Abstract: A gas turbine engine casing apparatus includes annular first and second annular cases connected by a spigot joint. The spigot joint includes an annular projection of the first annular case fitted into an annular recess of the second annular case. A plurality of circumferentially spaced apart scallops are formed on one of surfaces of the annular projection or of the annular recess, and are located in selective circumferential locations adjacent respective enhanced stiff areas of the first and second annular cases.

Abstract: A compressor rotor for a gas turbine engine includes a hub disposed about an axis of rotation and an outer surface forming a radially inner gaspath boundary, the outer surface defining a nominal hub diameter. A circumferential array of blades extends radially outwardly from the hub. A first inter-blade passage is defined between a first set of adjacent blades and has a first throat area. A second inter-blade passage is defined between a second set of adjacent blades and has a second throat area that is smaller than the first throat area. At least one scoop is disposed in the second inter-blade passage, the scoop defining a cavity extending radially into the outer surface of the hub relative to the nominal hub diameter.

Abstract: A fancase surrounds a set of fan blades mounted for rotation about a central axis of a turbine engine. The fancase has an annular casing structure having an inner wall defining an outer flow boundary surface of a gaspath. An annular recess is defined in the inner wall. A metallic bumper is disposed in the annular recess and bonded to the annular casing structure for limiting fan orbiting during fan unbalanced conditions.

Abstract: A rotor for a gas turbine engine. The rotor includes blades circumferentially distributed around a hub. The blades have airfoils with a span defined between a root and tip, a chord defined between a leading edge and a trailing edge, and a thickness defined between a pressure side surface and suction side surface. The blades include first blades and second blades. The airfoil of the first blades has a first thickness distribution defining a first natural vibration frequency of the airfoils of the first blades. The airfoil of the second blades has a second thickness distribution defining a second natural vibration frequency different than the first natural vibration frequency. The first thickness distribution is different than the second thickness distribution along a radially-inner half of the span, and the first thickness distribution matches the second thickness distribution along a radially-outer half of the span.

Abstract: A fan for a gas turbine engine comprises blades distributed around a hub. The blades include first and second blades, having geometric parameters and/or material properties that differ from each other to frequency mistune the fan. The blades are distributed about the hub with at least one second blade between adjacent first blades. The leading edge of the airfoil of the second blades is disposed axially aft of the corresponding leading edge of the first blades in at least a portion of the blade span.

Abstract: A vane assembly adapted to be disposed through an annular gas flow path defined between a fan case and an engine core of a gas turbine engine. The assembly comprises a vane having a vane body configured for extending between the engine core and the fan case, a vane head disposed at one end of the vane body, the vane head being adapted to be disposed outside the annular gas flow path, the vane head having an abutting surface configured for contacting an outer surface of the fan case, and a recess extending within the vane head and opening to the abutting surface. The recess is configured for circumferentially surrounding a longitudinal axis of the vane and forming a closed figure. The assembly also comprises a sealing member disposed within the recess.

Abstract: A gas turbine engine assembly comprises a casing defining a gas path, the casing including a shroud having an annular body having a surface defining a portion of gas path, the shroud having slots configured for receiving inserted vanes. The slots are delimited substantially about their perimeter by respective flanges, the flanges radially offset from the shroud gas path surface so as to be disposed outside of said gas path, the flanges defined by opposed flange surfaces. Vanes received in the slots. Grommets engage the vanes at the slots. Inserts extend between the shroud and the grommets, the inserts having slots configured for engaging both of the opposed flanges, the inserts extending in a radial direction from at least the respective flange to an adjacent said shroud gas path surface to substantially matchingly mate with an inner surface the adjacent shroud gas path surface.

Abstract: There is disclosed a shroud for a compressor stator, including: shroud segments extending circumferentially around an axis along portions of a circumference of the shroud. At least one of the shroud segments extending from a first lateral edge to a second lateral edge. The shroud segment has an inner face oriented toward the axis and an opposed outer face oriented away from the axis. At least one opening extends from the inner face to the outer face for receiving a vane of the compressor. A tab protrudes circumferentially from the second lateral edge and away from the first lateral edge. A slot extends circumferentially from the first lateral edge toward the second lateral edge. The tab is matingly received within a slot of an adjacent one of the shroud segments.

