sealing a region between a rotor and a rotor casing. The seal is spring loaded and will engage with the tip of the rotor to reduce pressure loss within the wave rotor disk engine.

Abstract: Provided is a turbo compressor. The turbo compressor includes a driving shaft, a first impeller, a second impeller, a first shroud, a second shroud, a first-stage outflow passage, and a second-stage outflow passage. Also, the turbo compressor includes a gap adjustment passage that is branched from at least one of the first-stage outflow passage or the second-stage outflow passage to extend to the first shroud.

Abstract: A face seal arrangement includes a seal seat rotating with the shaft and a non-rotating face seal including a sealing ring. The sealing ring includes a sealing face biased into contact with the seal seat. The sealing ring also has a groove defined remote from the seal face relative to the seal seat. The groove provides a pressure face that will be exposed to high pressure air outward of the bearing compartment. A coil spring biases the seal housing towards the seal seat, such that the sealing face of the sealing ring is biased into contact with seal seat by air pressure against the pressure face, and the coil spring. A balance ratio is defined between an area of the pressure face and an area of the sealing face, the balance ratio is between 0.5 and 0.64.

Abstract: The cooling fan module may include a fan assembly and a fan shroud. The fan shroud of the cooling fan module may include a first sidewall that may extend in a radial direction and define an opening. The fan assembly may include a hub disposed within the opening, a blade extending from the hub to a distal end that may include a first side and a second side converging to form a first leading edge with respect to the rotational direction. The fan assembly may include a winglet that may include a first fin, extending in the axial direction from the distal end of the blade and a second fin extending in the radial direction.

Abstract: The present disclosure relates to a geared gas turbine engine for an aircraft. Example embodiments include a gas turbine engine for an aircraft including: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; and a gearbox that receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of around 3.4 or higher, wherein the gas turbine engine is configured such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range from around 0.75 to around 0.82.

Abstract: First and second pressure relief passages extend from an oil pan in a branching manner, and merge with each other to form a merging portion. A pressure relief hole is arranged above the merging portion, and a first pressure relief passage is arranged below the merging portion. The minimum cross-sectional area of the second pressure relief passage is smaller than the minimum cross-sectional area of the first pressure relief passage. The second pressure relief passage includes a bent portion formed by bending the second pressure relief passage. The bent portion is configured to perform gas/liquid separation by crushing bubbles. When reaching the merging portion from the bent portion, oil is returned to the oil pan via the first pressure relief passage. When reaching the merging portion from the bent portion, gas is discharged to the outside of the housing via the pressure relief hole.

Abstract: A sealing device between a rotor part and a stator part, including at least one abradable coating cooperating with at least two upstream and downstream labyrinth seal lips. Axially upstream of the labyrinth seal lips, the sealing device includes a circumferential wall that radially extends beyond the upstream free axial sealing surface of the coating to create, at the free end of the upstream labyrinth seal lip, a separation of the circulating gas.

Abstract: A bladed disk preform system is provided. Multiple sheets are formed or provided. Each of the sheets may include multiple ceramic fibers. Slots may be cut in each of the sheets before or after forming. The slots extend from an outer edge of the sheets inward toward a center of the sheets. The sheets are stacked on top of each other to form a bladed disk preform. The bladed disk preform is porous. The bladed disk preform may include a plurality of blade portions extending outward from a root portion. Each of the slots is positioned between adjacent blade portions.