Abstract: A variable area exhaust nozzle defining an actual area ratio for a gas turbine engine that includes a shroud assembly having a fixed shroud and an adjustable cowl, an adjustable plug, a nozzle adjustment assembly, and a controller configured to receive data indicative of an operating speed of the gas turbine engine; determine a target area ratio based at least in part on received data indicative of the operating speed of the gas turbine engine; and operate the nozzle adjustment assembly to selectively adjust at least one of the adjustable cowl or the adjustable plug such that the actual area ratio is substantially equivalent to the target area ratio.

Abstract: A variable area exhaust nozzle defining an actual area ratio for a gas turbine engine that includes a shroud assembly having a fixed shroud and an adjustable cowl, an adjustable plug, a nozzle adjustment assembly, and a controller configured to receive data indicative of an operating speed of the gas turbine engine; determine a target area ratio based at least in part on received data indicative of the operating speed of the gas turbine engine; and operate the nozzle adjustment assembly to selectively adjust at least one of the adjustable cowl or the adjustable plug such that the actual area ratio is substantially equivalent to the target area ratio.

Abstract: A gas turbine engine includes a fan section having a single-stage fan, a core turbine engine, and a nacelle at least partially surrounding the fan of the fan section and the core turbine engine. The gas turbine engine also includes an outlet guide vane extending between a casing of the core turbine engine and the nacelle. The gas turbine engine is configured to define an acoustic ratio greater than or equal to 2.3, with the acoustic ratio being a ratio of an axial spacing between the trailing edge of the fan blade and the leading edge of the outlet guide vane at a radial location seventy-five percent (75%) along the span of the fan blade to the axial width of the fan blade also at the radial location seventy-five percent (75%) along the span of the fan blade.

Abstract: A gas turbine engine includes a fan section and a core turbine engine operable with the fan section. The gas turbine engine also includes a sump pressurization assembly including a valve, a first duct in airflow communication with the valve and a first pressurized air source, and a second duct in airflow communication with the valve and a second pressurized air source. Further, the sump pressurization assembly includes a supply duct selectively in airflow communication with the first duct and the second duct through the valve. The supply duct is located at least partially inward of a core air flowpath of the core turbine engine for pressurizing a sump of the gas turbine engine.

Abstract: A supersonic turbofan engine includes a fan section having a single-stage fan defining a fan pressure ratio greater than 1.9. The supersonic turbofan engine also includes a core turbine engine defining a core air flowpath. A nacelle at least partially surrounds the fan of the fan section and the core turbine engine. The supersonic turbofan engine defines a bypass ratio, the bypass ratio being greater than or equal to three.

Abstract: A gas turbine engine includes a fan section having a single-stage fan, a core turbine engine, and a nacelle at least partially surrounding the fan of the fan section and the core turbine engine. The gas turbine engine also includes an outlet guide vane extending between a casing of the core turbine engine and the nacelle. The gas turbine engine is configured to define an acoustic ratio greater than or equal to 2.3, with the acoustic ratio being a ratio of an axial spacing between the trailing edge of the fan blade and the leading edge of the outlet guide vane at a radial location seventy-five percent (75%) along the span of the fan blade to the axial width of the fan blade also at the radial location seventy-five percent (75%) along the span of the fan blade.

Abstract: A gas turbine engine includes a fan section and a core turbine engine operable with the fan section. The gas turbine engine also includes a sump pressurization assembly including a valve, a first duct in airflow communication with the valve and a first pressurized air source, and a second duct in airflow communication with the valve and a second pressurized air source. Further, the sump pressurization assembly includes a supply duct selectively in airflow communication with the first duct and the second duct through the valve. The supply duct is located at least partially inward of a core air flowpath of the core turbine engine for pressurizing a sump of the gas turbine engine.

Abstract: A gas turbine operational strategy is provided for the optimization of ISO-Day rated performance. The invention is realized by setting the nominal operating line equal to or slightly below the operating limit line, so that ISO full load performance is maximized. Thus, the first stage turbine nozzle throat area is sized such that, without inlet bleed heat addition, and at Base-Load pressure ratio and firing temperature, the compressor operates within its operational limits on a ISO day. As ambient conditions change, levels of inlet bleed heat and firing temperature are controlled and the inlet guide vane angle is set whereby the nominal operating line is equal to or slightly below the operating limit line.