A gas turbine engine having a blade assembly with a platform, an airfoil, and a shank. The airfoil has a plurality of cooling conduits, and the shank has a plurality of inlet passages to provide cooling fluid to the cooling conduits in the airfoil. The cooling fluid is vented through a plurality of cooling holes along the airfoil. The blade assembly has specific geometries that improve durability.

A gas turbine engine includes a turbomachine having an engine core including a high-pressure compressor, a combustion section, a high-pressure turbine, and a high-pressure shaft coupled to the high-pressure compressor and the high-pressure turbine. The engine core has a length (L.sub.CORE), and the high-pressure compressor has an exit stage diameter (D.sub.CORE). The high-pressure compressor defines a high-pressure compressor exit area (A.sub.HPCExit) in square inches. The gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn.sub.Total) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: Fn.sub.TotalEGT/(A.sub.HPCExit.sup.21000). The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of L.sub.CORE/D.sub.CORE is from 2.1 to 4.3.

A gas turbine engine includes a turbomachine having an engine core including a high-pressure compressor, a combustion section, a high-pressure turbine, and a high-pressure shaft coupled to the high-pressure compressor and the high-pressure turbine. The engine core has a length (L.sub.CORE), and the high-pressure compressor has an exit stage diameter (D.sub.CORE). The high-pressure compressor defines a high-pressure compressor exit area (A.sub.HPCExit) in square inches. The gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn.sub.Total) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: Fn.sub.TotalEGT/(A.sub.HPCExit.sup.21000). The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of L.sub.CORE/D.sub.CORE is from 2.1 to 4.3.

A turbomachine engine includes an engine core including a high-pressure compressor, a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine. The engine core has a length (L.sub.CORE), and the high-pressure compressor has an exit stage diameter (D.sub.CORE). A high-pressure shaft is coupled to the high-pressure compressor and the high-pressure turbine. The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of L.sub.CORE/D.sub.CORE is from 2.1 to 4.3.

A gas turbine engine includes a fan, an engine core, a fan case housing the fan and the engine core, a plurality of outlet guide vanes extending between the engine core and the fan case, and an acoustic spacing. Relationships between acoustic spacing and a high-speed shaft rating allow for a gas turbine engine that reduces noise emissions while maintaining high performance.

A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The acoustic spacing parameter, in combination with composite fan blades that include one or both of a fan leading edge to trailing edge compression factor (FLTCF) and a fan leading edge to trailing edge opening ratio (FLTOR), provide improved performance and reduced acoustic noise.

A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The acoustic spacing parameter, in combination with composite fan blades that include one or both of a fan leading edge to trailing edge compression factor (FLTCF) and a fan leading edge to trailing edge opening ratio (FLTOR), provide improved performance and reduced acoustic noise.

A turbomachine engine includes an engine core including a high-pressure compressor, a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine. The engine core has a length (L.sub.CORE), and the high-pressure compressor has an exit stage diameter (D.sub.CORE). A high-pressure shaft is coupled to the high-pressure compressor and the high-pressure turbine. The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of L.sub.CORE/D.sub.CORE is from 2.1 to 4.3.

A turbomachinery engine includes a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine includes four rotating stages. The low-pressure turbine includes an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine. In some instances, the area ratio is within a range of 2.0-5.1. Additionally (or alternatively) the low-pressure turbine includes an area-EGT ratio within a range of 1.05-1.6.

The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor. The unducted fan propulsors can also include a first VPF parameter within a range of 0.10 to 0.40 and defined as the hub-to-tip radius ratio divided by the fan pressure ratio and/or a second VPF parameter within a range of 1-30 lbf/in.sup.2 and defined as the bearing spanwise force divided by the fan area. In certain examples, the unducted fan propulsor further includes a pitch change mechanism and/or a gearbox.