Abstract: A system and methods for a self-rotating fluidic traverse actuator are presented. A turbine rotates in response to a fluid flow, and an outer cylinder with a longitudinal slot that ejects the fluid flow. An inner cylinder rotates inside the outer cylinder in response to a rotation of the turbine and at least one helical slot of the inner cylinder ejects the fluid flow into the longitudinal slot.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan section including a fan with a plurality of fan blades rotatable about an axis. Each of the plurality of fan blades includes a mid-span shroud and a speed change device in communication with the fan.

Abstract: A variable geometry turbine comprise a housing; a turbine wheel supported in the housing for rotation about a turbine axis; and an annular inlet passage upstream of said turbine wheel defined between respective inlet surfaces defined by an annular nozzle ring and a facing annular shroud. The nozzle ring and shroud are axially movable relative to one another to vary the size of the inlet passage; the nozzle ring having a circumferential array of inlet vanes extending across the inlet passage. The shroud covers the opening of a shroud cavity and defines a circumferential array of vane slots, each vane slot corresponding to an inlet vane, the vane slots and shroud cavity being configured to receive said inlet vanes to accommodate axial movement of the nozzle ring.

Abstract: An exhaust gas turbocharger (1) with a compressor (2) which has a compressor wheel (3) arranged in a compressor housing (4); with a turbine (5) which has a turbine wheel (6) arranged in a turbine housing (7); and with a bearing housing (8) which is arranged between the compressor housing (4) and turbine housing (7); and which has a bearing (9) for a rotor shaft (10) which bears the turbine wheel (6) and the compressor wheel (3), wherein the bearing (9) has at least one ceramic insert (11, 12) which is on the bearing housing side and is under compressive prestress, and at least one ceramic insert (11, 12) which is on the motor shaft side and is under tensile prestress.

Abstract: The invention relates to a blade (2) of a turbomachine, in particular a rotor blade of a gas turbine, which has a variable transition radius (Rv1) in the vicinity of at least one platform overhang (16), and to a turbomachine having at least one such blade (2).

Abstract: An impeller radial pump includes a pump chamber with a central suction and with a pump chamber outlet, wherein an impeller is provided in the pump chamber. Radially outside of the impeller, a circular ring shaped and circumferential pump chamber ring section is provided as a part of the pump chamber, wherein the pump chamber ring section essentially has an extension along the axial direction of the pump from the impeller against the suction direction. The pump chamber ring section has a varying width and is configured to be narrower in a compression region in the circulation direction.

Abstract: A hollow article includes a metallic hollow article formed from a having a first major surface, an internal cavity with an opening in the first major surface, and a socket around the opening; a cover of composite material received in the socket and covering the opening; and a filler material of foam in the internal cavity.

Abstract: An individual blade control system (IBCS) for a rotor system having a rotor hub is provided including a plurality of blade cuffs mounted to the rotor hub. Each blade cuff is configured to receive a rotor blade and rotate about a blade axis. A plurality of electrical actuators is mounted to the rotor hub adjacent at least one of the plurality of blade cuffs. Each electrical actuator is configured to rotate about an actuator axis. The plurality of blade axes and the plurality of actuator axes are arranged in a plane. Each electrical actuator is coupled to an adjacent blade cuff such that rotation of one of the plurality of electrical actuators causes a proportional rotation of one of the blade cuffs.

Abstract: An airfoil blade of a gas turbine engine includes a root configured for mating attachment with a cooperating rotor hub and an airfoil extending away from the root. The root is composed of a first metal and has a nanocrystalline metal coating, composed of a second metal, over at least a portion thereof. A method of protecting such a blade by applying a nanocrystalline metal coating to a portion of the blade root is also disclosed.

Abstract: An airfoil includes a pressure surface and a suction surface extending from a root section of the airfoil to a tip section of the airfoil. The airfoil also includes a leading edge and a trailing edge defining a chord length of the airfoil therebetween. The airfoil further includes a tip shelf formed along the tip section of the airfoil between the pressure surface and a tip shelf wall. The tip shelf wall is spaced between the pressure surface and the suction surface and the tip shelf extends from within 10% of the chord length measured from the leading edge to within 10% of the chord length measured from the trailing edge.

Abstract: A bimetallic ring for use as a turbine shroud in a gas turbine engine. The bimetallic ring forms a sealing surface as a hot gas flow path boundary in the engine. The ring is comprised of two materials. The first material, a wrought, oxidation resistant metal alloy comprises a first portion, which is the hot gas flow path sealing surface. The second material, a low cost low alloy steel, comprises a second portion that may be at least a pair of supporting side plates. A dissimilar weld joint joins the sealing surface to the second portion, the at least pair of supporting side plates.

Abstract: An axial compressor for compressing a flow medium is provided and includes compressor stator blades, which fastened on a stator blade carrier, and compressor rotor blades, which are fastened on a compressor disk of a compressor shaft, wherein two consecutive compressor disks enclose a hollow space lying between them, and wherein the last compressor disk, as seen in the flow direction of the flow medium, delimits a rear hollow space. A cooling medium feed duct leads through a stator blade carrier and a compressor stator blade, and is arranged upstream of the last compressor disk, which cooling medium feed duct, at the tip end of the compressor stator blade facing the compressor shaft, via a discharge opening arranged there, opens into the hollow space, wherein this hollow space, via a cooling medium transfer passage which is led through the subsequent compressor disks, is connected to the rear hollow space.

