Outlet guide vane for an aircraft turbomachine, comprising a lubricant cooling passage equipped with flow disturbing studs with simplified manufacturing

Patent metadata

Field	Value
Publication number	US-10975722-B2
Application number	US-201816614852-A
Country	US
Kind code	B2
Filing date	May 17, 2018
Priority date	May 22, 2017
Publication date	Apr 13, 2021
Grant date	Apr 13, 2021

Abstract

The invention relates to a guide vane intended to be arranged in all or part of an air flow of an aircraft bypass turbomachine fan, the vane comprising an aerodynamic part equipped with at least one interior lubricant cooling passage delimited in part by an intrados wall and an extrados wall of the vane, there being flow-disturbing lugs, made as one piece with one of the intrados and extrados walls, passing across the passage. According to the invention, in any plane of section passing orthogonally through the lugs, the space defined between these lugs has a geometry defined exclusively by a set of annulus shapes of the same dimensions, partially overlapping one another and each in part delimiting at least two of these lugs.

First claim

What is claimed is:

1. A guide vane ( 24 ) for being arranged in all or part of an airflow of a fan ( 15 ) of an aircraft dual flow turbomachine, the guide vane comprising a root ( 34 ), an edge ( 36 ), as well as a flow straightening aerodynamic part ( 32 ) arranged between the root and the edge of the vane, said aerodynamic part of the vane including at least one lubricant cooling internal passage ( 50 a , 50 b ) partly delimited by a front side wall ( 70 ) and a back side wall ( 72 ) of the vane, flow disturbing studs ( 80 ) made as a single piece with one of the front side ( 70 ) and back side ( 72 ) walls passing through said passage, wherein in any section plane orthogonally passing through the studs ( 80 ), a space ( 94 ) defined between these studs has a geometry exclusively defined by a set of rings ( 92 ) with a same dimension, partially covering each other and each partly delimiting at least two of these studs ( 80 ).

2. The vane according to claim 1 , wherein in said section plane, each stud ( 80 ) is delimited by four sides ( 82 ) each having an arc of circle shape.

3. The vane according to claim 1 , wherein said set of rings ( 92 ) is made by alternating first and second rows of rings (RA 1 , RA 2 ) partially overlapping each other and succeeding each other along a span direction ( 25 ) of the vane, each first row (RA 1 ) preferably comprising a number N of rings ( 92 ) and each second row (RA 2 ) preferably comprising a number N−1 of rings ( 92 ), the centres ( 88 ) of these rings being disposed in a staggered manner so as to define alternating first rows of studs (RP 1 ) each comprising a number N+1 of studs ( 80 ) and second rows of studs (RP 2 ) each comprising a number N of studs ( 80 ).

4. The vane according to claim 1 , wherein the rings ( 92 ) of the set have an external diameter between 20 and 50 mm and an internal diameter between 5 and 20 mm.

5. The vane according to claim 1 , wherein said one of the front side and back side walls ( 70 , 72 ) equipped with the studs ( 80 ) is part of a body ( 32 a ) of the vane, or a cap ( 32 b ) for closing this body.

6. The vane according to claim 1 , wherein the other of the elements from the front side and back side walls ( 70 , 72 ) is made as a single piece with additional flow disturbing studs ( 80 ′), and in that in any section plane orthogonally passing through the additional studs ( 80 ′), the space ( 94 ′) defined between these additional studs has a geometry exclusively defined by a set of additional rings with a same dimension, partially covering each other and each partly delimiting at least two of these additional studs ( 80 ′).

7. The vane according to claim 1 , wherein the additional studs ( 80 ′) penetrate the space ( 94 ) defined between the studs ( 80 ), and in that the studs ( 80 ) penetrate the space ( 94 ′) defined between the additional studs ( 80 ′).

8. An aircraft turbomachine ( 1 ), preferably a turbojet engine, comprising a plurality of guide vanes ( 24 ) according to claim 1 , arranged downstream or upstream of a fan ( 15 ) of the turbomachine.

9. A method for manufacturing a guide vane ( 24 ) for being arranged in all or part of an airflow of a fan ( 15 ) of an aircraft dual flow turbomachine, the guide vane comprising a root ( 34 ), an edge ( 36 ), as well as a flow straightening aerodynamic part ( 32 ) arranged between the root and the edge of the vane, said aerodynamic part of the vane including at least one lubricant cooling internal passage ( 50 a , 50 b ) partly delimited by a front side wall ( 70 ) and a back side wall ( 72 ) of the vane, flow disturbing studs ( 80 ) made as a single piece with one of the front side ( 70 ) and back side ( 72 ) walls passing through said passage,
comprising a step of making the studs ( 80 ) by repeatedly machining said one of the front side ( 70 ) and back side ( 72 ) walls using a chamfering cutter ( 86 ), this step being implemented such that the remaining parts not machined by said chamfering cutter form said studs ( 80 ).

10. The method according to claim 9 , wherein the repeated machining using the chamfering cutter ( 86 ) is made along parallel machining axes ( 88 , 88 a , 88 b ), preferably arranged in a staggered manner.

Assignees

• Safran Aircraft Engines

Inventors

• Marquie Dimitri Daniel Gabriel
• Leguery Geoffray Henri Robert
• Senejoux Pierre François

Classifications

• B23B51/04

  {Drills} for trepanning · CPC title

• F05D2250/141

  circular · CPC title

• F28D1/035

  with U-flow or serpentine-flow inside the conduits (F28D1/0341 takes precedence) · CPC title

• F05D2250/231

  cylindrical · CPC title

• F01D9/065 Primary

  Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids (see also F01D25/16, F01D25/24 and F01D25/26) · CPC title