Aircraft comprising a turbine engine incorporated into the rear fuselage with variable supply

Patent metadata

Field	Value
Publication number	US-10082040-B2
Application number	US-201615745704-A
Country	US
Kind code	B2
Filing date	Jul 21, 2016
Priority date	Jul 22, 2015
Publication date	Sep 25, 2018
Grant date	Sep 25, 2018

Abstract

The invention concerns an aircraft propelled by a turbine engine having contrarotating fans ( 7, 8 ), the turbine engine being incorporated at the rear of a fuselage ( 1 ) of the aircraft, in the extension of same and comprising at least two gas generators ( 2 a, 2 b ) that supply, via a shared central stream ( 4 ), a power turbine ( 3 ), the turbine ( 3 ) comprising two contrarotating rotors ( 5, 6 ) for driving two fans ( 7,8 ) disposed downstream from the gas generators ( 2 a, 2 b ), said aircraft comprising means ( 15 ) arranged for separating the gas flow in the power turbine ( 3 ) into at least two concentric streams ( 16, 17 ) and a device comprising first means for distributing the gas flow ( 21 - 24 ) between said streams ( 16, 17 ) from the central stream ( 4 ), the first distribution means being configured to be able to open or close the supply of at least one so-called sealable stream ( 16 ) of the streams ( 16, 17 ) of the power turbine ( 3 ).

First claim

The invention claimed is:

1. Aircraft propelled by a turbine engine having contra-rotating fans, the turbine engine being integrated at the rear of a fuselage of the aircraft, in a continuation of the fuselage of the aircraft, and comprising at least two gas generators which feed a power turbine through a common central flow duct, said power turbine comprising two contra-rotating rotors for driving two fans arranged downstream of the at least two gas generators, wherein the aircraft comprises separating means arranged so as to separate a gas flow in the power turbine into at least two concentric flow ducts and a device that comprises a first distribution means for distributing the gas flow between the at least two concentric flow ducts from the common central flow duct, said first distribution means being designed to be able to open or close a gas feed of at least one flow duct, referred to as at least one concealable flow duct, from among the at least two concentric flow ducts of the power turbine.

2. Aircraft according to claim 1 , wherein each contra-rotating rotor of the power turbine comprises at least one blade ring, the separating means for separating the flow into concentric flow ducts comprise fins extending circumferentially between a plurality of blades of each blade ring, at an intermediate radius between radial ends of the plurality of blades.

3. Aircraft according to claim 1 , wherein a radially innermost concentric flow duct of the at least two concentric flow ducts in the power turbine is said at least one concealable flow duct.

4. Aircraft according to claim 1 , wherein said at least one concealable flow duct comprises an intake opening in the common central flow duct, defined between two edges that are substantially defined by the same curve around an axis of the turbine engine and are offset along said axis.

5. Aircraft according to claim 4 , wherein the first distribution means comprises a part that is movable in translation along the axis of the turbine engine, and is designed such that said part opens or closes the intake opening of said at least one concealable flow duct of the power turbine depending on its translational position.

6. Aircraft according to claim 1 , wherein said aircraft comprises means designed to feed a ventilation air flow to said at least one concealable flow duct of the power turbine when the gas feed of said at least one concealable flow duct through the central flow duct is closed.

7. Aircraft according to claim 6 , wherein said aircraft comprises closing means designed to close an outlet of said at least one concealable flow duct of the power turbine when the gas feed of said at least one concealable flow duct through the central flow duct is closed.

8. Aircraft according to claim 7 , wherein the closing means for closing the outlet of said at least one concealable flow duct are controlled according to the difference between a resilient restoring force and a gas pressure at the outlet of said at least one concealable flow duct.

9. Aircraft according to claim 8 , wherein a plurality of support arms for supporting a downstream housing, which rotates together with a rotor of the two contra-rotating rotors of the power turbine, are distributed in a ring at the outlet of said turbine, said plurality of support arms supporting said closing means.

10. Aircraft according to claim 9 , wherein said closure means comprise flexible strips, each being secured to one support arm of the plurality of support arms in the region of a leading edge and extending in a circumferential direction up to a neighbouring support arm of the plurality of support arms when no force is exerted thereon in a direction originating from said at least one concealable flow duct.

11. Aircraft according to claim 1 , wherein the device comprises a second distribution means comprising a ring of substantially radial, variable-pitch stator vanes.

12. Aircraft according to claim 11 , wherein said aircraft comprises means that are intended for actuating the second distribution means and designed to vary a passage cross section for a primary flow in the common central flow duct depending on intake conditions of the at least two gas generator(s).

13. Aircraft according to claim 12 , wherein the second distribution means are arranged so as to vary the passage cross section for the primary flow in the common central flow duct up to a minimum value that is equal to or less than a passage cross section in one of the at least two concentric flow ducts not closed by the first distribution means.

14. Aircraft according to claim 11 , wherein the second distribution means are located downstream of the first distribution means in the gas flow.

15. Aircraft according to claim 1 , wherein a first flow duct of said at least two concentric flow ducts is radially inner in relation to a second flow duct of said at least two concentric flow ducts.

16. Method for managing a breakdown of the turbine engine of the aircraft according to claim 1 , for changing over from operation using two gas generators to operation using just one, comprising:
a closing step a), for closing a second distribution means so as to adjust the gas flow in the power turbine to a single gas generator in operation, and
a step b) consisting of closing the first distribution means while re-opening the second distribution means, so as to operate the power turbine using a single concentric flow duct of the at least two concentric flow ducts while maintaining the gas flow adapted to operation of the single gas generator.

17. Method for controlling the turbine engine of the aircraft according to claim 1 , wherein a position of a second distribution means is controlled according to parameters influencing the operating ability of the low-pressure compressors of the at least two gas generators.

Assignees

• Safran Aircraft Engines

Inventors

• Gallet Francois
• Tantot Nicolas Jerome Jean

Classifications

• F05D2220/323

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

• F02K3/072

  with counter-rotating {, e.g. fan} rotors · CPC title

• B64C11/48

  Units of two or more coaxial propellers · CPC title

• F05D2260/606

  Bypassing the fluid · CPC title

• F01D17/145

  by means of valves, e.g. for steam turbines (valves in general F16K) · CPC title