Abstract: A system is provided for an aircraft. This aircraft system includes a gas turbine engine and a sensor system. The gas turbine engine includes an inlet and a compressor section. A flowpath projects radially inward into the gas turbine engine from the inlet and extends through the compressor section. The sensor system includes a plurality of static pressure sensors at least partially within the flowpath. The sensor system is configured to determine a total pressure characteristic within the flowpath using the plurality of static pressure sensors.

Abstract: A stator vane provided rearward of a rotor blade, includes: an airfoil body having an airfoil profile. The maximum airfoil thickness position of the airfoil body on the airfoil profile satisfies following conditions at least on a tip side of the airfoil body: (a) on a plane expanded in a circumferential direction in which the stator vane is arranged, the position is located in a first region which is close to a trailing edge of the airfoil body from an intersection of the airfoil body and a line which is parallel to an extension line of a camber line of the rotor blade at a trailing edge of the rotor blade and passes through a leading edge of another stator vane adjacent in the circumferential direction, and (b) the position is located in a second region having a chord ratio from 0.2 to 0.8.

Abstract: Disclosed is a compressor of a dual-flow aircraft turboshaft engine extending longitudinally along an axis X, the compressor comprising a low-pressure compressor, a high-pressure compressor and an intermediate casing axially connecting the low-pressure compressor and the high-pressure compressor, the high-pressure compressor comprising an inlet guide stator comprising a retaining ring, an outer high-pressure compressor casing and a plurality of vanes, the compressor comprising an immobilising device comprising at least one first member secured to the intermediate casing and at least one second member secured to the retaining ring and configured to cooperate with the first member in order to prevent tangential movement while allowing axial movement and radial movement of the retaining ring relative to the intermediate casing according to the axis X.

Abstract: A variable-pitch stator vane for a turbine engine compressor stator, extending in a longitudinal radial plane including a radially inner plate and a radially outer plate between which at least one blade extends, each plate including a transverse longitudinal wall facing the other plate, characterised in that the transverse longitudinal wall of at least one plate includes at least one fin protruding radially towards the other plate.

Abstract: An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil body extending between leading and trailing edges and extending from a root section, and the airfoil body defining pressure and suction sides. The airfoil body defines a recessed region extending inwardly from at least one of the pressure and suction sides, and the airfoil body includes one or more ribs that define a plurality of pockets within a perimeter of the recessed region. A plurality of cover skins is welded to the airfoil body along the one or more ribs to enclose respective ones of the plurality of pockets. The plurality of cover skins are formed from a common cover having a perimeter that is dimensioned to mate with the perimeter of the recess.

Abstract: A propeller tilting apparatus of air mobility may include a guide bracket configured to extend in forward and rearward directions; a sliding mechanism provided in the guide bracket to be movable in the forward and rearward directions, and including a driving motor configured to generate rotation force, and a power conversion unit connected to the guide bracket and configured to convert the rotation force of the driving motor into rectilinear motion of the sliding mechanism; and a tilting mechanism engaged to the power conversion unit to be moved together with the sliding mechanism, and including a propeller motor configured to be tilted in the upward and downward directions, and an angle conversion unit connected to the propeller motor and the power conversion unit to tilt the propeller motor using the rotation force transmitted from the power conversion unit.

Abstract: A nozzle assembly for a gas turbine engine and methods for assembling a nozzle assembly are provided. In one example aspect, the nozzle assembly includes an outer wall and an inner wall radially spaced from the outer wall. The outer wall defines a plurality of mounting openings spaced circumferentially from one another. The inner wall defines a plurality of mounting openings spaced circumferentially from one another. The mounting openings defined by the inner wall are positioned circumferentially between adjacent mounting openings defined by the outer wall. The nozzle assembly includes vanes that are inserted through the mounting openings of the outer wall in a radially inward direction and vanes that are inserted through the mounting openings of the inner wall in a radially outward direction in an alternating manner.

Abstract: A cable holder, in particular for a cable of a wind turbine, to a cable harness, to a tower, to a wind turbine and also to a method for fastening a cable. In particular, a cable holder, in particular for a cable of a wind turbine, preferably for a medium-voltage cable connected to a medium-voltage transformer of a wind turbine, comprising a cable mount with a funnel-like cavity and also comprising an elastic insert, which is arranged within the funnel-like cavity, wherein the cable mount and the elastic insert are arranged and designed such that the cable can extend through the funnel-like cavity and the elastic insert can be clamped in between the cable mount and the cable.

Abstract: A ducted-rotor aircraft includes a fuselage and first and second ducts that are coupled to the fuselage. Each duct includes a duct ring, a rotor having a plurality of blades, a hub that positions the rotor such that the blades define a blade plane of rotation within the duct ring, and a plurality of stators that are coupled to the hub at respective locations aft of the blade plane of rotation. Each of the plurality of stators defines a leading edge that is slanted toward the blade plane of rotation. The leading edges of the stators are slanted to follow a contour defined by the blades. The leading edges may also be slanted to maintain a distance of at least one blade inboard chord length between the leading edges of the stators and respective trailing edges of the blades.

Abstract: A gas turbine engine includes a first component, a second component, an anti-rotation tab, and a liner. The first component has a first wall that includes a first slot, and the second component has a second wall that includes a second slot that is in register with the first slot to define a groove. The anti-rotation tab extends into the groove and limits rotation of the first component and the second component. The liner lines the groove to limit wear between the anti-rotation tab and the first component and between the anti-rotation tab and the second component.

