Blisk with low stresses at blade root, preferably for an aircraft turbine engine fan

Patent metadata

Field	Value
Publication number	US-9765637-B2
Application number	US-201414466044-A
Country	US
Kind code	B2
Filing date	Aug 22, 2014
Priority date	Sep 9, 2013
Publication date	Sep 19, 2017
Grant date	Sep 19, 2017

Abstract

An assembly for an aircraft turbine engine including an integral element including a disk and a plurality of blades is provided. Each blade has a connection zone connecting the blade to the disk, the connection zone including a first part arranged at the external flowpath delimitation surface provided on the disk, and at least one second end part arranged at a recess formed in the disk along the axial extension of the external surface, the average fillet radius defined by the second end part being larger than the fillet radius defined by the first part. The assembly further includes a flowpath reconstruction part formed in the aerodynamic continuity of the external surface, so as to cover the recess.

First claim

The invention claimed is:

1. An assembly for an aircraft turbine engine comprising:
an integral element including a disk and a plurality of blades projecting from the disk,
wherein at least at one of two opposite surfaces of at least one of the blades, the blade has a connection zone connecting said blade to the disk, said connection zone comprising a first part arranged at an external flowpath delimitation surface provided on the disk, and at least a second end part arranged at a recess in the disk located along an axial extension of said external flowpath delimitation surface, an average fillet radius defined by said second end part being larger than an average fillet radius defined by said first part, and
wherein the assembly comprises a flowpath reconstruction part formed in an aerodynamic continuity of said external flowpath delimitation surface, so as to cover said recess, a portion of said flow path reconstruction part being arranged around a portion of the two opposite surfaces of the at least one of the blades.

2. The assembly according to claim 1 , wherein a ratio between the average fillet radius defined by said second end part and the average fillet radius defined by said first part is between 1 and 5.

3. The assembly according to claim 1 , wherein said flowpath reconstruction part has an annular shape, and an end of said flowpath reconstruction part has notches through which blades can pass.

4. The assembly according to claim 1 , wherein said flowpath reconstruction part is fixed onto the integral element.

5. The assembly according to claim 4 , wherein said flowpath reconstruction part is fixed onto an attachment flange of the integral element.

6. The assembly according to claim 1 , wherein said connection zone comprises a single second end part, arranged at a trailing edge or leading edge of the blade.

7. The assembly according to claim 1 , wherein said connection zone comprises two second end parts, arranged on each side of said first part, at a trailing edge and a leading edge of the blade respectively.

8. The assembly according to claim 7 , wherein each second end part of the connection zone extends over a length equal to 15 to 35% of a total length of the connection zone.

9. The assembly according to claim 1 , wherein said integral element is a blisk.

10. A fan for an aircraft turbine engine comprising at least an assembly according to claim 1 .

11. An aircraft turbine engine comprising a fan according to claim 10 .

Assignees

• Snecma

Inventors

• Merlot Damien
• Sers Remi

Classifications

• F01D11/008

  by spacer elements between the blades, e.g. independent interblade platforms · CPC title

• F01D5/143

  Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour · CPC title

• F01D5/021

  for flow machines or engines with only one axial stage (for more than one stage F01D5/06) · CPC title

• F01D5/3069 Primary

  between two discs or rings · CPC title

• F05D2240/24

  for turbines · CPC title