A plant according to an embodiment includes a gas turbine; a heat recovery unit that includes a primary heat recovery steam generator in which a primary heat exchanging unit generates primary steam by exchanging heat that is the thermal energy of the flue gas from the gas turbine, and a secondary heat recovery steam generator that is installed independently from the primary heat exchanging unit, and in which a secondary heat exchanging unit generates secondary steam by exchanging heat that is the thermal energy of the flue gas partly having exchanged heat in the primary heat exchanging unit included in the primary heat recovery steam generator; a primary steam turbine; a CO.sub.2 recovery unit; and a first reboiler heat supply line.

A plant according to an embodiment includes a gas turbine; a heat recovery unit that includes a primary heat recovery steam generator in which a primary heat exchanging unit generates primary steam by exchanging heat that is the thermal energy of the flue gas from the gas turbine, and a secondary heat recovery steam generator that is installed independently from the primary heat exchanging unit, and in which a secondary heat exchanging unit generates secondary steam by exchanging heat that is the thermal energy of the flue gas partly having exchanged heat in the primary heat exchanging unit included in the primary heat recovery steam generator; a primary steam turbine; a CO.sub.2 recovery unit; and a first reboiler heat supply line.

An aircraft engine having a compressor, the compressor having at least one flutter-resistant blade, the blade having a leading edge (LE), a trailing edge (TE), a midchord (MC), a minimum radial height r.sub.hub, a maximum radial height r.sub.tip, and a radial extent between r.sub.hub and r.sub.tip, wherein, at every point along the radial extent of the blade, the blade has a modeshape value V.sub.1 for a blade first vibratory mode defined as

[00001]$V_1=\frac{.Math.D_{1\mathrm{LE}}.Math.}{.Math.D_{1\mathrm{MC}}.Math.},$

and wherein, when the engine is operating between its maximum speed and 70% of that maximum speed, at least 80% of the radial extent of the blade has a modeshape value V.sub.1 from 0 to 1.5.

An aircraft engine having a compressor, the compressor having at least one flutter-resistant blade, the blade having a leading edge (LE), a trailing edge (TE), a midchord (MC), a minimum radial height r.sub.hub, a maximum radial height r.sub.tip, and a radial extent between r.sub.hub and r.sub.tip, wherein, at every point along the radial extent of the blade, the blade has a modeshape value V.sub.1 for a blade first vibratory mode defined as

[00001]$V_1=\frac{.Math.D_{1\mathrm{LE}}.Math.}{.Math.D_{1\mathrm{MC}}.Math.},$

and wherein, when the engine is operating between its maximum speed and 70% of that maximum speed, at least 80% of the radial extent of the blade has a modeshape value V.sub.1 from 0 to 1.5.

A turbine engine comprising a compressor section and a turbine section in serial flow arrangement defining a working air flow path with a heat exchanger in fluid communication the working air flow path, and a nuclear fuel in thermal communication with the heat exchanger and a release valve in fluid communication with the working air flow path.

A turbine engine comprising a compressor section and a turbine section in serial flow arrangement defining a working air flow path with a heat exchanger in fluid communication the working air flow path, and a nuclear fuel in thermal communication with the heat exchanger and a release valve in fluid communication with the working air flow path.

A method for operating a hybrid-electric propulsion system of an aircraft is provided. The hybrid-electric propulsion system includes a gas turbine engine having a high pressure system, a low pressure system, and an electric machine coupled to one of the high pressure system or low pressure system. The method includes receiving data indicative of an actual or anticipated in-flight shutdown of the gas turbine engine; and adding power to the gas turbine engine through the electric machine in response to receiving data indicative of the actual or anticipated in-flight shutdown of the gas turbine engine.

A method for operating a hybrid-electric propulsion system of an aircraft is provided. The hybrid-electric propulsion system includes a gas turbine engine having a high pressure system, a low pressure system, and an electric machine coupled to one of the high pressure system or low pressure system. The method includes receiving data indicative of an actual or anticipated in-flight shutdown of the gas turbine engine; and adding power to the gas turbine engine through the electric machine in response to receiving data indicative of the actual or anticipated in-flight shutdown of the gas turbine engine.

A thrust reverser system is provided that includes a sleeve, a fixed cascade structure and a blocker door. The fixed cascade structure is within a cavity of the sleeve when the sleeve is in a sleeve stowed position. The blocker door is within the cavity of the sleeve when the sleeve is in the sleeve stowed position and the blocker door is in a blocker door stowed position. The blocker door projects in a radial inward direction away from the sleeve towards the centerline when the sleeve is in a sleeve deployed position and when the blocker door is in a blocker door deployed position. The blocker door includes a pivot attachment fixed at an end of the blocker door. The pivot attachment is configured to move in a forward direction from a first location to a second location.

A transition duct support apparatus and a method to support an exit frame in a transition duct in a gas turbine engine are provided. A stiffener (24) may be arranged to provide support to an outer edge (27) of an exit frame (12) in a transition duct (14). Stiffener (24) may be configured to circumferentially extend between mutually opposed corners (30) of the exit frame of the transition duct. A brace (26) may be connected to a centrally-disposed section (20) and may extend to support respective end portions (32) of the stiffener. The support apparatus is effective to provide a respective tuned level of stiffness support with respect to one or more axes of the exit frame in the transition duct. The apparatus and method may be effective for distributing mechanical stresses on the exit frame of the transition duct and/or neighboring regions in the transition duct.