Abstract: A guide vane assembly includes a fan case, a vane, and a retention ring. The fan case extends circumferentially partway about an axis. The vane extends radially inward away from the fan case and is coupled to the fan case. The retention ring extends circumferentially partway about the axis and is coupled with the fan case and supports the vane.

Abstract: An exterior nozzle member for a turbomachine extending along an axis oriented from upstream to downstream and comprising a primary duct configured to conduct a primary flow, an annular cavity of axis and a turbine, said exterior nozzle member comprising: an acoustic attenuation panel formed by an acoustic attenuation structure on which there are mounted an interior wall and an exterior wall facing the primary duct and the annular cavity, respectively, a nozzle flange connected to the panel and comprising an upstream end configured to be connected to a turbine casing flange and a sealing member connected to the panel and comprising an interior longitudinal branch extending radially inside the nozzle flange in order to thermally protect it from the primary flow, the nozzle flange being made of a first material and the sealing member being made of a second material, the first material having a coefficient of thermal expansion greater than that of the second material.

Abstract: A vane for a gas turbine engine, the vane including a platform with an airfoil extending radially from the upper surface of the platform. The platform includes a joint portion which includes a circumferentially extending flange and a recessed surface both formed on either the upper or lower surface of the platform. The flange and the recessed surface extend from opposing circumferential edges of the joint portion and each include a substantially radially-extending through hole.

Abstract: A composite platform for an aircraft turbine engine fan includes a wall of elongate shape that is configured to extend between two fan blades. The wall has an aerodynamic external face and an internal face on which is disposed a fixing tab configured to be fixed to a fan disc. A method for manufacturing the composite platform includes the steps of: a) producing a preform by three-dimensionally weaving of fibers, b) unbinding some of the fibers of the preform to detach at least one longitudinal layer of fibers from the rest of the preform, c) inserting a metal reinforcement between this layer and the rest of the preform, and d) injecting a resin into the preform so as to form said wall and secure the reinforcement to this wall.

Abstract: A hybrid electric propulsion system including: a gas turbine engine comprising a low speed spool and a high speed spool, the low speed spool comprising a low pressure compressor and a low pressure turbine, and the high speed spool comprising a high pressure compressor and a high pressure turbine; an electric motor configured to augment rotational power of the high speed spool or the low speed spool; at least one blade outer air seal positioned between an outer case of the high pressure turbine and a plurality of blades of the high pressure turbine; a clearance control system operably coupled to the at least one blade outer air seal, the clearance control system configured to vary a position of the at least one blade outer air seal with respect to the plurality of blades of the high pressure turbine; and a controller operably coupled to the electric motor and the clearance control system, wherein the controller is configured to operate the clearance control system based upon an operational state of the electri

Abstract: A fan blade platform assembly includes a pressure side platform half that includes a first attachment wall, a first gas path wall, and a first connecting wall attached to both the first attachment wall and the first gas path wall. A suction side platform half includes a second attachment wall, a second gas path wall, and a second connecting wall attached to both the second attachment wall and the second gas path wall. A seal engages the pressure side platform half and the suction side platform half.

Abstract: A regenerative blower-compressor includes an impeller mounted to a drive shaft within a housing including a channel extending from an inlet adjacent to a low fluid-pressure region of the channel to an outlet adjacent to a high fluid-pressure region of the channel, the impeller extends radially outward through an annular volume within the housing from the drive shaft to blades in the channel and is configured to rotate for rotating the blades through the channel for forcing fluid through the channel from the inlet to the outlet in response to rotation of the drive shaft, the drive shaft extends from the impeller within the annular volume to into a shaft chamber within the housing configured to receive fluid from the high fluid-pressure region of the channel, and a port configured to vent fluid directly from the shaft chamber to into the low fluid-pressure region of the channel.

Abstract: The invention relates to a straightening assembly (28) comprising two radially inner and outer coaxial shells (34) between which extend vanes (36) made of composite material, fixed to a first end portion on the radially inner shell and to a second end portion on the radially outer shell, characterized in that: for each vane (36), in a plane perpendicular to the axis of the radially inner and outer shells (34), a straight line passing through a junction between said first and second end portions and the useful part forms an angle ? with a radius of the radially inner shell, passing through the junction between said first end portion and the useful part of said vane (36), such that 0°<??30°; the radially inner shell has a diameter ranging from 1,000 mm to 1,600 mm; and in that the radially outer shell has a diameter ranging from 2,000 to 2,800 mm; the number of vanes (36) ranges from twenty-five to forty-five.

Abstract: A turbine comprises a casing, an outer metal shroud, an inner metal shroud and an annular distributor having a plurality of CMC ring sectors, each sector comprising a mast, an inner platform, an outer platform and at least one blade having a hollow profile that defines an inner housing, the inner and outer platforms each having an opening communicating with said inner housing, and the mast passing through said openings and the inner housing and being secured to said casing and connected to said annular sector. Each blade comprises at least one first radial shoulder projecting axially towards the inside of the blade, and each mast comprises at least one second shoulder projecting axially towards the outside of the mast configured to radially cooperate with a first shoulder and radially press the blade against the mast.

Abstract: Disclosed herein are a grooved turbine blade, and a turbine and a gas turbine including the same. According to the disclosure, since a groove part is formed on a root member to disperse torsional stress, it is possible to improve durability of a rotor disk to which the turbine blade is assembled.