Abstract: A gas turbine engine component includes an inner support structure surrounding an engine center axis and fixed to an engine static structure, an outer support structure spaced radially outward of the inner support structure, and a curved beam comprised of a plurality of curved beam spring segments that are positioned adjacent to each other to form a ring. The inner and outer support structures are coupled together around the curved beam to enclose the curved beam therebetween and form an assembly. A bearing is spaced radially inward of the assembly.

Abstract: A seal assembly adapted for use with a gas turbine engine includes a runner, a seal, and a compliant holder. The runner extends circumferentially about an axis and the compliant holder. The compliant holder engages the runner and is configured to support the runner radially relative to the axis. The seal extends circumferentially around the runner to block fluid flow through the seal assembly.

Abstract: A turbine nozzle for an aircraft turbine engine, the nozzle including at least one vane and a base having a platform. The nozzle is fitted with a passive system for reintroducing blow-by gas into the primary jet, the passive system including gas extraction ports on the base as well as gas reinjection ports on the radially outer surface of the platform and/or on a suction-side surface of the vanes, the gas reinjection ports being designed to redirect the gas flow in a reinjection direction having a circumferential orientation.

Abstract: A turbine vane for a gas turbine engine, including: a first airfoil including leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface extending from an inner platform in a radial direction to an outer platform, wherein a portion of the exterior airfoil surface of the first airfoil is formed in conformance with a set of Cartesian coordinates set forth in Table 1 as offset by corresponding values in Table 4, and wherein the values of Table 4 are offset from a point of origin that is a point where a radially outward surface of the inner platform meets a surface of a trailing edge of the inner platform and a surface of a mate face of the inner platform.

Abstract: In one embodiment, a local control system for a rotor assembly of an apparatus includes a first actuator disposed in the rotor assembly and configured to control motion of a first controllable element in the rotor assembly. The rotor assembly is mounted to the apparatus and is rotated responsive to torque and rotational energy provided thereto. The local control system also includes a first sensor disposed in the rotor assembly and configured to provide position feedback in relation to the first controllable element. The local control system also includes a first local control computer disposed in the rotor assembly and communicably coupled to a first central control computer disposed in the apparatus external to the rotor assembly, where the first local control computer is configured to transmit a control signal to the first actuator and receive a feedback signal from the first sensor.

Abstract: A turbine nozzle for a gas turbine engine includes an outer endwall, and an inner endwall spaced apart from the outer endwall. The turbine nozzle includes at least one airfoil coupled between the inner endwall and the outer endwall, and a compliant joint defined in at least one of the outer endwall and the inner endwall. The compliant joint includes at least one partially-fused seal that is configured to restrict a flow of fluid through the at least one of the outer endwall and the inner endwall and to form a slip-joint above a predetermined threshold stress. The slip-joint is configured to maintain a radial position of the at least one of the outer endwall and the inner endwall.

Abstract: There is disclosed a turbine comprising a stator vane ring 122 downstream of a rotor ring 116 of shrouded rotor blades 118. In vanes of the stator vane ring, an upstream portion of a mean camber line 136 diverges from a profile of the mean camber line at a half-span location, particularly to receive a tip leakage flow over the shrouded rotor blades 118.

Abstract: A turbine assembly for use with a gas turbine engine includes a bladed wheel assembly, a vane assembly, and an inner seal. The bladed wheel assembly is adapted to interact with gases flowing through a gas path of the gas turbine engine. The vane assembly is located upstream of the bladed wheel assembly and adapted to direct the gases at the bladed wheel assembly. The inner seal is configured to block gases from passing around the vane assembly.

Abstract: A gas turbine engine with a lubrication system includes a ball bearing assembly and a rotor circumscribing a rotational axis and journaled within the ball bearing assembly. The gas turbine engine also includes a lubrication system located radially outward from a rotational axis and radially outward and adjacent to the ball bearing assembly, which includes a lubrication channel having an inlet and an outlet and a dispersion cone adjacent to the outlet of the lubrication channel.

Abstract: A sealing arrangement is provided between inner and outer coaxial ring-shaped walls of an aircraft engine. The outer wall has an axial end with a U-shaped cross-section parallel to axis, the opening therein being oriented axially in a first direction. The inner wall has an axial end with a U-shaped cross-section parallel to axis, the opening therein being oriented axially in the opposite direction. The U-shaped end of the outer wall includes a ring-shaped free edge axially engaged in the opening in the axial end of the inner wall. The U-shaped end of the inner wall includes a ring-shaped free edge axially engaged in the opening in the axial end of the outer wall. The walls define therebetween a ring-shaped gas channel having a substantially S-shaped axial cross-section.

Abstract: An engine assembly includes an engine including a component and defining an opening and an interior, the component including a first side and an opposite second side, the second side positioned within the interior of the engine. The engine assembly also includes an inspection tool having a first member including at least one of a receiver or a transmitter and directed at the first side of the component. The inspection tool also includes a second member including the other of the receiver or the transmitter and positioned at least partially within the interior of the engine and directed at the second side of the component to communicate a signal with the first member through the component, the second member being a robotic arm extending through the opening of the engine.

