Abstract: There is provided a pitch control assembly for a rotor of an aircraft, the rotor rotatable about a longitudinal axis and having blades each rotatable about a span-wise axis to adjust a blade pitch angle thereof, the assembly comprising a plurality of position markers spaced circumferentially around the longitudinal axis and having lengths along the longitudinal axis which vary monotonically and incrementally, at least one sensor configured for producing at least one sensor signal in response to detecting a relative movement between the plurality of position markers and the at least one sensor, and a control unit communicatively coupled to the at least one sensor and configured to generate a feedback signal indicative of the blade pitch angle in response to the at least one sensor signal received from the at least one sensor.

Abstract: Systems and methods for identifying and mitigating gas turbine component misalignment using virtual simulation are disclosed herein. An example method may include capturing data associated with a first nozzle segment and a second nozzle segment of a gas turbine. The method may also include creating, based on the captured data, a virtual representation of the first nozzle segment and the second nozzle segment. The method may also include determining that a misalignment exists in a connection between the virtual representation first nozzle segment and the virtual representation of the second nozzle segment. The method may also include identifying, based on the determination that the misalignment exists, a third nozzle segment. The method may also include determining that a connection between a third nozzle segment and the first nozzle segment includes a smaller misalignment.

Abstract: This patent discloses improvements over a mechanism to modulate the pitch of the blades of a fan, turbine or propeller driven by a rotating shaft bearing a helical thread onto which a nut made of a ferromagnetic material was threaded; this nut was spun by electromagnets disposed about it on the stator system and its resultant axial motion manipulated the blades' pitch. This patent discloses refinement of control of the electromagnets employed to spin the axial modulator by embedding a plurality of ferrous protuberances in this nut—the rotational position of which is determined via metallic edge detection sensors whose logic levels are used to selectively energize electromagnets that apply a force onto the protuberances in order to induce rotation and axial motion on the nut. This additional degree of rotational freedom is controlled by this reluctance-motor like configuration while a second motor is used to spin the blades.

Abstract: A turbo blower operable in a surge area and, more particularly, a turbo blower operable in a surge area is provided. The turbo blower increases consistency of performance thereof and efficiency by preventing suspension thereof due to a temporarily generated surge by operating even in a surge area for a predetermined time in addition to a normal area in which the turbo blower normally operates.

Abstract: Provided are a flap drive device and a rotorcraft blade, and more particularly, a flap drive device using a two-section link mechanism, which may be applied to a rotorcraft blade, and a rotorcraft blade including the same.

Abstract: A containment system for an engine includes an engine case having an inner perimeter. The containment system includes a containment ring nested within the inner perimeter of the engine case and integrally formed with the engine case along a first interface and a second interface. The containment ring includes a first leg opposite a second leg, and the first interface is defined between the first leg and the engine case. The containment system includes a first plurality of perforations defined at the first interface, and the first leg of the containment ring is frangible along the first plurality of perforations to at least partially release the containment ring to protect the engine case during a containment event.

Abstract: A method for manufacturing an outer skin of a rotor blade includes forming an outer skin layer of the outer skin from a first combination of at least one of one or more resins or fiber materials. The method also includes forming an inner skin layer of the outer skin from a second combination of at least one of one or more resins or fiber materials. More specifically, the first and second combinations are different. Further, the method includes arranging the outer and inner skin layers together in a stacked configuration. In addition, the method includes joining the outer and inner skin layers together to form the outer skin.

Abstract: A gear system for a rotary wing aircraft includes a first gear including an inner gear portion having a first pitch and an outer gear portion having a second pitch that is distinct from the first pitch. A second gear includes a third pitch driven at a driving rotational speed. An actuator member is coupled to the second gear while driven at the driving rotational speed. The actuator member shifts the second gear along the first gear to engage with one of the inner gear portion and the outer gear portion to drive the first gear at a first rotational speed and engage with other one of the inner gear portion and the outer gear portion to drive the first gear at a second rotational speed. A controller is operable to control the actuator member to shift the second gear between the inner gear portion and the outer gear portion.

Abstract: A cooling structure for a trailing edge of a turbine blade is provided. The cooling structure for the trailing edge of the turbine blade comprising an airfoil shaped blade part including a leading edge, a trailing edge, a pressure surface and a suction surface connecting the leading edge and the trailing edge, and a cavity channel formed in the blade part and through which a cooling fluid flows, the cooling structure including slots and lands arranged alternately on the trailing edge along a span direction of the pressure surface by cutting a portion of the pressure surface, the slots communicating with the cavity channel and defined by adjacent lands where the pressure surface remains, wherein a pin-fin structure is disposed in the cavity channel on an upstream side of the slot, and wherein the cooling fluid is introduced through a micro-channel formed inside the pin-fin structure and is discharged through film cooling holes formed in the pressure surface.

Abstract: A propeller provided with interchangeable blades, and to a method of mounting such interchangeable blades on the propeller. The propeller comprises a hub, fastening fittings, rotation guide devices for guiding the fastening fittings in rotation relative to the hub, locking parts, and blades. The fastening fittings are configured to be assembled into the hub from the inside of the hub, and the locking parts lock respective ones of the fastening fittings onto the hub. The blades are fastened to respective ones of the fastening fittings outside the hub and can thus easily be replaced with other blades without removing the propeller.