Abstract: A turbofan engine is disclosed which includes a nacelle assembly, having an interior surface for directing airflow, and a flow disruptor positioned on the interior surface upstream of the fan, the flow disruptor extending towards the axis a height greater than the anticipated boundary layer height of the airflow. A turbofan engine which includes an array of circumferentially disposed flow disruptors extending from a fan case inner surface is also disclosed. A method of mitigating fan flutter in a gas turbine engine by generating a circumferential asymmetrically in the airflow, upstream of the fan, is also described.

Abstract: A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.

Abstract: A variable guide vane (VGV) apparatus has a variable guide vane (VGV) rotatable about a vane rotation axis. An actuating arm is mounted to the VGV. The actuating arm has a fork defining a slot for receiving a drive pin. The slot has a curved contour configured to act as a vane angle schedule adjustment.

Abstract: A variable guide vane (VGV) apparatus has a variable guide vane (VGV) rotatable about a vane rotation axis. An actuating arm is mounted to the VGV. The actuating arm has a fork defining a slot for receiving a drive pin. The slot is profiled to accommodate angular misalignment between the pin and the fork throughout a range of motion of the associated pin along the slot.

Abstract: A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.

Abstract: A gas turbine engine is described having a rotor with a plurality of rotor blades each extending radially between a blade root to a blade tip and axially between a leading edge and a trailing edge. An annular casing body housing the rotor blades includes an abradable segment of the inner surface facing the blade tips and having an abradable member. The abradable segment having one or more annular grooves, extending into the casing body from an inner surface thereof, and including an edge groove axially aligned with and facing the leading edge or the trailing edge of the rotor blades. The edge groove extends axially between a first position on the inner surface upstream of the leading edge or the trailing edge, and a second position on the inner surface downstream of the leading edge or the trailing edge.

Abstract: A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.

Abstract: A rotor for a gas turbine engine comprises a rotor having a hub and blades around the hub, and extending from the hub to tips. The tips include first and second tip portions between their respective tip leading edge and tip trailing edge. Tips are spaced from a rotational axis of the rotor by spans. A mean span of a first tip portion of a first blade is greater than a mean span of a corresponding first tip portion of a second blade. A mean span of a second tip portion the first blade is less than a mean span of a corresponding second tip portion of the second blade.

Abstract: A bleed of valve comprises a ring axially translatable between a retracted position in which the ring is configured to close an annular bleed off opening defined in a converging portion of a radially outer annular wall of a gas turbine engine gaspath and a deployed position in which the ring protrudes into the gaspath to mechanically scoop out incoming air and water/hail particles.

Abstract: A seal spacer for a gas turbine engine to provide sealing between a casing wall and a stator vane of the engine. The seal spacer comprises a circumferential body having a slot defined therethrough and configured to receive the stator vane therein. The circumferential body is configured to be disposed between and space apart the engine casing wall from a head of the stator vane when the vane is inserted in the slot and the vane is inserted in the opening in the casing wall. The circumferential body includes a support positioned to be engage and provide a gap between the casing wall and the head when the seal spacer is installed therebetween, and an elastomeric seal portion positioned to engage and compressingly provide sealing the casing wall and the head when the seal spacer is installed therebetween to seal the gap.

Abstract: A gas turbine engine casing apparatus includes annular first and second annular cases connected by a spigot joint. The spigot joint includes an annular projection of the first annular case fitted into an annular recess of the second annular case. A plurality of circumferentially spaced apart scallops are formed on one of surfaces of the annular projection or of the annular recess, and are located in selective circumferential locations adjacent respective enhanced stiff areas of the first and second annular cases.

Abstract: A controlled gap seal device is adapted to surround a shaft with an annular seal element. A ring is positioned on an outer surface of the seal element, the ring adapted to modify a diametrical dimension of the seal element by thermally expanding/contracting as a function of temperature variations. A housing assembly has an interior enclosing the seal element and the ring, with the seal configured to be generally stationary in the interior. The housing assembly has an air inlet and air outlet in fluid communication with a surrounding environment for directing a flow of gas from the surrounding environment onto the ring to controllably cool and shrink the ring. A method for modifying a diameter of a controlled gap seal relative to the shaft is also provided.