Turbine engine blade having improved stacking law

Patent metadata

Field	Value
Publication number	US-9650896-B2
Application number	US-201113993079-A
Country	US
Kind code	B2
Filing date	Dec 15, 2011
Priority date	Dec 15, 2010
Publication date	May 16, 2017
Grant date	May 16, 2017

Abstract

A turbine engine blade, including an airfoil which extends radially between a blade root and an airfoil tip, axially between a leading edge and a trailing edge, and tangentially between a pressure side and a suction side, the profile of the blade having a series of basic profiles, in a form of a vane section, stacked on one another along a stacking line connecting the center of gravity of all the vane sections. The projection of the stacking line of the airfoil on at least one plane extending radially from the blade root includes a double tangential inversion of the direction of the curvature thereof, located in the last thirty percent of the height of the airfoil, the projection plane being positioned substantially perpendicular to the chord of the blade.

First claim

The invention claimed is:

1. A turbomachine blade, comprising:
an airfoil which extends radially between a blade root and an airfoil tip, axially between a leading edge and a trailing edge, and tangentially between a pressure face and a suction face,
a profile of the blade including a series of elementary profiles, in a form of vane sections, stacked on one another along a stacking line joining centers of gravity of all of the sections,
wherein a projection of the stacking line of the airfoil onto a plane extending radially from the blade root comprises an axial inversion of curvature,
wherein the axial inversion is situated in a last 30 percent of height of the airfoil, the plane of projection of the axial inversion being oriented substantially parallel to a chord of the blade,
wherein directions of curvatures of the leading edge and the trailing edge change at a radial position of the axial inversion, and the leading edge and the trailing edge are both axially offset at a same amount and in a same direction from the stacking line at the radial position of the axial inversion, and
wherein the axial inversion is situated over a last 10 percent of the height of the airfoil.

2. The blade as claimed in claim 1 , further comprising a tangential inversion of curvature, the plane of projection of the tangential inversion being oriented substantially perpendicular to the chord of the blade.

3. The blade as claimed in claim 2 , wherein the projection includes first and second tangential inversions.

4. The blade as claimed in claim 3 , wherein the first and second tangential inversions are situated in the last 30 percent of height of the airfoil, and the first tangential inversion deforms in a direction of the suction face and the second tangential inversion deforms in a direction of the pressure face.

5. A turbomachine compressor comprising at least one rotor wheel made up of blades as claimed in claim 1 .

6. A turbomachine comprising a compressor as claimed in claim 5 .

7. A turbomachine turbine comprising at least one rotor wheel made up of blades as claimed in claim 1 .

8. A turbomachine blade, comprising:
an airfoil which extends radially between a blade root and an airfoil tip, axially between a leading edge and a trailing edge, and tangentially between a pressure face and a suction face,
a profile of the blade including a series of elementary profiles, in a form of vane sections, stacked on one another along a stacking line joining centers of gravity of all of the sections,
wherein a projection of the stacking line of the airfoil onto a plane extending radially from the blade root comprises a double axial inversion of curvature over a last 30 percent of height of the airfoil,
wherein the plane of projection of each axial inversion of the double axial inversion is oriented substantially parallel to a chord of the blade, and
wherein directions of curvatures of the leading edge and the trailing edge change at a radial position of each axial inversion of the double axial inversion, and the leading edge and the trailing edge are both axially offset at a same amount and in a same direction from the stacking line at the radial position of each axial inversion of the double axial inversion.

9. The blade as claimed in claim 8 , further comprising a tangential inversion of curvature, the plane of projection of the tangential inversion being oriented substantially perpendicular to the chord of the blade.

10. The blade as claimed in claim 9 , wherein the projection includes first and second tangential inversions.

11. The blade as claimed in claim 10 , wherein the first and second tangential inversions are situated in the last 30 percent of height of the airfoil, and the first tangential inversion deforms in a direction of the suction face and the second tangential inversion deforms in a direction of the pressure face.

12. A turbomachine compressor comprising at least one rotor wheel made up of blades as claimed in claim 8 .

13. A turbomachine comprising a compressor as claimed in claim 12 .

14. A turbomachine turbine comprising at least one rotor wheel made up of blades as claimed in claim 8 .

Assignees

• Pesteil Agnes
• Perrot Vincent
• Snecma

Inventors

• Pesteil Agnes
• Perrot Vincent

Classifications

• F01D5/14 Primary

  Form or construction (selecting particular materials, measures against erosion or corrosion F01D5/28) · CPC title

• F05D2250/713

  inflexed · CPC title

• F04D29/324

  Blades · CPC title

• F01D5/141 Primary

  Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147) · CPC title

• F05D2240/307

  related to the tip of a rotor blade · CPC title