Abstract: A turbine containment system is provided. The turbine containment system includes a first containment member surrounding a portion of a turbine including a plurality of rotor disks and rotor blades; and a second containment member in communication with the first containment member, wherein the first containment member and the second containment member together contain each of the rotor disks and the rotor blades therein.

Abstract: A turbine containment system is provided. The turbine containment system includes a first containment member surrounding a portion of a turbine including a plurality of rotor disks and rotor blades; and a second containment member in communication with the first containment member, wherein the first containment member and the second containment member together contain each of the rotor disks and the rotor blades therein.

Abstract: A strut assembly for a gas turbine engine includes an outer structural case. The outer structural case includes a first mounting pad for mounting a first strut and a second mounting pad for mounting a second strut. The outer structural case further includes a case ligament extending between the first mounting pad and the second mounting pad in a substantially straight direction to reduce an amount of bending stress on the outer structural case.

Abstract: A strut assembly for a gas turbine engine includes an outer structural case. The outer structural case includes a first mounting pad for mounting a first strut and a second mounting pad for mounting a second strut. The outer structural case further includes a case ligament extending between the first mounting pad and the second mounting pad in a substantially straight direction to reduce an amount of bending stress on the outer structural case.

Abstract: Methods, systems, and apparatus for controlling a turbine clearance in an aircraft engine are provided. A method includes activating a turbine clearance control based on a flight phase of an aircraft using the aircraft engine, and adjusting the turbine clearance based on a preselected turbine clearance value.

Abstract: Methods, systems, and apparatus for controlling a turbine clearance in an aircraft engine are provided. A method includes activating a turbine clearance control based on a flight phase of an aircraft using the aircraft engine, and adjusting the turbine clearance based on a preselected turbine clearance value.

Abstract: An integrated engineering analysis system that determines at least one final output value in response to at least one initial input value. A first subprocesses provides at least one output value in response to the initial input value. A plurality of subprocesses provide a plurality of output values in response to the output values of the first subprocess. The plurality of output values and the output value of the first subprocess are inputted into at least one of the plurality of subprocesses. A command code executes each of the plurality of subprocesses in response to the output values and determines which one of the plurality of subprocesses to run and a final subprocess provides a final output. The final output is the result of a plurality of calculations executed by the plurality of subprocesses and the command code compares the final output with the initial input.

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 vane includes pressure and suction sides extending chordally between leading and trailing edges, and longitudinally between a root and a tip. The vane narrows in chord to a waist between the root and tip. The vane may also be bowed at its trailing edge in cooperation with the narrow waist.

Abstract: A new stator core cooling concept is provided that can be implemented with a conventional core structure having cooling ducts interspersed at intervals along the axial length of the stator core. To enhance cooling of the stator coils, the cooling ducts are radially closed off so that cooling air is forced to flow circumferentially, preferably through small holes, channels, or slots, to impinge directly on the armature bar surface. Accordingly, rather than flowing cooling gas through the cooling ducts and immediately discharging it into the annular gap between the rotor and stator, cooled gas flows radially down through the cooling duct and then must flow circumferentially whereupon it impinges directly on the armature bar surface itself.

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: An arrangement is provided for statically sealing between the inner shroud 12 of individual stator vanes of a gas turbine, and a seal housing 14 provided with a radially outwardly open channel 22 into which is received a plurality of outer seal segments 24, and which includes a radially outwardly open channel 33 throughout the length of each outer seal segment and with a plurality of inner seal segments received in the channel, with the outer seal segments being biased outwardly by the springs 36, and the inner seal segments 34 being biased outwardly into sealing contact with individual inner shrouds by the concentric inner compression springs 38 carried by the outer seal segments 24.