Abstract: The gas turbine engine can have an engine core; a core output shaft drivable by the engine core; a power output shaft; an auxiliary power shaft; and a reduction gearbox having gears, the gears drivingly connecting the core output shaft to the auxiliary power shaft. The gears can include an epicyclic gearing drivingly connecting the core output shaft and the auxiliary power shaft to the power output shaft. The gas turbine engine can further have a second auxiliary power shaft interconnected to the auxiliary power shaft, the power output shaft, and the core output shaft by the gears.

Abstract: A method of manufacturing a shroud segment for gas turbine engine includes providing an insert having a plurality of pins that extend into a platform cavity portion of a mold cavity. A powder injection molding feedstock is injected. When the green part is disengaged from the mold, each elongated feature is slid out of the green part to define a respective elongated cooling passage in the platform. The method may include, after debinding and sintering, projecting a coating material while defining an obstruction between source of coating material and the open end of each elongated feature with a shoulder of the element to prevent the coating material from reaching the open end, followed by machining to remove at least a part of the shoulder.

Abstract: The gas turbine engine can have an engine core; a core output shaft drivable by the engine core; a power output shaft; an auxiliary power shaft; and a reduction gearbox having gears, the gears drivingly connecting the core output shaft to the auxiliary power shaft. The gears can include an epicyclic gearing drivingly connecting the core output shaft and the auxiliary power shaft to the power output shaft. The gas turbine engine can further have a second auxiliary power shaft interconnected to the auxiliary power shaft, the power output shaft, and the core output shaft by the gears.

Abstract: The gas turbine engine can have an engine core; a core output shaft drivable by the engine core; a power output shaft; an auxiliary power shaft; and a reduction gearbox having gears, the gears drivingly connecting the core output shaft to the auxiliary power shaft. The gears can include an epicyclic gearing drivingly connecting the core output shaft and the auxiliary power shaft to the power output shaft. The gas turbine engine can further have a second auxiliary power shaft interconnected to the auxiliary power shaft, the power output shaft, and the core output shaft by the gears.

Abstract: A method of manufacturing a shroud segment for gas turbine engine includes providing an insert having a plurality of pins that extend into a platform cavity portion of a mold cavity. A powder injection molding feedstock is injected. When the green part is disengaged from the mold, each elongated feature is slid out of the green part to define a respective elongated cooling passage in the platform. The method may include, after debinding and sintering, projecting a coating material while defining an obstruction between source of coating material and the open end of each elongated feature with a shoulder of the element to prevent the coating material from reaching the open end, followed by machining to remove at least a part of the shoulder.

Abstract: A method of manufacturing a gas turbine engine element, for example a shroud segment. An insert has at least one elongated feature received in a mold cavity. A powder injection molding feedstock is injected. When the green part is disengaged from the mold, each elongated feature is slid out of the green part to define a respective elongated passage. The cross-sectional dimension of the elongated feature may be 0.020 inches or less, and/or a ratio between the length and cross-sectional dimension of the elongated feature may be at least 25. The method may include, after debinding and sintering, projecting a coating material while defining an obstruction between source of coating material and the open end of each elongated feature with a shoulder of the element to prevent the coating material from reaching the open end, followed by machining to remove at least a part of the shoulder.

Abstract: A rotor blade assembly tool for coupling a plurality of circumferentially spaced rotor blades to a rotor disc of a turbine rotor assembly, includes a base ring with an array of circumferentially spaced resilient fingers axially extending from the base ring. The resilient fingers are configured each to radially abut a blade seal, damper or other engine component against radially inner facing surfaces of platforms of the respective blades when the blades are being seated onto the disc during an assembly procedure.

Abstract: A method of manufacturing a gas turbine engine element, for example a shroud segment. An insert has at least one elongated feature received in a mold cavity. A powder injection molding feedstock is injected. When the green part is disengaged from the mold, each elongated feature is slid out of the green part to define a respective elongated passage. The cross-sectional dimension of the elongated feature may be 0.020 inches or less, and/or a ratio between the length and cross-sectional dimension of the elongated feature may be at least 25. The method may include, after debinding and sintering, projecting a coating material while defining an obstruction between source of coating material and the open end of each elongated feature with a shoulder of the element to prevent the coating material from reaching the open end, followed by machining to remove at least a part of the shoulder.

Abstract: A rotor blade assembly tool for coupling a plurality of circumferentially spaced rotor blades to a rotor disc of a turbine rotor assembly, includes a base ring with an array of circumferentially spaced resilient fingers axially extending from the base ring.

