Abstract: An airfoil includes a platform and an airfoil section that are formed of a ceramic matrix composite that includes core fiber plies, skin fiber plies, platform fiber plies, and a cooling passage that extends through selected ones of the plies to provide a cooling circuit through the airfoil section and into the platform.

Abstract: The present invention belongs to the technical field of gas turbine engines used as propulsion systems for supersonic aircraft. In particular, the invention relates to thrust vectoring convergent-divergent nozzles and, more in particular, to an actuation mechanism for vectoring said variable geometry nozzle.

Abstract: A casting core assembly for a gas turbine engine component according to an implementation includes a skin core corresponding to a first cooling passage of an airfoil. The first cooling passage includes a first section and a tip flag section joined at a first bend. The skin core includes a first portion corresponding to the first section and a tip flag portion corresponding to the tip flag section. The skin core includes at least one arcuate slot corresponding to at least one turning vane of the airfoil. A method of forming an airfoil for a gas turbine engine is also disclosed.

Abstract: An adaptive surface texturing method is provided for use with a part formed of ceramic matrix composites (CMCs) to be coated with an environmental barrier coating (EBC). The method includes coating a CMC surface of the part with an initial coating, determining a contour and an undulation pattern of the CMC surface and designing a texturing pattern to follow the contour and the undulation pattern and to be formed in the initial coating based on a thickness of the initial coating and an average particle size of a slurry of the EBC.

Abstract: A rotor system for a turbine engine. The rotor system includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades that rotate. The stator assembly includes a plurality of stator vanes arranged circumferentially about the stator assembly and includes at least one pair of non-uniform gaps between adjacent stator vanes. The plurality of stator vanes includes a first group of stator vanes having a first non-uniform gap between adjacent stator vanes, a second group of stator vanes having a second non-uniform gap between adjacent stator vanes, and a third group of stator vanes having a uniform spacing between adjacent stator vanes. The first non-uniform gap is positioned 180° from the second non-uniform gap. The plurality of rotor blades directs air through the plurality of stator vanes.

Abstract: An air supply system has an air intake that provides air to an air vehicle engine. An inlet port supplies air to the air intake. First apertures on the surface of the air intake provide an uninterrupted air supply when the inlet port is blocked. A rotary cylinder is mounted to be rotatable around the air intake. The first apertures and second apertures that are on the surface of the rotary cylinder form an open position so air is introduced into the air intake and a closed position in which air supply to the air intake is blocked. An actuator triggers a first gear on the rotary cylinder to rotate around the air intake of the rotary cylinder. A second gear provides motion transmission between the actuator and the first gear. A sensor unit measures temperature, pressure, humidity and/or flow values of air passing through the air intake.

Abstract: A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft.

Abstract: The present disclosure provides an electric propulsor. The electric propulsor has a propeller having a hub and blades extending from the hub. The blades are movable along respective flapping axes between a deployed position and a stowed position. The electric propulsor also has an actuator and an electric motor coupled with the hub and arranged to rotatably drive the propeller. Further, the electric propulsor has one or more processors configured to: receive an input indicating an azimuthal position of the blades; control the electric motor to align the blades with a target azimuthal position based at least in part on the input; and with the blades of the propeller arranged at the target azimuthal position, cause the actuator to translate a control rod coupled thereto and with the hub to actively fold or unfold the blades along the respective flapping axes of the blades.

Abstract: A system for controlling the pitch setting of a propeller blade for an aircraft turbine engine, includes a blade having a vane connected to a root. A hub accommodates the root of the blade, and a ring is mounted around the root and in the hub. Stops are also mounted around the root and in the hub. A nut is threadedly engaged with a thread of the ring and is configured to bear axially against the hub to secure the assembly.

Abstract: An air inlet of a nacelle of an aircraft propulsion assembly assembly comprising an inner wall and an outer wall which are connected upstream by an air inlet lip. The air inlet lip comprising thick portions and thin portions distributed alternately over the circumference of the air inlet lip. Each thick portion comprising an upstream leading edge to separate an upstream air flow into an outer air flow guided by the outer wall and an inner air flow guided by the inner wall during a thrust phase. Each thin portion comprising a downstream leading edge situated longitudinally downstream of each upstream leading edge so as to detach a reverse air flow at the air inlet lip during a thrust reversal phase.

Abstract: A fluid duct includes a body and an elongate rib. The body extends along a central axis and defines a lumen configured to receive a first fluid extending therethrough. The body has an inner surface and an outer surface. The central axis defines an axial direction along the central axis, a radial direction perpendicular to the central axis, and a circumferential direction oriented rotationally about the central axis. The elongate rib is positioned on the inner surface of the body and extends radially inward from the inner surface of the body. The elongate rib defines a channel extending through the elongate rib in a primary direction. The channel is configured to receive a second fluid and isolate the second fluid from the first fluid.