Abstract: A control system for a fire fighting device is provided. The control system may include a controller configured to operably connect over a network to a first valve controller controlling a status of a first valve in response to an input from the controller, a second valve controlling a status of a second valve in response to an input from the controller, and a governor operably controlling a power source in response to an input from the controller. The control system may further comprise a touchscreen display operably connected to the controller, the display configured to display the status of the first and second valves and configured to receive an input from a user, wherein controller generates a manipulatable icon displayed on the touchscreen display, a manipulation of the icon generating the input from the controller to the first valve controller.

Abstract: A first stage compressor disk of a gas turbine engine includes a body. The body includes a forward end, an aft end, and an outer surface. The body also includes a plurality of forward balancing holes through the outer surface. The forward balancing holes align circumferentially about the body. The body further includes a plurality of aft balancing holes through the outer surface. The aft balancing holes align circumferentially about the body and are located aft of the forward balancing holes. The first stage compressor disk also includes a radial flange at the aft end of the body. The radial flange extends radially outward from the body. The radial flange includes slots for mounting airfoils.

Abstract: A rotating circular airfoil system that includes a disc body that has an airfoil shape in cross section. The air foil shape has an angle of attack that is symmetrical. The disc body's includes a circular inner edge which forms a circular center opening and the airfoil shape's leading edge. The disc body also includes an outer edge that forms the airfoil shape's trailing edge. Intersecting ribs extend across the center opening and include a shaft opening that attaches to a rotating shaft. Formed on the top and bottom surfaces of the disc body are evenly spaced apart, extending fins. The fins extend continuously from the inner edge to the outer edge. When a rotating shaft is inserted into the shaft opening and imparts sufficient rotation to the disc body, the fins contact the air forcing it to flow along disc body's top and bottom surfaces and creating a lifting force.

Abstract: A laminated abutment (10) comprising an inner strength member (11), an outer strength member (12), and an elastomeric member (15) comprising a stack of main layers (20, 30) that are flexible and of secondary layers (40) that are rigid. The stack comprises at least one first main layer (20) and at least one second main layer (30), each first main layer (20) presenting, in a lead-lag plane (PTRA), a first thickness (e1) that varies between a minimum first thickness (e1min) on the pitch-variation axis (AXPAS) and a maximum first thickness (e1max) on the flapping axis (AXBAT), and each second main layer (30) presenting, in a lead-lag plane (PTRA), a second thickness (e2) that varies between a minimum second thickness (e2min) on the flapping axis (AXBAT) and a maximum second thickness (e2max) on the pitch-variation axis (AXPAS).

Abstract: A casing for an aircraft engine includes an outer ring and an inner hub defining an airflow passage therebetween, the outer ring having an axis defining an axial direction; a plurality of struts arranged in a circumferential array and extending radially from the inner hub to the outer ring to mount the inner hub to the outer ring; wherein the outer ring is defined by a double skin including an axially-extending annular outer skin of sheet metal concentrically surrounding and radially-spaced from an annular inner skin of sheet metal, the outer and inner skins generally parallel to one another, an annular front end ring and an annular rear end ring welded or brazed to the outer and inner skins adjacent respective front and rear edges of the skins to define an annular cavity between them, and the outer ring further comprising a plurality of circumferentially spaced axially-extending ribs interconnecting the outer and inner skins to reinforce the double skins.

Abstract: A rotor blade includes an airfoil having a tip plate that extends across an outer radial end. A rim extends radially outward from the tip plate and surrounds at least a portion of the airfoil and includes a concave portion opposed to a convex portion. A plurality of dividers extend between the concave and convex portions to define a plurality of pockets between the concave and convex portions at the outer radial end. A plurality of cooling passages through the tip plate provide fluid communication through the tip plate to the plurality of pockets. A first fluid passage in at least one divider provides fluid communication between adjacent pockets across the at least one divider.

Abstract: A wind turbine blade (10) for a rotor of a wind turbine (2) is provided with a longitudinally extending flow guiding device (70, 170) attached to the profiled contour. The flow guiding device comprises: a base (90, 190) having a first longitudinal end (91, 191) nearest the root end (16) and a second longitudinal end (92, 192) nearest the tip end (14), a first side (93, 193) nearest the leading edge (18) and a second side (94, 194) nearest the trailing edge (20), as well as a first surface (95, 195) and a second surface (96, 196), the first surface of the base being attached to the profiled contour, and the second surface facing away from the profiled contour. A longitudinally extending substantially plate-shaped element (97, 197) protrudes from the second surface (96, 196) of the base (90, 190).

Abstract: An airfoil comprises pressure and suction surfaces extending from a root section to a tip section of the airfoil. The airfoil also comprises a leading edge and trailing edge defining a chord length therebetween. A tip shelf is formed along the tip section between the pressure surface and a tip shelf wall, the tip shelf wall being spaced between the pressure surface and the suction surface. A squealer pocket is formed along the tip section between the tip shelf wall and a squealer tip wall extending from the suction surface. The tip shelf extends from within 10% of the cord length measured from the leading edge to within 10% of the chord length measured from the trailing edge. The squealer pocket extends from within 10% of the chord length measured from the leading edge to terminate less than 85% of the chord length measured from the trailing edge.