ROTATING BLADE

摘要

The present invention relates to a rotating blade for a turbomachine, in particular a compressor stage or a turbine stage of a gas turbine, particularly of an aircraft engine, having a blade element (10) for deflecting the flow, with a pressure side and a suction side (11), these sides being joined by a leading edge and a trailing edge (12, 13), wherein a stacking axis (S) of the blade element (10), in the radial direction (R) over a radius r from a root of a blade element at r=0 to a tip of a blade element at r=H, has a course x(r) in a first downstream direction (X) perpendicular to the radial direction (R) and parallel to a principal axis of the turbomachine and has a course y(r) in a second direction (Y) perpendicular to the radial direction (R) and to the first direction (X).

主权项

1. A rotating blade for a compressor stage or a turbine stage of an aircraft engine, comprising:
a blade element (10) for deflecting the flow, with a pressure side and a suction side (11), these sides being joined by a leading edge and a trailing edge (12, 13), wherein a stacking axis (S) of the blade element (10), in the radial direction (R) over a radius r from a root of a blade element at r=0 to a tip of a blade element at r=H,
has a course x(r) in a first downstream direction (X) perpendicular to the radial direction (R) and parallel to a principal axis of the turbomachine, this course deviating from a superimposed course xl(r)+xs(r), which is composed of an addition of a particularly positive, linear course xl(r)=ax·(r/H+bx) and a local sine function,xs(r)={A·[sin(B·(rH+C))+1]⇔-π2≤B·(rH+C)≤3·π20⇔anyother by a maximum of 0.100·ax·(1+bx), at least in a radial region between r=0.200·H and r=0.400·H; and
has a course y(r) in a second direction (Y) perpendicular to the radial direction (R) and to the first direction (X), this course deviating from a particularly negative, linear course ay·(r/H+by) by a maximum of 0.100·ay·(1+by), at least in the radial region between r=0.200·H and r=0.400·H.