Rotating blade
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 for deflecting the flow, with a pressure side and a suction side, these sides being joined by a leading edge and a trailing edge, wherein a stacking axis of the blade element, in the radial direction 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 perpendicular to the radial direction and parallel to a principal axis of the turbomachine and has a course y(r) in a second direction perpendicular to the radial direction and to the first direction.
1. A rotating blade for a compressor stage or a turbine stage of an aircraft engine, comprising:
a blade element for deflecting the flow, with a pressure side and a suction side, these sides being joined by a leading edge and a trailing edge, wherein a stacking axis of the blade element, in the radial direction 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 perpendicular to the radial direction and parallel to a principal axis of the turbomachine, this course deviating from a superimposed course x l (r)+x s (r), which is composed of an addition of a particularly positive, linear course x l (r)=a x ·(r/H+b x ) and a local sine function,
x
s
(
r
)
=
{
A
·
[
sin
(
B
·
(
r
H
+
C
)
)
+
1
]
⇔
-
π
2
≤
B
·
(
r
H
+
C
)
≤
3
·
π
2
0
⇔
else
by a maximum of 0.100·a x ·(1+b x ), 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 perpendicular to the radial direction and to the first direction, this course deviating from a particularly negative, linear course a y ·(r/H+b y ) by a maximum of 0.100·a y ·(1+b y ), at least in the radial region between r=0.200·H and r=0.400·H,
wherein:
A<0,
1.800·π≤B≤2.200·π or 6.300·π≤B≤7.000·π,
−0.300≤C≤−0.200 or −0.100≤C≤−0.050,
0<a x ,
a y <0,
a X ≤0.036·H,
−0.100≤b X ≤0.100,
−0.006·H≤a y ,
−0.100≤b y ≤0.100, and
−0.005·H≤A, and
wherein H is the radial height of the blade element.
2. The rotating blade according to claim 1 , wherein the radial height H of the blade element is between 150 mm and 200 mm.
3. The rotating blade according to claim 1 , wherein the blade has a shroud at the tip of the blade element.
4. The rotating blade according to claim 1 , wherein the at least one rotating blade is in a compressor stage or a turbine stage of an aircraft engine.
5. The rotating blade according to claim 1 , wherein the rotating blade is fastened in detachable manner to a rotor.
6. The rotating blade according to claim 5 , wherein the rotating blade is in an aircraft engine, having at least one compressor stage or one turbine stage.