Abstract: A rotor blade assembly tool for coupling a plurality of circumferentially spaced rotor blades to a rotor disc of a turbine rotor assembly, includes a base ring with an array of circumferentially spaced resilient fingers axially extending from the base ring. The resilient fingers are configured each to radially abut a blade seal, damper or other engine component against radially inner facing surfaces of platforms of the respective blades when the blades are being seated onto the disc during an assembly procedure.

Abstract: A rotor containment structure for gas turbine engine comprises an inner containment layer having a single integral body defining an annular structure about the rotor, and having a support on the inner surface of the inner containment layer for at least one shroud segment. An outer containment layer provides containment strength to contain blade fragments and defining an annular structure about the inner containment layer. Air passage through the outer containment layer for air to pass from an exterior of the outer containment layer to an interior of the outer containment layer; and the inner containment layer being connected at a first end to the outer containment layer with a gap defined between the inner surface of the outer containment layer and the outer surface of the inner containment layer, the gap being in fluid communication with the air passage such that air flows through the gap, beyond a free second end of the inner containment layer.

Abstract: A blade outer air seal (BOAS) of a gas turbine engine has a segmented support ring to support a segmented turbine shroud. The support ring has a cooling air distribution system which includes a plurality of inlet cavities extending axially and inwardly to communicate with an inner cooling air passage within the respective support segments. The inlet cavities each are formed with two recesses defined in respective adjacent two support segments.

Abstract: A balancing assembly for a rotor assembly of a gas turbine engine includes a flange engaging portion defining a flange receiving opening for mating with the flange in a slidable manner. A hole extends through at least part of the flange engaging portion, the hole corresponding in size to a recess and adapted to receive a rivet or other suitable fastener to fix the balancing weight with respect to the disc.

Abstract: A retaining ring arrangement is provided for axially holding a component on a rotating component of a gas turbine engine. The retaining ring arrangement comprises a split retaining ring mounted in a circumferential groove defined in a radially outer surface of the rotating component. The inner diameter of the retaining ring is biased inwardly in radial contact with a radially outer facing seat provided on one of the two components to be assembled. An anti-rotation feature is provided at the inner diameter of the retaining ring for restraining the ring against rotation. A sleeve surrounds the retaining ring to limit radial expansion thereof when subject to centrifugal forces during engine operation.

Abstract: A rotor blade assembly tool for coupling a plurality of circumferentially spaced rotor blades to a rotor disc of a turbine rotor assembly, includes a base ring with an array of circumferentially spaced resilient fingers axially extending from the base ring.

Abstract: An apparatus and method for balancing a gas turbine engine rotor includes a plurality of balancing weights adapted to be selectively attached to at least one of inlets or outlets of a cooling passage of the rotor. The weights include cooling access which permits coolant to communicate with the cooling passage.

Abstract: A retaining ring arrangement is provided for axially holding a component on a rotating component of a gas turbine engine. The retaining ring arrangement comprises a split retaining ring mounted in a circumferential groove defined in a radially outer surface of the rotating component. The inner diameter of the retaining ring is biased inwardly in radial contact with a radially outer facing seat provided on one of the two components to be assembled. An anti-rotation feature is provided at the inner diameter of the retaining ring for restraining the ring against rotation. A sleeve surrounds the retaining ring to limit radial expansion thereof when subject to centrifugal forces during engine operation.

Abstract: A rotor containment structure for gas turbine engine comprises an inner containment layer having a single integral body defining an annular structure about the rotor, and having a support on the inner surface of the inner containment layer for at least one shroud segment. An outer containment layer provides containment strength to contain blade fragments and defining an annular structure about the inner containment layer. Air passage through the outer containment layer for air to pass from an exterior of the outer containment layer to an interior of the outer containment layer; and the inner containment layer being connected at a first end to the outer containment layer with a gap defined between the inner surface of the outer containment layer and the outer surface of the inner containment layer, the gap being in fluid communication with the air passage such that air flows through the gap, beyond a free second end of the inner containment layer.

Abstract: An apparatus and method for balancing a gas turbine engine rotor includes a plurality of balancing weights adapted to be selectively attached to at least one of inlets or outlets of a cooling passage of the rotor. The weights include cooling access which permits coolant to communicate with the cooling passage.

Abstract: A blade outer air seal (BOAS) of a gas turbine engine has a segmented support ring to support a segmented turbine shroud. The support ring has a cooling air distribution system which includes a plurality of inlet cavities extending axially and inwardly to communicate with an inner cooling air passage within the respective support segments. The inlet cavities each are formed with two recesses defined in respective adjacent two support segments.