Impeller blade of rotor of compressor of low-pressure gas turbine (versions)

FIELD: engines. SUBSTANCE: invention relates to engine building. Second stage impeller vane with disc with slots and blade ring with frontal line of grid of feather profiles, in low-pressure compressor rotor of gas turbine engine (GTE), comprising flow part confined in peripheral outline by engine housing, having power turbine, comprises stem and wing with convex-concave profile. Blade feather is made with helical twisting relative to axis of feather, creating variable on height of feather an angle of γ ust of installation of feather profile, defined as angle between common tangent, connecting leading and trailing edges, forming a chord profile, and front line of profile grid in flat scanning of cylindrical section of blade rim, having in root section of feather a value γ ust.k = (65.2÷73.2)°, and in peripheral section a value U ust.k = (35.8÷43.8)°. Blade has variable on height of feather an angle γ of installation of feather profile relative to front line of profile grid of blade rim, decreasing with radial distance from rotor axis with gradient (G) y.p = (196.3÷282.2) [deg/m]. Blade feather has leading and trailing edges diverging to peripheral end with chord increase gradient G y.x = (7.4÷10.7)·10 -2 [m/m]. Blade feather has, variable on width and height, blade thickness. Maximum thickness of blade feather profile is highest in root section and decreases on height of feather to peripheral end with gradient G y.t = (1.14÷1.63)·10 -2 [m/m]. EFFECT: technical result consists in improved geometrical configuration, spatial stiffness, structural and aerodynamic parameters of second stage of gas turbine engine low-pressure compressor rotor shaft, as well as high efficiency and wider range of GDU modes of compressor with longer service life of blade. 25 cl, 3 dwg