Compressor with segmented inner shroud for an axial turbine engine

Patent metadata

Field	Value
Publication number	US-10690147-B2
Application number	US-201815872242-A
Country	US
Kind code	B2
Filing date	Jan 16, 2018
Priority date	Jan 26, 2017
Publication date	Jun 23, 2020
Grant date	Jun 23, 2020

Abstract

An assembly for a turbojet, wherein the assembly includes an outer shroud and an inner shroud that are concentric, wherein the inner shroud is segmented and includes circumferential clearances between the segments thereof. The assembly additionally includes an annular row of stator vanes connecting the inner shroud to the outer shroud, a drive with a reduction ratio that is intended to be coupled to a fan, and a circuit for cooling and for lubricating the drive. The circuit is configured to heat up at least the outer shroud during the operation of the turbine engine such as to circumferentially reduce the circumferential clearances between the segments.

First claim

What is claimed is:

1. An assembly for an axial turbine engine, said assembly comprising:
an outer shroud and an inner shroud that are coaxial, the inner shroud comprising segments and circumferential clearances between the segments thereof;
an annular row of stator vanes connecting the inner shroud to the outer shroud;
a reduction ratio drive intended to be coupled to a fan; and
a cooling circuit for the reduction ratio drive, which cooling circuit is configured to heat up the outer shroud during the operation of the axial turbine engine such as to circumferentially reduce the circumferential clearances between the segments.

2. The assembly according to claim 1 further comprising a circular splitter to which the outer shroud is fixed and surrounding the outer shroud, the cooling circuit being structurally and functionally configured to deice said circular splitter.

3. The assembly according to claim 1 , wherein the outer shroud comprises a composite material with an organic resin and fibres, the fibres including carbon fibres and/or glass fibres.

4. The assembly according to claim 1 , wherein at least one stator vane of the annular row of stator vanes includes metal and/or a shape memory material.

5. The assembly according to claim 1 , wherein the cooling circuit is structurally and functionally configured to heat up the stator vanes in order to deice the stator vanes.

6. The assembly according to claim 1 , wherein the cooling circuit comprises a heat exchanger in thermal contact with at least one stator vane of the annular row of stator vanes and/or in thermal contact with the outer shroud in order to deice the at least one stator vane and/or the outer shroud.

7. The assembly according to claim 1 , wherein at least one circumferential clearance of the circumferential clearances forms an inclined straight line in space with respect to an axis of rotation of the axial turbine engine, wherein the axis of rotation and the inclined straight line are not co-planar.

8. The assembly according to claim 1 , wherein the inner shroud has an external annular surface with an outer diameter variation along an axial direction.

9. The assembly according to claim 1 , wherein the stator vanes have a radial height variation from upstream to downstream.

10. The assembly according to claim 1 , wherein the inner shroud surrounds the reduction ratio drive.

11. A turbojet engine, said engine comprising:
a fan;
a compressor; and
an assembly, wherein the assembly comprises:
an outer shroud and an inner shroud that are coaxial, the inner shroud comprising segments and circumferential clearances between the segments thereof;
an annular row of stator vanes connecting the inner shroud to the outer shroud;
a reduction ratio drive for driving the fan;
a cooling circuit for the reduction ratio drive, wherein the cooling circuit is configured to heat up the outer shroud during the operation of the turbine engine so as to circumferentially modify the circumferential clearances between the segments.

12. The turbojet engine according to claim 11 further comprising a fan hub frame supporting the fan and the outer shroud, the reduction ratio drive being placed in the fan hub frame.

13. The turbojet engine according to claim 11 , wherein the reduction ratio drive is a gear reduction suitable for reducing a fan rotation speed with respect to a first shaft.

14. The turbojet engine according to claim 11 , wherein the reduction ratio drive is suitable for increasing a rotation speed of the compressor with respect to a turbine shaft.

15. The turbojet engine according to claim 11 , wherein the fan and the compressor rotate at a respective rotation speed, wherein the reduction ratio drive is coupled to the fan and to the compressor, and wherein the rotation speed of the compressor is greater than the rotation speed of the fan.

16. The turbojet engine according to claim 11 , wherein the inner shroud and the outer shroud define therebetween an inlet of the compressor.

17. The turbojet engine according to claim 11 , wherein the turbojet engine has a bypass ratio greater than or equal to at least one of 5, or 8, or 12, or 15.

18. A method for controlling circumferential clearances between turbojet engine inner shroud segments, the turbojet engine comprising an assembly that includes a reduction ratio drive driving a fan, an inner shroud segmented by circumferential clearances, an outer shroud, and an annular row of stator vanes connecting the inner shroud to the outer shroud, the method comprising the following steps:
operating the turbojet engine, and
exchanging heat, wherein the turbojet engine also comprises a circuit for heat exchange with two or three of:
the reduction ratio drive,
the outer shroud or,
the stator vanes;

wherein during the step for exchanging heat, the circuit reducing the circumferential clearances between the inner shroud segments.

19. The method according to claim 18 , wherein during the step of exchanging heat, the circumferential clearances are closed by a change in shape of a shape memory material, the inner shroud segments touching one another along the circumference.

20. The method according to claim 18 further comprising a step of stopping the turbojet engine, the circumferential clearances being open to a greater extent during the stopping step than during the step for heat exchange.

Assignees

• Safran Aero Boosters Sa

Inventors

• Hiernaux Stephane

Classifications

• F01D25/02 Primary

  De-icing means for engines having icing phenomena · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F01D9/041

  using blades (F01D5/148 takes precedence) · CPC title

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title

• F05D2300/603

  Composites; e.g. fibre-reinforced · CPC title