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.

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 TiAl alloy impeller is composed of a TiAl alloy and is mountable on a vehicle turbocharger. The TiAl alloy has a lamellar structure in which an .sub.2 layer containing Ti.sub.3Al and a layer containing TiAl are alternately stacked. The TiAl alloy impeller includes a shaft portion and blades, each of the blades includes an outer edge, and a region of each of the blades including the outer edge has a processed surface. The processed surface is provided with one or more grooves. At least one of the grooves satisfies a relation of D.sub.g>2R.sub.g or D.sub.g>W.sub.g, where D.sub.g represents a depth of the groove and R.sub.g represents a curvature radius of the groove at a bottom surface of the groove or W.sub.g represents a width of the groove. Thus, a TiAl alloy impeller having high performance as an impeller is provided.

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.

A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given

[00001]$\frac{N_{\mathrm{FB}}{D}_{\mathrm{FT}}}{L_{\mathrm{AXIAL}}\left(\frac{R_{\mathrm{TB}}}{N_{\mathrm{FB}}}\right)}.$

N.sub.FB is a number of the fan blades, D.sub.FT is a fan tip diameter of the fan blades, R.sub.TB is a thrust bearing radius of the radial thrust bearings, and L.sub.AXIAL is an axial length from a fan hub tip to the fan bearings.

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.

A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A set of blades are circumferentially arranged in the turbine section and the compressor section. A set of dovetails mounts the set of blades to a disk, rotated about the engine centerline. Each dovetail can include a set of lobes defining complementary recesses, collectively defining a neck for mounting to the disk.

A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A set of blades are circumferentially arranged in the turbine section and the compressor section. A set of dovetails mounts the set of blades to a disk, rotated about the engine centerline. Each dovetail can include a first upper lobe and a first lower lobe defining a first intervening recess, and a second upper lobe and a second lower lobe defining a second intervening recess, collectively defining a neck for mounting to the disk.

A turbomachine engine can include a fan assembly, a vane assembly, a core engine, a gearbox, and an overall engine efficiency rating. The fan assembly can include a plurality of fan blades. The vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include a low-pressure turbine. The gearbox includes an input and an output. The input of the gearbox is coupled to the low-pressure turbine of the core engine and comprises a first rotational speed, the output of the gearbox is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.0-4.0.

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.