Abstract: A method for assembling a gas turbine engine includes providing a first fan assembly including a plurality of rotor blades that are configured to rotate in a first rotational direction at a first rotational speed, rotatably coupling a second fan assembly axially aft of the first fan assembly, wherein the second fan assembly includes a plurality of rotor blades that are configured to rotate in a second rotational direction, and coupling a gearbox to the second fan assembly that is configured to rotate the second fan assembly at a second rotational speed that is different than the first rotational speed.

Abstract: A method for assembling a gas turbine engine that includes providing a low-pressure turbine inner rotor that includes a first plurality of turbine blade rows configured to rotate in a first direction, and rotatably coupling a low-pressure turbine outer rotor to the inner rotor, wherein the outer rotor includes a second plurality of turbine blade rows that are configured to rotate in a second direction that is opposite the first rotational direction of the inner rotor and such that at least one of the second plurality of turbine blade rows is coupled axially forward of the first plurality of turbine blade rows.

Abstract: A method for assembling a gas turbine engine includes providing a high pressure turbine, providing a low-pressure turbine inner rotor that includes a first plurality of turbine blade rows configured to rotate in a first direction, providing a low-pressure turbine outer rotor that includes a second plurality of turbine blade rows configured to rotate in a second direction that is opposite the first direction, and coupling a plurality of bearings between the outer rotor and a turbine mid-frame between the high pressure turbine and at least one of the inner and outer rotors such that the plurality of bearings support a forward end of the outer rotor.

Abstract: A method for assembling a gas turbine engine that includes providing a low-pressure turbine inner rotor that includes a first plurality of turbine blade rows configured to rotate in a first direction, providing a low-pressure turbine outer rotor that includes a second plurality of turbine blade rows configured to rotate in a second direction that is opposite the first direction, coupling a turbine mid-frame assembly including a plurality of spokes within the engine such that the spokes are spaced axially forward of the inner rotor, coupling a bearing between the turbine mid-frame assembly and the inner rotor such that the inner rotor is rotatably coupled to the turbine mid-frame, and adjusting the plurality of spokes to align the bearing in a radial direction.

Abstract: A method for assembling a gas turbine engine includes providing a low-pressure turbine inner rotor that includes at least a first pair of adjacent rows of turbine blades that are configured to rotate in a first direction providing a low-pressure turbine outer rotor that includes a forward end, an aft end, and at least one row of turbine blades that is rotatably coupled between the first pair of inner rotor turbine blades, wherein the at least one row of outer rotor turbine blades is configured to rotate in a second direction that is opposite the first direction, and coupling a support assembly between the first pair of inner rotor turbine blades such that the support assembly supports the outer rotor forward end.

Abstract: A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a first drive shaft; a booster compressor positioned downstream of and in at least partial flow communication with the fan section including a plurality of stages, each stage including a stationary compressor blade row and a rotating compressor blade row connected to a drive shaft and interdigitated with the stationary compressor blade row; a core system positioned downstream of the compressor, where the core system further includes an intermediate compressor positioned downstream of and in flow communication with the booster compressor, the intermediate compressor being connected to a second drive shaft, and a combustion system for producing pulses of gas having increased pressure and temperature from a fluid flow provided to an inlet thereof so as to produce a working fluid at an outlet; and, a low pressure turbine

Abstract: A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a drive shaft; a booster compressor positioned downstream of the fan section including a plurality of stages, where each stage includes a stationary compressor blade row and a rotating compressor blade row connected to the drive shaft and interdigitated with the stationary compressor blade row; and, a combustion system for producing pulses of gas having increased pressure and temperature of a fluid flow provided to an inlet thereof so as to produce a working fluid at an outlet.

Abstract: A gas turbine engine includes a compressor powered by a turbine. The turbine includes a nozzle having vanes extending between outer and inner bands. Each vane includes an internal cooling plenum and a bypass tube extending through the bands. First and second manifolds surround the outer band and are disposed in flow communication with the plenums and bypass tubes, respectively. A bleed circuit joins the compressor to the manifolds for providing pressurized air thereto. A control valve modulates airflow to the first manifold and in turn through the cooling plenums of the vanes.

Abstract: A method enables a gas turbine engine including a first rotor assembly and a second rotor assembly coupled in axial flow communication downstream from the first rotor assembly to be assembled. The method comprises coupling an upstream end of an extension duct to an outlet of the first rotor assembly, wherein the extension duct includes a plurality of panels coupled circumferentially, and coupling a downstream end of the extension duct to an inlet of the second rotor assembly using at least one fish mouth seal.

Abstract: A method enables a gas turbine engine including a first rotor assembly and a second rotor assembly coupled in axial flow communication downstream from the first rotor assembly to be assembled. The method comprises coupling an upstream end of an extension duct to an outlet of the first rotor assembly, wherein the extension duct includes a plurality of panels coupled circumferentially, and coupling a downstream end of the extension duct to an inlet of the second rotor assembly using at least one fish mouth seal.

Abstract: A gas turbine engine turbine assembly includes a high pressure spool having a high pressure turbine drivingly connected to a high pressure compressor by a high pressure shaft which is rotatable about an engine centerline. A low pressure turbine has counter rotatable low pressure inner and outer shaft turbines drivingly connected to coaxial low pressure inner and outer shafts respectively which are at least in part rotatably disposed co-axial with and radially inwardly of the high pressure spool. The low pressure inner shaft turbine is drivingly connected to a forward fan blade row by the low pressure inner shaft. The low pressure outer shaft turbine is drivingly connected to a aft fan blade row by the low pressure outer shaft. A single direction of rotation booster is drivenly connected to the low pressure outer shaft and axially located aft and downstream of the aft fan blade row.