Abstract: Low latency pitch adjustable rotors are disclosed. A disclosed example rotor includes a rotor hub to rotate about a rotational axis, rotor blades coupled to the rotor hub, the rotor blades being pitch adjustable and having corresponding pitch angles, and a reaction hinge operatively coupled between the rotor hub and the rotor blades, the reaction hinge to move relative to the rotor hub in response to an angular acceleration or deceleration of the rotor hub to adjust the pitch angles.

Abstract: A turbine blade includes an airfoil body, and a plurality of cooling passages extending along a blade height direction inside the airfoil body and being in communication with each other to define a serpentine flow passage. The plurality of cooling passages include first turbulators on an inner wall surface of an upstream cooling passage of the plurality of cooling passages, and second turbulators on an inner wall surface of a downstream cooling passage of the plurality of cooling passages. A second angle formed by the second turbulators with respect to a flow direction of a cooling fluid in the downstream cooling passage is smaller than a first angle formed by the first turbulators with respect to the flow direction of the cooling fluid in the upstream cooling passage.

Abstract: The invention relates to a side-channel compressor (1) for a fuel cell system (37) for conveying and/or compressing a gas, particularly hydrogen, comprising a housing (3), the housing (3) comprising a housing upper part (7) and a housing lower part (8), a compressor chamber (30) located in the housing (3), comprising at least one peripheral side channel (19), a compressor wheel (2) arranged in the housing (3), which is rotatably arranged about an axis of rotation (4), the compressor wheel (2) comprising blades (5) arranged on the periphery thereof in the region of the compressor chamber (30), and respectively a gas inlet (14) embodied in the housing and a gas outlet (16) which are fluidically interconnected by means of the compressor chamber (30), particularly the at least one side channel (19).

Abstract: A rotor assembly for a gas turbine engine includes a turbine shaft disposed about a longitudinal axis, a first rotor and a second rotor configured for rotation about the longitudinal axis, and an intermediate shaft positioned radially between the turbine shaft and the second rotor. The second rotor is mounted to and axially adjacent the first rotor. The intermediate shaft is mounted to the turbine shaft on an inner radial side of the intermediate shaft. The intermediate shaft is mounted to the second rotor on an outer radial side of the intermediate shaft.

Abstract: A structure for assembling turbine blade seals, a gas turbine including the same, and a method of assembling turbine blade seals are provided. The structure for assembling turbine blade seals includes a turbine blade including an airfoil, a platform, and a root, a turbine rotor disk to which the root of the turbine blade is mounted, a seal plate mounted between the platform and one side of the turbine rotor disk to seal a cooling channel defined within the root and the platform, an insertion pin inserted through the turbine rotor disk and the seal plate to fix the seal plate to the turbine rotor disk, and a retainer configured to fix the insertion pin and to prevent the insertion pin from falling out.

Abstract: The multi-staged cowl described herein allows to increase and maximize water mass flow and pressure drop at the runner cross-section of a hydrokinetic turbine so as to maximize produced power output, while respecting dimensional constraints provided by a shallow body of water, a river for example, in which the hydrokinetic turbine can be submerged. The multi-staged cowl described herein can thus be configured so as to allow water to flow through the hydrokinetic turbine at a substantially stable water mass flow, eliminating instability, avoiding vortices, minimizing cavitation and avoiding fluid separation to negligible levels, and can include an inlet, an outlet and multiple stages which can extend between the inlet and the outlet, so that water can flow therethrough in a water flow direction.

Abstract: An airfoil includes a root and a tip, which define a span of the airfoil therebetween. The airfoil also includes a leading edge and a trailing edge downstream of the leading edge along a flow direction. The leading edge and the trailing edge each extend across the span of the airfoil from the root to the tip. The airfoil further includes a pressure side surface and a suction side surface. The airfoil also includes a thermal barrier coating on the pressure side surface and the suction side surface. The thermal barrier coating includes a base layer and a top coat. A thickness of the base layer varies across each of the pressure side surface and the suction side surface with a maximum thickness of the base layer at the leading edge.

Abstract: A vertical axis turbine includes a rotatable hub assembly that is configured to be connected to an energy sink and rotatable about an axis of rotation. At least two blades are mounted on the hub assembly, each blade including a leading edge and a trailing edge, the blades being oriented so that the respective leading edges face in a common rotational direction. Each blade further includes a straight section that is substantially parallel to the axis of rotation and two helical sections, the straight section being interposed between the helical sections, and the helical sections extending at least partially around the axis of rotation.

Abstract: Impingement insert for an airfoil of a blade/vane of a gas turbine is provided. The impingement insert includes a double-walled section having an outer and an inner walls, that define—an inner channel at an inner surface of the inner wall, an outer channel at an outer surface of the outer wall and a middle channel between the outer and the inner walls. Impingement cooling holes are provided in the outer wall that use the cooling air of the middle channel to eject impingement jets into the outer channel. The impingement insert includes at least one extraction duct that extends between the outer and the inner walls across the middle channel, and has an inlet at the outer channel, and an outlet at the inner channel, for flowing the cooling air, after impingement, from the outer channel into the inner channel.

Abstract: A rotor blade system having a plurality of rotor blades, wherein at least one of the rotor blades includes an outer surface having generally opposing first and second surfaces, the rotor blade including a fluid flow altering surface positioned relative to one of the first or second surfaces which is moveable between first and second positions, wherein movement of the fluid flow altering surface is effected by an expandable member.