Fan blade with zero tip dihedral at the head
A blade ( 3 ) of a fan ( 1 ), made of composite material, of a turbomachine. On a portion ( 15 ) of the blade ( 7 ) which extends from a lower limit ( 16 ) located at a predetermined distance (d) from the shank ( 6 ) of the blade equal to at least 80% of the predetermined height (h) to the top ( 11 ) of the blade ( 7 ), a dihedral angle (D) measured at least one predefined point on the chord of the blade ( 3 ), located on the chord of the blade ( 3 ) at the upstream end ( 13 a ) of the nose ( 13 ) of the shield ( 12 ), is greater than or equal to −3° and less than or equal to 0°.
1. A turbomachine fan blade comprising:
a structure of composite material comprising a fibrous reinforcement obtained by three-dimensional weaving and a matrix in which the fibrous reinforcement is embedded, the composite material structure forming an airfoil with an aerodynamic profile comprising a leading edge, a trailing edge, a root configured to be attached to a fan disk of the turbomachine fan and a shank extending between the root and the airfoil, the airfoil having a predetermined height between the shank and a top of the airfoil along a stacking axis; and
a metallic shield applied and attached to the leading edge of the airfoil, the metallic shield comprising a nose attached to the leading edge and fins that delimit with the nose a cavity housing the leading edge of the airfoil;
a chord of the blade being defined, in a plane normal to the stacking axis, between an upstream end of the nose of the shield and the trailing edge;
wherein the airfoil further comprises a portion, which extends from a lower limit to the top of the airfoil, wherein the lower limit is located at a predetermined distance from the shank at least equal to 80% of the predetermined height,
wherein a dihedral angle is greater than or equal to −3° and less than or equal to 0° in the portion of the airfoil over a distance at least equal to 10% of the predetermined height,
wherein the dihedral angle is a local direction difference between a gas flow through the turbomachine and the blade in projection in a plane that is axial and tangent to the direction of rotation of the turbomachine,
wherein the dihedral angle is measured along the leading edge, wherein the dihedral angle changes sign at least one time along the leading edge from the shank to the top of the airfoil.
2. The turbomachine fan blade according to claim 1 , wherein the dihedral angle is greater than or equal to −3° and less than or equal to 0° at several predefined points of the chord of the blade within the portion of the blade.
3. The turbomachine fan blade according to claim 1 wherein, within the portion, the dihedral angle is greater than or equal to −3° and less than or equal to −1° over a distance at least equal to 15% of the predetermined height.
4. The turbomachine fan blade according to claim 1 wherein, within the portion, the dihedral angle is greater than or equal to −3° and less than or equal to −1° over the entire height of the portion.
5. The turbomachine fan blade according to claim 1 , wherein, within the portion, a variation of the dihedral angle along the stacking axis over a distance equal to 10% of the height of the airfoil is equal to 3° at most.
6. The turbomachine fan blade according to claim 1 , wherein, within the portion, the dihedral angle is greater than or equal to −2° and less than or equal to −1°.
7. The turbomachine fan blade according to claim 1 , wherein the fins extend over a length greater than or equal to 15% and less than or equal to 25% of a chord length of the airfoil.
8. A fan for a turbomachine comprising a plurality of turbomachine fan blades according to claim 1 , the turbomachine fan blades being connected to a fan disk of the fan.
9. The fan according to claim 8 , comprising at most twenty fan blades.
10. The fan according to claim 8 , having an outer diameter of the fan comprised between eighty inches and one hundred inches.
11. The fan according to claim 8 , having an outer diameter of the fan comprised between eighty inches and ninety inches.
12. A turbomachine comprising a fan according to claim 8 .
13. The turbomachine according to claim 12 having a bypass ratio greater than or equal to 10.
14. The turbomachine according to claim 12 having a bypass ratio comprised between 10 and 80 inclusive.
15. An aircraft comprising at least one turbomachine according to claim 12 .