A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor section including a propulsor that delivers flow to a core flowpath and a compressor section including first and second compressors. The core flowpath passes through the first compressor. The core flowpath in the first compressor has an outer diameter relative to the engine longitudinal axis. The outer diameter has a slope angle relative to the axis.

An impeller includes a shroud including an inlet, a hub facing the shroud, and a plurality of blades disposed between the hub and the shroud and arranged circumferential direction along a circumference of the inlet. Each of the plurality of blades has a slot which is positioned adjacent to the inlet.

An air blower and a combustion device including the air blower are provided.

An air blower includes an impeller housed in a casing and rotating. The impeller includes multiple blades to form multiple inter-blade passages, and a central space portion. When the impeller rotates, air flowing into the central space portion through an air intake port of the casing passes through the inter-blade passages at an outward side. To reduce noise generated when the air passes through the inter-blade passages, each blade includes: multiple first concavo-convex portions, with a first surface side being concave and a second surface side on the other side being convex, from among a first and a second surface corresponding to a front and a rear surface of each blade; and multiple second concavo-convex portions, with the second surface side being concave and the first surface side being convex, contrary to the first concavo-convex portions.

A lifting fan for hovercraft of the present disclosure comprises an upper shroud that is higher in its center and lower on its outer side, an air inlet part formed in the center of the upper shroud, a lower shroud that is higher in its center and lower on its outer side, a rotating shaft coupled to the center of the lower shroud, a plurality of blowing blades formed between the upper shroud and the lower shroud, a blowing passage formed with an inclination between the upper shroud, the lower shroud, and the plurality of blowing blades, and an air outlet part formed on the outer side of the upper shroud, on the outer side of the lower shroud, and at distal ends of the plurality of blowing blades, and thus has the effect of being excellent in the efficiency of air flow and of reducing the amount of noise and vibration generated in the process in which the air introduced through the air inlet part is discharged through the air outlet part by way of the inclined blowing passage.

A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A three-bladed impeller providing a cooling airflow in air, with increased heat-dissipating efficiency and reduced noise includes a hub and three blades, the blades are arranged around the hub. Each blade is arched along its axial length from the front of fan to the back and also arched radially from the hub end of each blade to the outside tip. The back edge of each blade includes first and second slots, arranged alternately, the width of each first slot is λ1, the width of each of each second slot is λ2, the comparative sizes between λ2 and λ1 are in a ratio range of 1.6:1 to 1.8:1 (λ2:λ1).

A centrifugal compressor comprises an impeller including a hub and a plurality of blades. The hub is provided with a through hole. The hub has an external radial surface having an inner external radial surface and an outer external radial surface. The outer external radial surface is formed closer to a back surface than an imaginary curved surface having as a radius a radius of curvature of the inner external radial surface at a radially outer edge thereof.

A propeller fan includes a shaft portion disposed on a rotation axis, and a blade disposed on an outer peripheral side of the shaft portion and including a leading edge and a trailing edge. The blade includes a negative pressure surface in which a plurality of recesses are formed, and the plurality of recesses include a first recess and a second recess disposed on the trailing edge side than the first recess in a circumferential direction about the rotation axis as a center. The first recess has a depth larger than a depth of the second recess.

A propeller fan includes a shaft portion disposed on a rotation axis, and a blade disposed on an outer peripheral side of the shaft portion and including a leading edge and a trailing edge. The blade includes a negative pressure surface in which a plurality of recesses are formed, and the plurality of recesses include a first recess and a second recess disposed on the trailing edge side than the first recess in a circumferential direction about the rotation axis as a center. The first recess has a depth larger than a depth of the second recess.

A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.