Abstract: A sealing washer for use in a rotor of a gas turbine, sealing being brought about on one side of the sealing washer by an approximately radially oriented bearing face and on the other side by way of a supporting face which is oriented in an inclined manner with respect to the rotor axis. The sealing washer is of split configuration and has at least one washer segment with a circumferential washer section. At the dividing point, a pressing section and a triangular section overlap along a dividing face. Here, the dividing face intersects the bearing face and makes sealing both with the pressing section and with the triangular section possible. In order to also make sealing on the supporting face with the pressing section and the triangular section possible, the triangular section is reduced to at most 0.3 times the cross-sectional area of the washer section.

Abstract: A runner for a hydraulic turbine or pump includes a plurality of blades, each blade being defined by a pressure surface, an oppositely facing suction surface, a leading edge and a spaced apart trailing edge. At least one blade has a device for supplying a flow of oxygen containing gas to the trailing edge of at least one of the blades. The profile of the suction side surface of the blade along a cross section through a point P1 and a point P2 is concave. The point P1 is located on the suction side surface of the trailing edge where an opening is located, the point P2 is spaced apart from the point P1 by less than 3% of the runner outlet diameter D and the point P2 is located upstream of the point P1 on a line perpendicular to the trailing edge starting at the point P1.

Abstract: An engine core including a turbine, compressor, and a core shaft connecting the turbine and compressor; a fan located upstream of the engine core including a plurality of fan blades; and a gearbox. The gearbox is arranged to receive an input from the core shaft and to output drive to the fan to drive the fan at a lower rotational speed than the core shaft. The gearbox is an epicyclic gearbox and includes a sun gear, a plurality of planet gears, ring gear, and planet carrier to the mounted planet gears. The planet carrier has an effective linear torsional stiffness and the gearbox has a gear mesh stiffness between the planet gears and the ring gear. A carrier to ring mesh ratio of: the ? ? effective ? ? linear ? ? torsional ? ? stiffness ? ? of ? ? the ? ? planet ? ? carrier gear ? ? mesh ? ? stiffness ? ? between ? ? the ? ? planet ? ? gears ? ? and ? ? the ? ? ring ? ? gear is greater than or equal to 0.2.

Abstract: An improved stator assembly for use in a gas-turbine engine is disclosed. The stator assembly may comprise a vane, an inner diameter (ID) ring, an outer diameter (OD) ring, a vane disposed between the ID ring and the OD ring, a potting component coupling the vane to at least one of the OD ring or the ID ring, and a potting embedded component disposed within the potting component. The potting embedded component may prevent disbond of the potting component during operation of the gas-turbine engine.

Abstract: A liner for a gas turbine engine is provided that includes an outer wall, an inner wall, and a plurality of structural elements disposed between the inner wall and the outer wall. The outer wall is spaced apart from the inner wall and is disposed radially outside of the inner wall. The outer wall and inner wall form an annular structure having a circumference, and an annular region disposed between the first interior surface of the outer wall and the second interior surface of the inner wall. Each of the structural elements is attached to the outer wall and the inner wall. The outer wall includes at least one inspection port extending through the outer wall configured for access to the annular region disposed between the first interior surface of the outer wall and the second interior surface of the inner wall.

Abstract: This aircraft gas turbine is provided with a rotor, a rotor blade (24), a casing, a stationary blade (26), and a guide member (53). The rotor blade (24) has fins (43) projecting from the outer peripheral surface (42a) of a rotor blade shroud (42). A stationary blade shroud (51) of the stationary blade (26) forms a cavity (Ct) with at least the rotor blade shroud (42) and the fins (43). The guide member (53) is provided inside the cavity (Ct) and extends inward from the radial outer side, and an inner-side end part (53a) on the radial inner side faces the outer peripheral surface (42a) of the rotor blade shroud (42) with a gap therebetween.

Abstract: A vibration suppression device for a rotary machine according to at least one embodiment of the present disclosure is a vibration suppression device for a rotor of a rotary machine including a damper pin movably provided inside a gap of the rotor, the damper pin including a magnet, and a magnetic force generation portion provided in the rotor at a periphery of the gap. The magnetic force generation portion is configured to exert, against the magnet, a magnetic force in a direction pushing the damper pin away from a stick region of the damper pin located on a radially outward side of the rotor in the gap.

Abstract: A rotating machine includes a plurality of rotors. Each of the rotors includes a rotor bore protruding radially inward from a platform. A ring shaped cover plate is interfaced with each rotor bore via at least one snap. The at least one snap includes a first arm extending from the cover plate and having a convex facing contact surface, and multiple second arms extending axially from each rotor bore and having contact surfaces facing the convex surface.

Abstract: A hub assembly for a ducted-rotor aircraft includes a plurality of stator supports, a first integral fitting that attaches to first ends of the stator supports, and a second integral fitting that attaches to second ends of the stator supports. The first fitting includes a plurality of motor attachment portions configured to locate and support a motor of the ducted-rotor aircraft, a plurality of fairing mounts configured to support attachment of an aerodynamic fairing to the hub assembly, and a plurality of stator locators configured to locate respective ones of the plurality of stator supports for attachment to the hub assembly. The first fitting has a body portion with an annular wall that supports the motor attachment portions, the fairing mounts, and the stator locators. The first fitting may be fabricated as a monolith that includes the body portion, motor attachment portions, fairing mounts, and stator locators.