Abstract: A measuring rake, configured to be arranged on a link rod, includes a sheath including a front face and two side walls delimiting between them a recess configured to receive the link rod, an electronic circuit arranged on the front face of the sheath and including at least one sensor, a patched leading edge fixed removably to the sheath, and a seal arranged between the patched leading edge and the electronic circuit, the patched leading edge including a plurality of air intakes, each forming a fluidic passage between an outer face and an inner face of the patched leading edge opening out facing at least one sensor of the electronic circuit, the measuring rake making it possible to obtain a measuring tool which can be assembled simply and quickly, with easy access to the electronic circuit, and making it possible to avoid problems of orifice blockages or air leaks.

Abstract: A turbine engine includes a core having a fan section, a compressor section, combustion section, and turbine section in serial flow arrangement and defining an engine centerline. The turbine section has a set of turbine stages, having paired sets of non-rotating, circumferentially spaced vanes defining a respective nozzle, and rotating circumferentially spaced blades. A first stage turbine nozzle has a solidity that is a ratio of an axial chord length of the turbine vanes to the circumferential spacing between adjacent vanes.

Abstract: A layout for asymmetric trip strips including a flow passage having a lower wall and an upper wall opposite the lower wall, each of the lower wall and the upper wall including an inner surface, the flow passage having a passage inlet and a length L and a diameter d; multiple skewed trip strips extending from at least one inner surface of the lower wall or the upper wall; and at least one periodic reflection of the skewed trip strips along the flow passage downstream of the passage inlet at a frequency with a length-to-diameter ratio of L/d?20.

Abstract: A rotor blade including: a blade body extending in a blade height direction that is a direction perpendicular to an airfoil cross-section; and a platform provided to an end at a hub side, from between a tip side and the hub side which are of the blade height direction. The blade body includes: an anterior edge and a posterior edge; a ventral surface and a dorsal surface linking the anterior edge and the posterior edge together; and a cooling passage extending in the blade height direction between the ventral surface and the dorsal surface. The ventral/dorsal wall thickness ratio at any position between the hub-side position and the middle position in the cooling passage is smaller than the ventral/dorsal wall thickness ratio at any position between the middle position and the tip-side position.

Abstract: A gas turbine engine is provided. The gas turbine engine includes a stage of airfoil segments having a first ceramic matrix composite (CMC) airfoil segment, the first CMC airfoil segment includes: an end band having a radial surface defining a cold side surface area, ACS; and an airfoil coupled to or formed with the end band, the airfoil and the end band defining a joint therebetween, the joint including one or more radial interfaces defining a total radial interface surface area, ARI, and one or more axial interfaces defining a total axial interface surface area, AAI. The first CMC airfoil segment defines a first ratio of the cold side surface area, ACS, to the total axial interface surface area, AAI, and a second ratio of the total axial interface surface area, AAI to the total radial interface surface area, ARI.

Abstract: A ventilation ring for a bearing support member of an aircraft turbine engine includes two or more tubular walls and one or more spacer walls. The two or more tubular walls extend opposite each other, and one or more of the tubular walls includes a main portion and a base, the base having a thickness greater than a thickness of the main portion. The one or more spacer walls connect the two or more tubular walls. The one or more spacer walls have one or more apertures extending into the base and a rim around the aperture. An entirety of the rim has a thickness greater than a thickness of the spacer wall at a distance from the rim.

Abstract: A turbine engine comprising an engine core having a rotor and a stator, and a floating seal assembly sealing at least portions of the rotor and the stator relative to a higher-pressure area and a lower-pressure area. The floating seal assembly can comprise a carriage assembly at least partially defining a seal cavity and a seal body floating within the seal cavity. The floating seal assembly can further include at least one of a seal located between the seal body and the carriage assembly, and a seal face located between the seal body and the carriage assembly.

Abstract: A containment structure for a rotor includes a shroud and a shroud reinforcement. The shroud is coaxial with and partially surrounds the rotor and includes a tubular section, a transition section, and a flange section. The tubular section extends axially past a first side of the rotor. The transition section connects to the tubular section and is adjacent to a curved side of the rotor. The flange section connects to the transition section opposite the tubular section. The flange section extends radially past a radially outer side of the rotor. The shroud reinforcement is connected to a radially outer surface of the transition section. The shroud reinforcement encloses the transition section and includes a support scaffold and a reinforcing material. The support scaffold includes a series of geometric retaining features encircling a radially outer surface of the transition section. The reinforcing material couples to the support scaffold and restricts shroud radial expansion.

Abstract: A de-aerator is provided that includes a body, a cover panel, a fluid inlet, a helical fluid passage, at least one partition, and at least one fluid outlet. The body extends between opposing first and second axial ends. The body has a sidewall, a base panel, and an internal cavity. The fluid inlet is in communication with the body and is configured to direct fluid tangentially into the internal cavity. The helical fluid passage is disposed within the internal cavity and has entry and exit ends. The helical fluid passage has circumferential turns that each include an air passage. The at least one partition is disposed within the internal cavity at or below the passage exit end, spaced above the base panel. The at least one fluid outlet is configured to permit liquid passage from the internal cavity of the body to outside the body.