Abstract: A gas turbine engine turbine assembly includes a high pressure spool having a high pressure turbine drivingly connected to a high pressure compressor by a high pressure shaft which is rotatable about an engine centerline. A low pressure turbine has counter rotatable low pressure inner and outer shaft turbines with a row of low pressure variable vanes disposed therebetween and drivingly connected to coaxial low pressure inner and outer shafts respectively which are located radially inwardly of the high pressure spool. The low pressure inner and outer shaft turbines are drivingly connected to a forward and aft fan blade rows by the low pressure inner and outer shafts. A single direction of rotation booster is drivenly connected to the low pressure outer shaft and axially located aft and downstream of the aft fan blade row.

Abstract: An aircraft gas turbine engine includes a low pressure turbine having a low pressure turbine flowpath and low pressure inner and outer shaft turbines with counter rotatable low pressure inner and outer shaft turbine rotors, respectively. The low pressure inner and outer shaft turbine rotors include low pressure first and second turbine blade rows disposed across the turbine flowpath which are drivingly connected to first and second fan blade rows by low pressure inner and outer shafts, respectively. At least one row of low pressure variable vanes is operably disposed across the low pressure turbine flowpath between the low pressure inner and outer shaft turbines. The low pressure first turbine blade rows of the low pressure inner shaft turbine may be in tandem with or interdigitated with the second turbine blade rows of the low pressure outer shaft turbines.

Abstract: A counterrotatable booster compressor assembly for a gas turbine engine having a counterrotatable fan section with a first fan blade row connected to a first drive shaft and a second fan blade row axially spaced from the first fan blade row and connected to a second drive shaft, the counterrotatable booster compressor assembly including a first compressor blade row connected to the first drive shaft and a second compressor blade row interdigitated with the first compressor blade row and connected to the second drive shaft. A portion of each fan blade of the second fan blade row extends through a flowpath of the counterrotatable booster compressor so as to function as a compressor blade in the second compressor blade row.

Abstract: A low pressure turbine section for an aircraft gas turbine engine includes a low pressure turbine flowpath and counter rotatable low pressure inner and outer shaft rotors having inner and outer shafts, respectively. The low pressure inner and outer shaft rotors include low pressure first and second turbine blade rows disposed across the turbine flowpath and drivingly connected to first and second fan blade rows, respectively. The first low pressure turbine blade rows are disposed downstream of the second low pressure turbine blade rows along said low pressure turbine flowpath. Rows of non-rotatable low pressure vanes are disposed across the low pressure turbine flowpath between the first low pressure turbine blade rows and between the second low pressure turbine blade rows. An annular vaneless gap may be located between an aftmost row of the second low pressure turbine blade rows and a forwardmost row of the first low pressure turbine blade rows.

Abstract: A low pressure turbine section for an aircraft gas turbine engine includes a low pressure turbine flowpath and counter rotatable low pressure inner and outer shaft rotors having inner and outer shafts, respectively. The low pressure inner and outer shaft rotors include low pressure first and second turbine blade rows disposed across the turbine flowpath and drivingly connected to first and second fan blade rows, respectively. The first low pressure turbine blade rows are disposed downstream of the second low pressure turbine blade rows along said low pressure turbine flowpath. Rows of non-rotatable low pressure vanes are disposed across the low pressure turbine flowpath between the first low pressure turbine blade rows and between the second low pressure turbine blade rows. An annular vaneless gap may be located between an aftmost row of the second low pressure turbine blade rows and a forwardmost row of the first low pressure turbine blade rows.

Abstract: An aircraft gas turbine engine includes a low pressure turbine having a low pressure turbine flowpath and counter rotatable low pressure inner and outer shaft rotors having inner and outer shafts, respectively. The low pressure inner and outer shaft rotors include low pressure first and second turbine blade rows disposed across the turbine flowpath and drivingly connected to first and second fan blade rows by low pressure inner and outer shafts, respectively. At least one of the low pressure first and second turbine blade rows is interdigitated with an adjacent pair of one of the turbine blade rows. At least one row of non-rotatable low pressure vanes is disposed across the low pressure turbine flowpath between a non interdigitated adjacent pair of one of the turbine blade rows.

Abstract: An aircraft gas turbine engine includes a low pressure turbine having a low pressure turbine flowpath and counter rotatable low pressure inner and outer shaft rotors having inner and outer shafts, respectively. The low pressure inner and outer shaft rotors include low pressure first and second turbine blade rows disposed across the turbine flowpath and drivingly connected to first and second fan blade rows by low pressure inner and outer shafts, respectively. At least one of the low pressure first and second turbine blade rows is interdigitated with an adjacent pair of one of the turbine blade rows. At least one row of non-rotatable low pressure vanes is disposed across the low pressure turbine flowpath between a non interdigitated adjacent pair of one of the turbine blade rows.

Abstract: A method of extending a useful serviceable life of a gas turbine engine by increasing a limiting exhaust gas temperature margin degraded by engine use without replacing engine hardware. The method includes adjusting at least one engine parameter selected from a first group of engine parameters including a nozzle area and a rotor speed.

Abstract: A counterrotatable booster compressor assembly for a gas turbine engine having a counterrotatable fan section with a first fan blade row connected to a first drive shaft and a second fan blade row axially spaced from the first fan blade row and connected to a second drive shaft, the counterrotatable booster compressor assembly including a first compressor blade row connected to the first drive shaft and a second compressor blade row interdigitated with the first compressor blade row and connected to the second drive shaft. A portion of each fan blade of the second fan blade row extends through a flowpath of the counterrotatable booster compressor so as to function as a compressor blade in the second compressor blade row.