Single-piece blisk for turbomachine fan comprising an upstream and/or downstream recess making its blades more flexible

Patent metadata

Field	Value
Publication number	US-9863252-B2
Application number	US-201514942106-A
Country	US
Kind code	B2
Filing date	Nov 16, 2015
Priority date	Nov 17, 2014
Publication date	Jan 9, 2018
Grant date	Jan 9, 2018

Abstract

The invention relates to a single-piece fan blisk ( 14 ) for a turbojet, this blisk ( 14 ) comprising a hub ( 16 ) with a general shape of revolution about a rotation axis (AX) and comprising an external peripheral face ( 18 ) of revolution extending radially towards the rotation axis at an upstream face ( 19 ) and a downstream face both of which are in the form of rings. This hub ( 16 ) supports radially oriented blades ( 17 ) at its external peripheral face ( 18 ), each comprising a base through which it is connected to this external peripheral face ( 18 ), each blade ( 17 ) having a leading edge ( 22 ) and a trailing edge ( 23 ) that are radially oriented. The spacing between the upstream face ( 19 ) and the downstream face of the hub ( 16 ) along the rotation axis (AX) is less than the distance separating the leading edge ( 22 ) from the trailing edge ( 23 ) of each blade.

First claim

The invention claimed is:

1. Single-piece fan blisk ( 14 ) for a turbofan such as a turbojet, this single-piece blisk ( 14 ) comprising a hub ( 16 ) with a general shape of revolution about a rotation axis (AX), this hub ( 16 ) comprising an external peripheral face ( 18 ) extending radially towards the rotation axis (AX) at an upstream face ( 19 ) and a downstream face ( 21 ) both of which are in the form of rings, this hub ( 16 ) supporting blades ( 17 ) each comprising a base through which it is connected to an external peripheral face ( 18 ) and a leading edge ( 22 ) and a trailing edge ( 23 ) that are radially oriented, in which the spacing between the upstream face ( 19 ) and the downstream face ( 21 ) (AX) is less than the distance separating the leading edge ( 22 ) from the trailing edge ( 23 ) of each blade along the rotation axis, wherein:
the upstream face ( 19 ) of the hub ( 16 ) is located along the rotation axis (AX) between the leading edges ( 22 ) of the blades and the trailing edges ( 23 ) of the blades, each blade ( 17 ) comprising a first prolongation on the side of its leading edge ( 22 ) towards the rotation axis (AX) through which it is connected to the upstream face ( 19 );
and/or in that the downstream face ( 21 ) of the hub ( 16 ) is located along the rotation axis (AX) between the leading edges ( 22 ) of the blades and the trailing edges ( 23 ) of the blades, each blade ( 17 ) comprising a second prolongation (P) on the side of its trailing edge ( 23 ) towards the rotation axis (AX) through which it is connected to the downstream face ( 21 ).

2. Blisk according to claim 1 , in which the ratio (L 1 /L 2 ) of the length (L 1 ) separating the leading edge ( 22 ) of the blade from the upstream face ( 19 ) of the hub ( 16 ) divided by the length (L 2 ) separating the leading edge ( 22 ) from the trailing edge ( 23 ) of the blade, is between two tenths and four tenths.

3. Blisk according to claim 1 , in which the ratio (L 1 ′/L 2 ) of the length (L 1 ′) separating the trailing edge ( 23 ) of the blade from the downstream face ( 21 ) of the hub ( 16 ) divided by the length (L 2 ) separating the leading edge ( 22 ) from the trailing edge ( 23 ) of the blade, is between two tenths and four tenths.

4. Blisk according to claim 1 , in which the ratio of the sum of the length (L 1 ) separating the leading edge ( 22 ) of the blade from the upstream face ( 19 ) of the hub and the length (L 1 ′) separating the trailing edge ( 23 ) of the blade from the downstream face ( 21 ) of the hub, divided by the length (L 2 ) separating the leading edge ( 22 ) from the trailing edge ( 23 ) of the blade, is between two tenths and four tenths.

5. Turbofan fan comprising a disk according to claim 1 .

6. Turbojet type aircraft engine, comprising a single-piece blisk according to claim 1 .

Assignees

• Snecma

Inventors

• Merlot Damien
• Ide Etienne
• Joly Philippe Gérard Edmond

Classifications

• F05D2240/304

  related to the trailing edge of a rotor blade · CPC title

• F05D2260/941

  particularly aimed at mechanical or thermal stress reduction · CPC title

• F04D29/321

  for axial flow compressors · CPC title

• F05D2230/53

  by integrally manufacturing a component, e.g. by milling from a billet or one piece construction · CPC title

• F04D29/329

  Details of the hub · CPC title