Abstract: A gas turbine engine rotor assembly including a rotor having a radially outer rim with an outer surface shaped to reduce circumferential rim stress concentration between each blade and the rim. Additionally, the shape of the outer surface directs air flow away from an interface between a blade and the rim to reduce aerodynamic performance losses between the rim and blades. In an exemplary embodiment, the outer surface of the rim has a concave shape between adjacent blades with apexes located at interfaces between the blades and the rim.

Abstract: A gas turbine engine includes a compressor rotor assembly which directs air at a sufficent pressure and temperature to a compressor rotor bore. The compressor assembly includes a compressor including a plurality of rotors joined with a plurality of couplings. Each of the rotors includes a radially outer rim, a radially inner hub, and a web extending therebetween. The web includes a flange including a plurality of slots which define a plurality of radial vanes that are airfoil-shaped. The vanes effectively deswirl bleed air and re-direct the airflow to a compressor bore.

Abstract: A compressor stator includes a casing having a circumferentially extending slot therein in which are mounted a plurality of vane sectors. Each sector includes a plurality of vanes extending between outer and inner bands, with each outer band being mounted in the slot. A plurality of tabs are ranged in interlocked pairs between the outer bands and casing for unloading adjacent sectors from each other.

Abstract: A compressor casing is configured to surround blade tips in a compressor stage. The casing includes stall grooves with adjoining lands defining respective local gaps with the blade tips. At least one of the lands is offset to locally increase a corresponding one of the gaps larger than the nominal gap for the casing to reduce tip rubbing thereat.

Abstract: A high flow interstage air extraction system for a gas turbine engine includes a 360 degree annular slot formed in a compressor stator casing by a first wall and a second wall. A plurality of baffle plates extend from the second wall to eliminate circumferential airflow. A plurality of stator vane platforms form at least a portion of the first wall. The annular slot results in large amounts of interstage compressor air being extracted from the compressor without adding length to the compressor and the engine.

Abstract: A bladed disk assembly includes a rotor disk having an axial dovetail slot, and a rotor blade having a dovetail disposed in the slot. A blade retainer is fixedly joined to the disk for retaining the dovetail in the slot. A spacer is disposed in the slot between the blade retainer and the dovetail and is selectively compressible for allowing the dovetail to slide in the slot toward the blade retainer for generating friction forces to dissipate energy upon impact of a predeterminedly sized foreign object.

Abstract: A system for changing pitch in aircraft propfans is disclosed, in which a motor, located at a stationary position, changes pitch of the blades by of a gear train. Three features of the system are that a feather brake can override the motor and drive the blades toward feather; a unison brake can prevent pitch from undergoing excursions toward fine pitch when propeller speed exceeds a limit; and the motor is non-rotating in the absence of pitch change. A pair of such systems can be used in a counterrotating propfan system, in order to independently change pitch of the propfans.

Abstract: The invention concerns aircraft propellers of the variable pitch type. Each propeller blade is carried by a trunnion which is rotatably mounted to a common ring. Each trunnion contains a reduction gear set which is driven by a radial shaft. The reduction gear sets multiply the torque applied by the shafts, thus allowing shafts of reduced diameter to be used. The reduced diameter is important when the shafts must pass through a turbine flowpath, and thus be concealed within turbine blades. The reduced diameter shafts can be easily concealed within turbine blades without redesigning and compromising blade aerodynamics.

Abstract: The invention relates to pitch-change systems in aircraft propellers. The invention includes redundant planetary gear systems, each of which can independently change pitch if the other fails. The invention also includes a system for reducing displacement of the planets, which can occur during a failure.