Aircraft turbine engine with planetary or epicyclic gear train

Patent metadata

Field	Value
Publication number	US-11047252-B2
Application number	US-201615560464-A
Country	US
Kind code	B2
Filing date	Mar 22, 2016
Priority date	Mar 26, 2015
Publication date	Jun 29, 2021
Grant date	Jun 29, 2021

Abstract

Aircraft turbine engine comprising a low-pressure spool that comprises a low-pressure shaft ( 24 ), means ( 44 ) for taking off power from said low-pressure shaft, and a fan ( 28 ) that is driven by said low-pressure shaft by means of a reduction gear ( 32 ), said reduction gear comprising at least one first element ( 50 ) that is connected to said low-pressure shaft for conjoint rotation, at least one second element ( 56 ) that is connected to said fan for conjoint rotation, and at least one third element ( 52 ) that is connected to a stator casing of the turbine engine, characterised in that said at least one third element is connected to said stator casing by disengageable connection means ( 60 ), and comprising at least one member that can move from a first position in which said at least one third element is fixedly connected to said stator casing into a second position in which said at least one third element is separated from said stator casing and is free to rotate about said longitudinal axis.

First claim

The invention claimed is:

1. An aircraft turbine engine comprising
a low-pressure spool that comprises a low-pressure shaft that connects a rotor of a low-pressure compressor to a rotor of a low-pressure turbine, and
a high-pressure spool that comprises a high-pressure shaft that connects a rotor of a high-pressure compressor to a rotor of a high-pressure turbine, the low-pressure and high-pressure shafts extending along the same longitudinal axis (A), the turbine engine further comprising a device for removing power from said low-pressure shaft, and
a fan that is driven by said low-pressure shaft by means of
a planetary or epicyclic reduction gear, said reduction gear comprising at least one first element that is connected to said low-pressure shaft for conjoint rotation, at least one second element that is connected to said fan for conjoint rotation, and at least one third element that is connected to a stator casing of the turbine engine, wherein said at least one third element is connected to said stator casing by
a disengageable connection device, said disengageable connection device comprising at least one member that is movable from a first position in which said at least one third element is fixedly connected to said stator casing into a second position in which said at least one third element is separated from said stator casing to be rotatable about said longitudinal axis only when in said second position,
and wherein the first element is a planetary shaft which is centered on the longitudinal axis (A) and which is arranged in an upstream extension of the low pressure shaft, the planetary shaft being rotatable when said at least one third element is separated from said stator casing.

2. The aircraft turbine engine according to claim 1 , wherein said third element is an external ring gear of the reduction gear.

3. The aircraft turbine engine according to claim 2 , wherein the external ring gear is fixedly connected to the stator casing, the stator casing being a stator casing of an inter-duct compartment which separates a primary duct from a secondary duct.

4. The aircraft turbine engine according to claim 1 , wherein said third element is a planet carrier of the reduction gear.

5. The aircraft turbine engine according to claim 4 , wherein the planet carrier is fixedly connected to the stator casing, the stator casing being a stator casing of an inter-duct compartment which separates a primary duct from a secondary duct.

6. The aircraft turbine engine according to claim 1 , wherein said connection device comprises an annular flange that is supported by said third element, said at least one member being movably mounted in at least one stirrup fixed to the stator casing and mounted on said flange.

7. The aircraft turbine engine according to claim 6 , wherein said at least one member, which is a piston, is designed to come into abutment on the annular flange and to clamp said flange when said member is in the first position mentioned above.

8. The aircraft turbine engine according to claim 7 , wherein at least one of said at least one stirrup and said at least one member comprises a support plate made of a material having a predetermined friction coefficient.

9. The aircraft turbine engine according to claim 8 , wherein said at least one member is biased against the annular flange by at least one spring.

10. The aircraft turbine engine according to claim 8 , wherein said at least one member is designed to be moved in translation by means of a screw.

11. The aircraft turbine engine according to claim 7 wherein said at least one member is biased against the annular flange by at least one spring.

12. The aircraft turbine engine according to claim 7 , wherein said at least one member is designed to be moved in translation by means of a screw.

13. The aircraft turbine engine according to claim 6 , wherein the stirrup is pressed against a wall of the stator casing.

14. The aircraft turbine engine according to claim 6 , wherein the at least one stirrup is provided with a plurality of stirrups regularly spaced about the longitudinal axis.

15. The aircraft turbine engine according to claim 1 , wherein said disengageable connection device is connected to a first actuator that is connected to a computer of the turbine engine.

16. The aircraft turbine engine according to claim 15 , wherein the first actuator is a hydraulic actuator.

17. The aircraft turbine engine according to claim 1 , wherein said turbine engine has a bypass ratio of greater than 10, or even of greater than 12.

18. A method for starting up an aircraft turbine engine according to claim 1 , wherein said method comprises disengaging said disengageable connection device in order to move said movable member from the first position thereof into the second position thereof.

19. The aircraft turbine engine according to claim 1 , wherein said fan is decoupled from the low-pressure spool when said third element is separated from said stator casing.

20. The aircraft turbine engine according to claim 1 , wherein said device for removing power from said low-pressure shaft comprises a take-off shaft which extends through an arm of an intermediate casing.

21. The aircraft turbine engine according to claim 1 , wherein said device for removing power from said low-pressure shaft comprises an inner radial end having a first toothed wheel that meshes with a second toothed wheel mounted on the low pressure shaft.

22. The aircraft turbine engine according to claim 1 , wherein the stator casing is a stator casing of an inter-duct compartment which separates a primary duct from a secondary duct.

23. The aircraft turbine engine according to claim 1 , wherein the reduction gear is arranged between the fan and the low pressure spool.

Assignees

• Safran Aircraft Engines

Inventors

• Bordoni Nils
• Kubiak Guillaume Patrice
• Lemarchand Kevin Morgane

Classifications

• F02C7/277

  the starter being a {separate} turbine · CPC title

• F05D2260/85

  Starting · CPC title

• F05D2260/40311

  of the epicyclical, planetary or differential type · CPC title

• F05D2260/4023

  through a friction clutch · CPC title

• F02C7/36

  Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H) · CPC title